スペースシャトル（）は、かつてアメリカ航空宇宙局 (NASA) が1981年から2011年にかけて135回打ち上げていた再使用をコンセプトに含んだ有人宇宙船である。

もともと「再使用」というコンセプトが強調されていた。しかし、結果として出来上がったシステムでは、オービタ部分は繰り返し使用されたものの、打ち上げられる各部分の全てが再利用できていたわけではなく、外部燃料タンクなどは基本的には使い捨てである。

 概要
初飛行は1981年、2回目の飛行は1982年で、2011年7月の135回目の飛行を最後に退役した。

スペースシャトルは宇宙輸送システム (Space Transportation System, STS) あるいはスペースシャトル計画の一環としてもちいられた。STSの開発とシャトルの飛行は、基本的にアメリカの資金によって行われた。主な使用目的は、NASAのおかれた様々な政治的状況や起こしてしまったシャトルの事故も影響して、およそ10年ごとに大きく変遷してきた地球ドラマチック「さようならスペースシャトル 〜栄光と挫折の30年〜 」前編および後編。イギリスの放送局の制作した番組。日本ではNHKによる放送、前編2012年1月21日および後編1月28日（再放送、1月30日および2月6日）。が、数々の人工衛星や宇宙探査機の打ち上げ、宇宙空間における科学実験、国際宇宙ステーション (International Space Station, ISS) の建設などである。なおシャトルはNASAによってだけでなく、米国国防総省、欧州宇宙機関、ドイツ等の軌道上実験にも使用された。

シャトルは再使用型宇宙往還機であり、軌道船 (Orbitor Vehicle, OV)、外部燃料タンク (External Tank, ET)、固体燃料補助ロケット (Solid Rocket Booster, SRB) の三つの部分によって構成されているShuttle Basics. NASA.。ETとSRBは上昇中に切り離され、軌道船 (OV) のみが地球周回軌道に到達する。発射時には機体は通常のロケットと同じように垂直に打ち上げられるが、軌道船は水平に滑空して帰還・着陸し、再使用のために整備された。SRBはパラシュートで海に降下し、回収船で回収されて整備した後、推進剤を再充填して再利用された。

 構造・飛行等の概略
thumb|250px|飛行の概略。打ち上げ、固体燃料補助ロケット (SRB) の切り離し、外部燃料タンク (ET) の切り離し、軌道上での作業、減速、大気圏再突入、着陸 の概念図
まずシャトルの構造および打ち上げ〜着陸の概略を説明する。

通常は5名から7名の飛行士が搭乗した。なお、最も初期の頃に行われた、STS-1からSTS-4の4回の試験飛行のように、機長と操縦士の2名だけでも飛行できた。

発射時のシャトルの構成は、おおまかに
 オレンジ色の外部燃料タンク (External Tank, ET)最初の打ち上げ（コロンビア号）の際は、外部燃料タンクも白色に塗装されていた。"NASA Takes Delivery of 100th Space Shuttle External Tank" . NASA, August 16, 1999. Quote: "…orange spray-on foam used to insulate…""Media Invited To See Shuttle External Fuel Tank Ship From Michoud". NASA, December 28, 2004. Quote: "The gigantic, rust-colored external tank…"
 2本の白色で細長い固体燃料補助ロケット (Solid Rocket Boosters, SRB)
 宇宙飛行士と貨物を搭載する軌道船 (Orbiter Vehicle，OV)
の三つの部分から構成されていた。なお、上記に加えて、STSのために開発された、PAMとIUSと呼ばれる人工衛星打上げ用の2種類の固体ロケットを用いれば、搭載物をさらに高い軌道に運ぶこともできた。なお、シャトルには全体でおよそ250万個もの部品が使われており、人間がこれまでに製造した中で最も複雑な機械であると言われている。
（→#構造・メカニズム・諸元）

シャトルは通常のロケットと同じように、発射台からは垂直に離陸する。その際の推力を生むのは2本のSRBおよび、（軌道船の後部に装着している）3基のメイン・エンジン (Space Shuttle Main Engine, SSME) であり、SSMEの推進剤（液体水素と液体酸素）は外部燃料タンクから供給される。上昇の手順はおおまかに、
 SRBも含めてすべてのロケットが噴射される第一段階
 SRBが役目を終えSSMEだけで推進する第二段階
のふたつに分かれていて、打上げからおよそ2分後に第二段階に移り、SRBは切り離され落下、パラシュートで海に着水し再使用のため船で回収される。機体（軌道船およびET）はその後も上昇を続け、軌道に到達するとSSMEが燃焼を停止し、ETも役目を終えて切り離される。切り離され自由落下を始めたET（巨大なオレンジ色のタンク）は通常は大気圏に再突入して空気抵抗と熱によって消滅する。ただし、様々な用途に使用することは、構想としてはあった。

軌道船はその後さらに軌道操縦システム (Orbital Maneuvering System, OMS) を噴射することでミッションの目標としている軌道へと向かう。軌道上での姿勢は、姿勢制御システム (Reaction Control System, RCS) を噴射することで制御する。

シャトルが従来の宇宙船とは際だって異なった特徴の一つに、軌道船の胴体部分のほとんどを占めるほどの大きさの貨物搭載室を備えていることと、そこに大きな観音開きのドアがついていることである。これによって、飛行士や宇宙ステーションの建設資材などを、地球周回低軌道や大気圏上層部、さらには熱圏などに運ぶことができた。例えば、ハッブル宇宙望遠鏡のような大きなものを搭載し軌道に投入することや故障した衛星などがあれば、その軌道へ向かい、貨物室に回収して地球に持ち帰ったりすることもできた。

任務が終了すると、軌道船はOMSを逆噴射して速度を落とし大気圏に再突入した。降下している間、シャトルは大気の様々な層を通過し、主に空気抵抗を用いて機体の速度を極超音速状態から減速させる。大気圏下層部に到達し着陸態勢に入るとグライダーのように滑空飛行し、フライ・バイ・ワイヤ方式の操縦系統で油圧によって動翼を制御した。着陸の際には、長い滑走路が必要とされた。シャトルの形態は、帰還時に極超音速飛行および旅客機のような低速飛行の双方をしなければならない、という二律背反する要求を満たすために作られた妥協の産物であり、その結果として軌道船は着陸寸前には、普通の航空機には見られないような急激な降下（高い降下率）を経験することになる。

（→#飛行手順の詳細）

 かさんだコストと危険性
当初は通常のロケットより一回あたりの飛行コストを安くできるという見込みでこの計画がスタートし製造されたが、実際の運用で発生した事故に対する安全対策により、当初の予想より保守費用が大きくなっていき、結果的に使い捨てロケットよりもコストが高くなった。（→#甘すぎた予測と膨らんだ費用と危険性）

 呼称の指す範囲
「スペースシャトル」という言葉は、一般には軌道船（オービタ）の単体を指していることもある。シャトル（往復を繰り返すもの）という表現に合致しているのは基本的にオービタ部分であるし、形状という点でも、「シャトル」という用語の源となっている織物のシャトルと形が類似し連想させるのはオービタ単体であるからである。ただし技術的な観点、つまり宇宙飛行システム、飛行に必要な技術的な要素、という意味では、軌道船以外にも外部燃料タンク・固体燃料補助ロケットが結合されて、はじめてシャトルは完成状態となり飛行可能となるので、NASAのエンジニアなどは三つが合体した状態を「スペースシャトル」と呼ぶ。そして、紛らわしさを避けるために「オービタ」「SRB」「ET」などの呼称を用いて呼び分けている。

完成状態にする作業はスペースシャトル組立棟で行われる。なお、この建物は元々はシャトルのものではなく、アポロ計画のサターン5型ロケットを組み立てるために作られたものである。

「スペースシャトル」という用語で、スペースシャトルをコアとした計画全体（スペースシャトル計画）を指して用いられていることもある。

 計画・設計・製造
 計画の初期段階
thumb|180px|スペースシャトルの当初のコンセプト図の一部
シャトルの設計と製造は1970年代初頭に始まったが、その概念はそれより20年も前、1960年代のアポロ計画よりも早い段階に存在していたPlease refer to Space Shuttle design process.Please refer also to North American X-15#Design_and_development.。宇宙から宇宙船を水平に着陸させるという構想は1954年に国立航空諮問委員会(NACA)が描いていたもので、それは後にX-15航空工学実験調査機として実現することになった。NACAに対してこの提案を行ったのは、ヴァルター・ドルンベルガーである。

1957年、X-15をさらに発展させたXシリーズ宇宙往還機計画が提案された。宇宙飛行士ニール・アームストロングはX-15とX-20両方のテスト・パイロットに選抜されたが、X-20は計画されただけで実機が飛行することはなかったPlease refer to X-20 Dyna-Soar.。

X-20は実現されなかったが、同様のコンセプトを持つHL-10実験機は数年後に開発され、1966年1月にNASAの元へと届けられた。HLとは、「Horizontal Landing（水平着陸）」の意味であるPlease refer to Northrop HL-10.。

1960年代半ば、空軍は次世代宇宙輸送システムに関する一連の極秘調査計画を行い、「一部再使用型の宇宙船こそが最も安上がりな方法だ」と判断した。彼らの提案では、使い捨て型の宇宙船とロケット（クラスI）の開発に直ちに取りかかり、それに続いて一部再使用型（クラスII）の開発を続け、最終的には完全再使用型（クラスIII）に達するべきである、とされた。1967年、NASA長官ジョージ・ミューラー (George Mueller) は幹部80人を集め、将来的な選択肢に関する1日間の討論会を開催した。会議では、初期の頃の空軍のX-20計画を含む様々な提案がなされた。

1968年、NASAは地球と宇宙を往復することを目的とした「統合往還機 (Integrated Launch and Re-entry Vehicle, ILRV)」の研究を開始し、同時に複数の企業に対してメイン・エンジン (SSME) の開発を競わせた。ヒューストンとハンツビルにあるNASAの事務局は共同で、宇宙に貨物を運ぶだけでなく大気圏を滑空して地球に帰還できるような宇宙船の設計を公募した。その中の一つに、巨大なロケットと小型の軌道船によって構成されたDC-3と呼ばれた案があったPlease refer to North_American_DC-3#History.。

1969年、ニクソン大統領はスペースシャトル計画を進行させることを正式に決定したPlease refer to Space_Shuttle_design_process#Decision-making_process.。

1973年8月、X-24Bが飛行に成功したことにより、大気圏に再突入した宇宙船が水平に着陸するのが可能であることが証明された。

 開発・設計
thumb|140px|1975年に行われた模型を使用しての風洞試験。再突入で機体が高温の空気に包まれる状態を、イオンガスを吹きつけることで再現した。
スペースシャトルは、再使用することを目的に設計された宇宙船としては初めてのものである。シャトルは様々な搭載物を低軌道に運び、ISS（国際宇宙ステーション）の人員を交代させることができ、軌道船は地球を周回する人工衛星その他の物体を回収し地上に持ち帰ることもできるように設計された。各軌道船は「100回の飛行もしくは10年間の使用に耐えられるように」との考えで設計されたが、後にその期間は延長された。STS（宇宙輸送システム）の設計責任者は、マーキュリー計画、ジェミニ計画、アポロ計画などでも宇宙船の設計を担当したである。軌道船の大きさや形状を決定する際の最も重要な要素となったのは、当時計画されていた商業衛星や秘密衛星の最大のサイズのものを搭載できるようにすることと、極軌道から一周回で離脱するという空軍の秘密計画に対応できるような飛行範囲を持っていることである。衛星を宇宙空間に配置するための高い搭載能力が欲しいという国防総省の要求、および再使用できる機器を持つ宇宙船を開発することによって宇宙開発予算を削減したいというニクソン政権の要求の双方に応えるため、固体燃料補助ロケットと使い捨て型の燃料タンクの併用という方式が選択された。

 耐熱タイル
シャトル開発でひとつの大きな壁になったのが、大気圏に再突入時の熱からオービタを守り、繰り返し使用可能な熱シールドの開発である。オービタは機体を軽量にするために、基本的に航空機と同様のアルミニウムで出来ているが、アルミニウムという素材はわずか200度程度の温度で柔らかくなってしまい、大気圏再突入時に発生する1600度以上の熱に耐える事は出来ない。そこで、断熱材として素材にシリカガラス繊維を用いた耐熱タイルが開発された。シリカは熱を伝える速度が非常に遅いので、それを用いた耐熱タイルを用いれば機体のアルミを護ることができる。だが、まだ問題があった。機体のアルミは熱で膨張するのに対し、耐熱タイルのほうはほとんど膨張しない為、そのまま接着しては温度上昇とともに耐熱タイルは剥がれて脱落してしまう。試行錯誤が繰り返された結果、機体と耐熱タイルの間にフェルトをはさむ事で機体とタイルの膨張率の違いを受け止める方法が浮上した。これは特殊なフェルトではなく、カウボーイハットなどに用いられるごく普通のフェルトである。機体とフェルトと耐熱タイルの接着についても、アメリカの家庭にありふれた浴槽の防水コーキング用のゴムが接着剤として用いられた。耐熱タイルは2万5千枚製造され、オービタの曲面を覆うため、部分ごとに形状の異なるものがジグソーパズルのように機体に貼り付けられた。

素材選択や接着方法の開発が難航した耐熱タイルは、やはりスペースシャトルの弱点のひとつとなり、繰り返される飛行で何度も脱落を起こし、大事故の原因にもなった。安全確保のため、帰還後の点検で毎回毎回タイルひとつひとつの状況や履歴を記録しつつ手作業で検査・修復しなければならず、シャトルの不安要因のひとつ、大きな重荷のひとつとしてつきまとうことになった。

 製造
飛行可能な機体は6機製造された。初号機エンタープライズは宇宙に行けるようには作られてはおらず、もっぱら滑空試験のためのみに使用された。実用化されたのは、コロンビア、チャレンジャー、ディスカバリー、アトランティス、エンデバーの5機である。当初はエンタープライズも進入着陸試験が終了した後に実用機として改造される予定だったが、構造試験のために製造されたSTA-099をチャレンジャー (OV-099) に改造したほうが安上がりだと判断された。チャレンジャーは1986年、発射から73秒後に爆発事故を起こして機体が失われたため、機体構造の予備品として残っていたものを集めて新たにエンデバーが製作された。コロンビアは2003年に空中分解事故を起こして消滅した。
thumb|center|400px|シャトル5機。このうちの2機が事故で失われることになった。

 甘すぎた予測と膨らんだ費用と危険性
スペースシャトル計画の始まりの段階で、NASAの関係者には「一回の飛行あたり1200万ドルほどのコストで飛ばすことができる」などと主張する者もいて、そうした甘い見込みのもとに計画は進んでしまった。

シャトルを繰り返し安全に飛ばすため、再使用する機体の部品は飛行のたびに徹底的な検査が行われたが、シャトルを構成する膨大な数の部品の検査にかかる費用は巨額のものとなった。

エンデバーの製作にかかった費用は約17-18億ドルで、シャトルの一回の飛行にかかる費用は2002年の時点では約4億5,000万ドルだった。だが、コロンビアの事故以降は安全対策のコストが上昇し、2007年には1回の飛行につき約10億ドルを要するようになったhttp://www.nasa.gov/centers/kennedy/about/information/shuttle_faq.html。

スペースシャトルには技術的な困難だけでなく、官僚主義に侵されたNASAという巨大組織の抱える問題も影響した。チャレンジャー号の事故は予測・回避できた可能性が高かったにもかかわらず、NASAの幹部は「事故は起きないだろう」と充分な対策を行わず、米国が行った宇宙飛行中の事故では初の死者を出している。コロンビア号の事故においても、発射時の映像を確認した職員によって上昇中に剥離した断熱材がオービタに衝突した可能性が指摘されたものの、NASA幹部は提供された情報を軽視したという経緯がある。

政治学者のロジャー・A・ピールケ・Jr. (Roger A. Pielke, Jr.) は、2008年度初頭までにシャトル計画にかかった費用は総額で1,700億ドル（2008年度換算）ほどと算定した。これによれば打ち上げ一回あたりのコストは15億ドルということになる The Rise and Fall of the Space Shuttle, Book Review: Final Countdown: NASA and the End of the Space Shuttle Program by Pat Duggins, American Scientist, 2008, Vol. 96, No. 5, p. 32.。

最終的には、スペースシャトルの計135回の打ち上げで2090億ドルもの費用がかかっていた。

 構造・メカニズム・諸元
 軌道船
thumb|left|300px|軌道船解剖図
thumb|left|300px|軌道船解剖図

軌道船は多くの航空機と似たような形状をしており、主翼は内側が81°、外側が45°の後退角を持った二重デルタ翼で、垂直尾翼の後退角は50°である。主翼の後端には4枚の動翼が取りつけられている。垂直尾翼後端には空力ブレーキも兼ねた方向舵が設置されていて、降下と着陸の際に高揚力装置（フラップ）とともに作動して機体を制御する。

胴体部分のほとんどは直径4.6m、長さ18mの貨物搭載室が占めていて、観音開きの保護ドアによって覆われている。搭載物は通常は機体が水平の状態にあるときに格納され、その後機体とともに発射台上に垂直に設置される。無重力の宇宙空間では、搭載物は飛行士が操縦するロボットアームや船外活動によって放出される。搭載物自体が持っているロケットによって、さらに高い軌道へと投入されることもある。

thumb|upright|中央に見えるのが軌道船の3基のメイン・エンジン。その両脇にある2基の小さいロケットは軌道操縦システム。さらにその上にあるのが垂直尾翼。
機体の後端には、メイン・エンジンが三角状に配置されている。エンジンのノズルは上下方向に10.5°、左右方向に8.5°傾けることが可能で、上昇中に推力の向きを変えて機体の進行方向を制御する。軌道船の機体構造は主にアルミニウム合金によって作られているが、エンジン部分の支持構造にはチタニウム合金が使用されている。

軌道船は飛行目的に応じて、軌道実験室（スペースラブ、スペースハブ）、搭載物をより高い軌道に投入するためのロケット（慣性上段ロケット (IUS)、ペイロード・アシスト・モジュール (PAM)）、軌道滞在期間延長機器（EDO (Extended Duration Orbiter) キット）、カナダ・アームなど様々な追加機器を搭載することができる。

製造された機体の中で実際に宇宙に行くことができたのは、OV-099チャレンジャー号、OV-102コロンビア号、OV-103ディスカバリー号、OV-104アトランティス号、OV-105エンデバー号の5機である。

オービタに加えられた主な機器の画像

 外部燃料タンク

外部燃料タンク (ET) の主な機能は、軌道船のメイン・エンジンに燃料の液体水素と酸化剤の液体酸素を供給すると同時に、2本のSRBと軌道船を接続し、全体を支える骨組みとなることである。ETはシャトルの中では唯一再使用されない部分で、飛行のたびに投棄されているが、軌道に投入して（宇宙ステーションに接続するなどして）利用することは構想としては検討されていたNASA-CR-195281, "Utilization of the external tanks of the space transportation system". NASA, August 23–27, 1982.STS External Tank Station . astronautix.com。

 固体燃料補助ロケット

固体燃料補助ロケット (SRB) は2基合計で発射時に必要とされる推力の83%、約1,250万ニュートン（1,276.8トン）を発生しColumbia Accident Investigation Board Report, Vol II, Appendix D.7 . NASA, October 2003.、打上げから2分後、高度約15万フィート(46km)に達したところで切り離され、パラシュートで海に着水して回収される。外殻は厚さ13mmの鋼鉄でできている。SRBは何度も再使用されるもので、一例を挙げれば2009年に試験発射されたアレスI-Xロケットは、過去48回のシャトルの飛行で使用されたSRBの部品を寄せ集めて作られたものであり、その中には1981年の初飛行 (STS-1) で使われたものも含まれていたNASA Ares I First Stage Motor to be Tested August 25. NASA, July 20, 2009.。

 飛行制御系統
シャトルはコンピュータ制御されたフライ・バイ・ワイヤ方式のデジタル飛行制御システムを採用した、初期のころの機種の一つである。これは飛行士が操作する操縦桿やペダルと、機体の操縦翼面や姿勢制御システムの間に機械的なリンクや油圧系統などが一切存在しないということを意味する。飛行士が入力した操作は電気信号に変換され、電線（ワイヤ）を介して操縦装置に伝えられる。

フライ・バイ・ワイヤ方式の最大の懸念は信頼性の問題であり、シャトルのコンピューターシステムについては多くの研究開発が行われた。シャトルは IBM製の5台のAP-101と呼ばれる、それぞれ独立して冗長性を持ち、組み込みシステムを構成する32ビット汎用コンピューターを使用している。このうち4台は主飛行電子ソフトウェアシステム (Primary Avionics Software System, PASS) という特製のソフトウェアで稼働し、残りの1台はこれとは別の、バックアップ飛行システム (Backup Flight System, BFS) というソフトを使用している。これらを総称して「データ処理システム (Data Processing System, DPS)」と呼ぶ。

シャトル用DPS設計の到達目標は、フェイルセーフを達成して信頼性を向上させることだった。DPSは、もし5台のコンピューターのうち1台が故障してもミッションを継続することができ、2台が故障しても安全に着陸できるように設計されている。

4台の汎用コンピューターは、相互に監視し合いながら稼働している。もし1台が他と違う指令を出した場合は、3台が「投票」を行い、違う指令を出している1台を機体の制御から除外する。残りの3台のうち1台がまたもや違う指令を出した場合は、残った2台が投票をしてその1台を除外する。極めて稀な場合だが、もし4台の「主張」が2対2に別れた場合は、どちらか一方のグループが無作為に選ばれる。

BFS（バックアップ飛行システム）は5台のコンピューターの中で独立して開発されたソフトで、4台のメインシステムが故障した時にのみ稼働する。BFSが開発されたのは、メインシステムはハードウェア的には冗長性を持たせているものの全く同じソフトで稼働しているため、もし何らかのエラーが発生した時には4台すべてが故障してしまう可能性があるからである。埋め込み式アビオニクスソフトは、一般の商用ソフトとは全く違う環境のもとで開発されている。コードラインの数は商用ソフトに比べればごく限られたもので、変更がなされることは滅多になく、広範な試験が行われ、ほんのわずかなコンピューターコードのために開発要員や試験要員も含めて多くの人員が関わっている。しかし、どんなに万全を尽くしても故障というのは常に起こりうるものであり、そのような不測の事態に備えてBFSは用意された。シャトルが退役するまでの間、実際にBFSが操縦を引き継ぐような事態が発生することは一度もなかった。

シャトルのコンピューターのソフトウェアは、PL/Iに似たHAL/Sと呼ばれる高級プログラミング言語で書かれている。これはリアルタイム組み込みシステム環境のために、特別に設計されたものである。

IBM製AP-101コンピューターは、もともと1台あたり約424KBの磁気コアメモリを持ち、CPUは毎秒40万回の計算を行うことができた。ハードディスクはなく、ソフトは磁気テープカートリッジからロードした。

1990年、AP-101はAP-101Sという上位機種に置きかえられた。記憶容量はこれまでの2.5倍の約1MBに、演算速度は3倍の毎秒120万回に向上し、さらに記憶装置は磁気コアメモリからバックアップ電池つきの半導体メモリに改良された。

1983年11月から、シャトルにはグリッド・コンパス (GRiD Compass) と呼ばれる、世界で最も初期の頃に作られたラップトップ型コンピュータを使用していた。コンパスはせいぜい8,000ドル程度で売られていた安価なものだが、その重量や大きさに比して不釣り合いなほどの性能を当時発揮し、NASAはその重要な顧客の一つだった。なお、グリッド・コンパスは飛行制御系統には関係せず、シャトルの飛行軌跡を2周回分表示させるのに使用された。

thumb|100px|シャトル計画の標章

 機体の塗装と標章
操縦室の窓と貨物搭載室ドアの間の機体側面には、軌道船の名称が書かれている。搭載室ドア後部の下側には、NASAの標章と「United States」の文字および星条旗が描かれている。国旗は右側主翼にももう一つある。文字に使用されている書体はHelveticaである。

 改良
シャトルは1970年代に開発された宇宙船であるため、その当時から安全面における性能や信頼性を向上させるべく多くの改良や改造が施されてきた。
thumb|STS-101の操縦席。アトランティス号でグラスコックピットが初めて採用された。
内部構造のほとんどは初期に設計されたものとそれほど変わってはいないが、アビオニクス（飛行用電子機器）は大きく変貌した。たとえばコンピューターのアップグレード（性能向上）に関して言えば、初期の頃のアナログ式のメーター類は廃止され、最新型のエアバスA380やボーイング777に使われているような、グラスコックピットと呼ばれるフルカラーの液晶表示板に改められた。HP-41Cのようなプログラム入力可能な電卓も、依然として使われている。ISS（国際宇宙ステーション）の登場により、ISSに補給物資を届ける飛行でより多くの貨物をミッドデッキに搭載できるよう、内部エアロックは外部エアロックに置き換えられた。外部エアロックの上部には、ISSとのドッキングに使うロシアのアンドロジナスドッキング機構が使われた。

SSME（メイン・エンジン）もまた、信頼性と出力を向上させるべく何度も改良を施されてきた。発射時に「エンジンの出力を104%に上げる」という言い回しが存在することはその名残である。これは安全上の限界を超えてエンジンを噴射するという意味ではなく、初期のエンジン出力と比較しての値を指す。長い開発期間のうちに製造元のロケットダイン社は、安全出力を当初の設計値の104%にまで向上させることができたのだが、これまでに作成した膨大な量の文書やソフトを書き直す必要を避けるため104%という言い回しが残ることとなった。SSMEの進歩の歴史は、フェーズII、ブロックI、ブロックIA、ブロックIIA、ブロックII のような「ブロック番号」となって残されている。これらの改良によってエンジンの信頼性・メンテナンス性・性能は大きく向上し、2001年にはブロックIIエンジンを109%の推力にまで到達させることができた。ただし通常使用される最大推力は104%までで、106%または109%が実現されるのは緊急事態が発生して飛行が中止される時だけである。

最初の二回の飛行STS-1とSTS-2では、外部燃料タンクが太陽光を吸収して内部の温度が上昇するのを防ぐため全体が白色に塗られた。しかし地上での試験で必要ないことが分かったので次回からは廃止され、その塗料の分だけ軌道に投入できる搭載量が増えることとなった。他のところでは、液体水素タンク内部の桁のいくつかも不要なことが判明したため、軽量化のために取り除かれた。改良を施された軽量タンクはほとんどの飛行で使用されてきたが、STS-91からは超軽量タンク(SLWT)に置きかえられた。改良型の超軽量タンクにはアルミニウム/リチウム合金2195が使用されていて、最終型の軽量タンクに比べ3.4トンの減量に成功した。シャトルは無人では飛行できない設計になっているため、これらは実際の飛行で試してみる以外に手段がなかった。

SRB（固体燃料補助ロケット）もまた、何度も改良されてきた。代表的なところではチャレンジャー号爆発事故の後、本体接合部分の密閉性を確保するOリングが三重に強化された。

SRBには他にも性能や安全性を高めるためのいくつかの改良が試みられたが、実現されることはなかった。その中の一つに、より簡略かつ低コストで、安全面や性能にも格段の向上を果たしたと考えられる発展型SRB（Advanced Solid Rocket Booster, ASRB）があった。ASRBは1990年代半ばに宇宙ステーション計画支援のため製造が開始されたが、開発費が22億ドルにまではね上がったため中止が決定された。この代替案として、搭載能力を向上させるために超軽量タンクが開発されたが、安全性は向上しなかった。空軍は独自に、分割式ではない一体成形型の軽量SRBを開発していたが、こちらもまたキャンセルされた。

1995年、発射台上で準備作業をしていたディスカバリー号のETの発泡断熱材にキツツキが穴を空けたため、発射が遅れるという事態が発生した。この時以来、NASAは発射台周辺に市販の鳥よけのためのフクロウの模型や風船を配置するようになった。これらは打ち上げの直前にすべて取り除かれる。ET断熱材は発泡スチロールのようにもろい物質であるため、発射の際の衝撃や空気抵抗ではがれ落ち、軌道船を大気圏再突入の熱から保護する耐熱タイルを傷つける事故がこれまでにもたびたび発生してきた。断熱材の剥落は2003年2月1日に発生したコロンビア号空中分解事故の原因になり、その後も何度も打上げスケジュールの延期の原因になった。

人間が搭乗せず、搭載物だけを宇宙に送る無人の発射計画も1980年代以来何度も提案されてきたが、そのたびに却下された。「」と呼ばれるこれらの計画は、シャトルで蓄積されてきた技術を応用し、再使用という特性を放棄することとひきかえに、大幅なコストの削減が期待できるはずだった。

最初の4回の飛行では、飛行士は離陸時と帰還時には完全密閉型のヘルメットを着用し、空軍の高々度用与圧服を改良した宇宙服を着た。5回目の飛行からはこの与圧服は廃止され、青いワンピースのフライトスーツと部分与圧ヘルメットを着用するようになったが、チャレンジャー号事故による2年間の中断の後に再開された1988年の飛行からは、打上げ/帰還時にはあまりかさばらないように改良されたヘルメットつきのオレンジ色の部分与圧服(Launch-Entry Suit: LES)を着用するようになった。1995年からは、完全与圧式の改良型与圧服 (Advanced Crew Escape Suit: ACES) に置き換えられた。

軌道船がISSとドッキングして宇宙に滞在できる期間を延長するために、ステーション・シャトル電力供給システム (Station-to-Shuttle Power Transfer System, SSPTS) が導入された。SSPTSはISSが発生した電力を使用して軌道船の消耗品の消費を抑えるもので、STS-118から実用化された。

 技術的詳細
thumb|軌道船とソユーズ宇宙船比較図（同縮尺）
thumb|主翼解剖図

軌道船諸元（OV-105エンデバー号）
 全長：37.237m
 全幅：23.79m
 全高：17.86m
 空虚重量：78,000kg
 離陸時総重量：111,000kg
 最大着陸重量：100,000kg
 主エンジン：ロケットダイン社製ブロックII-SSME3基。1基あたり海面推力1.752MN（178トン、104%推力発生時）
 最大搭載量：25,060kg
 貨物室寸法：4.6m×18.0m
 運用高度：190〜960km（100〜520海里）
 最大速度：秒速7.743km(時速27,870km マッハ22.57相当)
 軌道範囲：2,009km（1,085海里）
 定員：飛行によって異なる。初期の頃は最小人員の2名で飛行したが、後の多くの飛行では5名になり、その後7名（船長、パイロット、数人の搭乗運用技術者、まれに航空機関士(フライトエンジニア)）で構成するのが一般的になった。STS-61-AとSTS-71の2回の飛行では8名が搭乗した。STS-3xxと呼ばれる緊急救助飛行では、11名(4人乗りで打ち上げて、7人を移乗)を搭乗できるよう検討されていた。

外部燃料タンク諸元（超軽量タンク）
 全長：46.9m
 直径：8.4m
 燃料容量：2,025m
 空虚重量：26,535kg
 発射時重量：756,000kg

固体燃料補助ロケット諸元
 全長：45.46m
 直径：3.71m
 空虚重量（1機あたり）：68,000kg
 発射時総重量（1機あたり）：571,000kgSpace Shuttle Propulsion Systems, p. 153. NASA, June 26, 1990.
 推力（発射時、海面推力）：12.5MN（1,281,360kg）

完成型詳細
 全長：56m
 発射時総重量：2,000,000kg
 発射時総推力：30.16MN(3,091,680kg)

 飛行手順の詳細
 発射
200px|thumb|STS-79で、移動式発射台上に設置されるアトランティス号。両主翼上部に見える灰色の箱状のものは、シャトルに推進薬を充填したり電力供給などを行うための地上設備。
thumb|200px|2004年、ケネディ宇宙センター39番発射台で行われた騒音抑制装置の試験の様子。発射時には、爆音で機体が損傷することを防ぐために41秒間で1,100m³の水が放出される。
thumb|200px|発射台から上昇してゆくところ。（アトランティス号、STS-115）
thumb|160px|打ち上げられた後に地上から見える、空中に残された白い筋（コロンビア号、STS-107）
thumb|200px|高高度を上昇中のスペースシャトル。この写真は望遠撮影（アトランティス号、STS-27）
シャトルの発射は、すべてケネディ宇宙センターで行われる。発射時に適用される天候基準は以下のとおりである。ただし、これだけに限定されるものではない。
 発射台周辺や飛行経路に一切の降雨があってはならない。
 気温は摂氏2℃以上37℃以下でなければならない。
 高度2,400mに上昇するまでの間に機体の姿を隠してしまうような雲が存在してはならない。
 高度9,000mに到達するまでの間、雷が発生する確率が20%を超えてはならない。
特に落雷が起きる可能性がある場合には、シャトルは絶対に発射されない。航空機はしばしば雷の直撃を受けることがあるが、構造が伝導体であることや、電気的に接地されていないために電流が空気中に放電されることなどにより、機体が悪影響を受けることはない。これに対してシャトルは、機体構造は通常のジェット旅客機と同じように伝導性のアルミニウムで作られているので内部機器が電流の影響を受けることはないが、発射時に噴射される噴煙が機体と地面をつなぐ電線の役目を果たしてしまう。このためNASAの基準では、周辺10海里以内に積乱雲が発生している場合には発射を行ってはならないことになっているWeather at About.com. What is the Anvil Rule for Thunderstorms?. Retrieved 2008-06-10.。当日は気象担当官が発射台周辺のみならず、大西洋を越えた緊急着陸地点やSRB（固体燃料補助ロケット）の回収点の天候なども監視し、最終的に発射を行うかどうかを判断するNASA Launch Blog. . Retrieved 2008-06-10.。シャトルは雷に対してはまず安全だとは思われるが、アポロ12号が発射された時には実際に落雷で船内が一時停電する事故が発生したため、NASAはこの件については特に慎重になっている。

長い間、シャトルは12月31日と1月1日をまたがっては飛行できなかった。1970年代に開発されたシャトル用のソフトウェアは年越しができるようには設計されておらず、もし飛行中にそれを強行するとコンピューターをリセットしなければならなくなり、予測できないようなエラーが発生する可能性が生じるからである。NASAの技術者がこの問題を解決したのは2007年のことで、これによってようやくシャトルは年を越えて飛行できるようになった。

発射当日はTマイナス9分前の最後のホールド（待機）が解除された後、いよいよ最終的な準備段階に入り、管制センターに設置された地上の打上げ管制装置 (Ground Launch Sequencer, GLS) が秒読み作業を引き継ぐが、もしシャトルに搭載された機器に重大な問題が発生した場合には秒読みは自動的に停止される。発射31秒前には、「オート・シークエンス・スタート (Auto Sequence Start)」と呼ばれる作業工程によって秒読み作業がGLSからシャトルのメイン・コンピューターに引き継がれる。

発射16秒前（Tマイナス16）、騒音抑制装置が作動し、猛烈な音響で機体が損傷を負わないようにするために移動式発射台やSRBの火炎偏向板（フレームトレンチ）に1,100m³の水が放出されはじめるNational Aeronautics and Space Administration. "Sound Suppression Water System"  Revised 2000-08-28. Retrieved 2006-07-09.。

発射10秒前（Tマイナス10）、SSME（メイン・エンジン）のノズル内に停滞している水素ガスを燃焼させて除去するために、ノズルの下で電気火花が飛ばされはじめる。エンジン周辺にこれらのガスが残っていると、点火する過程で搭載した検知機が異常を感知して、異常な加圧を招いたり爆発したりする可能性がある。この時、SSMEのターボ・ポンプが作動して燃焼室内に液体酸素や液体水素を供給しはじめる。この間、軌道船の4台のコンピューターは相互に指令を交わし、点火に必要なすべての動作を制御する。

発射6.6秒前（Tマイナス6.6）、SSMEの点火が始まる。点火指令は軌道船のGPC（汎用コンピューター）を経由して、3番エンジン（右側）、2番エンジン（左側）、1番エンジン（中央）の順に120ミリ秒の間隔を置いて送られる。GPCはSSMEの推力を90%にまで到達させると同時に、ノズルの向きを所定の位置に固定するNational Aeronautics and Space Administration. "NASA - Countdown 101". Retrieved 2006-07-10.。エンジンに点火されると、騒音抑制装置の水が蒸発して大量の水蒸気となり、南側に向かって噴出される。3基のSSMEの推力はそれから3秒以内に100%に達しなければならず、もしそれが実現しなかった場合はGPCがエンジンを緊急停止させる。逆に正常に推力が発生されていることが確認されれば、SRBを発射台に固定している8本の爆発ボルトが吹き飛ばされ、SRBに点火される。この時間こそが「Tマイナス0」と規定されている発射の瞬間であり、この直後に機体は上昇を開始する。そしてSRBは、いったん点火されたら燃料をすべて消費するまで燃焼を停止することはできない。SRBの排気ガスは北側に向かって掘られた火炎坑に沿って音速に近い速度で噴出され、しばしば衝撃波を発生させる原因となる。GPCは、4台の汎用コンピューターに設定された「発射手順制御装置（Master Events Controller）」と呼ばれるプログラムを介して点火の手順を実行する。上昇中に様々な異常事態が発生したときの緊急対応手順（中止方法）は、広範囲なものが用意されている。その大部分を占めるのは最も複雑で大きな負荷がかかるSSMEに関するもので、SRBが原因でチャレンジャー号爆発事故が発生した後には、緊急対応手順はより拡充されたものになった。

SSMEに点火されSRBが発射台から解放されるまでの間、機体はエンジンの推力によって機首下げの方向にわずかに（操縦席付近で約2m）傾く。この運動は、NASAの隠語で「うなずき (nod)」あるいは「はじき (twang)」などと呼ばれている。その後機体は約6秒かけてまた元の位置に揺れ戻ってきて、完全に垂直になった瞬間にSRBに点火されて上昇を開始する。

発射整備塔を離れた直後、シャトルは予定軌道に対応するためロール運動とピッチ運動を開始し、ETとSRBが上になった裏返しの姿勢になる。機体はゆるやかな弧を描きながら上昇し、燃料はどんどん消費されて重量が軽くなっていくため、加速度は徐々に増加していく。発射直後の加速度は1.2Gで、SRBが切り離される直前は2.5Gに増大し、SRB切り離し直後はいったん0.9Gに落ち、その後SSMEが燃焼を停止する直前には3Gにまで達する。地球周回軌道に乗るためには垂直方向よりもむしろ水平方向への加速がより多く必要とされるが、機体が視界から消える前はほぼ垂直に上昇していくため、水平方向への運動はほとんど確認することはできない。ISSが周回している高度380km付近での周回速度は秒速7.68km、時速27,650kmで、地表付近ではマッハ23に相当する。ISSは赤道に対して51.6°の傾斜角をもって地球を周回しているので、シャトルがランデブーをするためにはその角度に合わせる必要がある。
thumb|upright|速度マッハ2.46、高度約20,000mに達した頃の機体表面の圧力図。気圧が低い部分から高くなるに従って青から赤へと色分けされている。灰色は機体をとりまく空気の密度を表している。「オーバーフロー (Overflow)」というソフトを使用して作成。
マックスQ付近では、機体の、特に主翼などの弱い部分にかかる空気力学的圧力を抑えるため一時的にSSMEの推力が65%にまで絞られる。その前後では、空気の急激な圧縮と断熱膨張によりベイパーコーン (vapor cone) やプラントル・グロワートの特異点が起こる。

発射126秒後、SRBをETにつなぎとめていたボルトが爆薬で切断される。SRBはブースター分離モーターを噴射して機体の後方へと押しのけられ、残った推力を偏向し180度のターンを行い燃焼を完全に終了し、真下を向いて落下する。SRBはパラシュートで海に着水して再使用のため回収されるが、シャトルはSSMEの推力でなおも上昇を続ける。この時点では、機体はSRBがなくなったことで推力と重量の比は1を下回っているため、SSMEの力だけでは地球の重力を振り切ることはできなくなる。しかし燃焼を続けるうちに燃料が消費されて徐々に機体が軽くなり、やがて推力：重量比は再び1を超え、最終的に軌道に到達するまで二度と1を下回ることなく加速を続ける。

機体はその後も機首をやや上に向けた姿勢で徐々に軌道を水平に近づけ、SSMEの力で加速する。発射から約5分45秒後、地上との直接通信が終了し、背面が宇宙空間に向いた姿勢になるよう機体を反転させる。地上との交信は、その後は追跡およびデータ中継衛星 (Tracking and Data Relay Satellite, TDRS) を介して行われる。

最後の10秒間には機体は相当に軽くなっているため、飛行士に負担をかけないよう加速度が3G以下になるように推力が絞られる。

メイン・エンジンは空転すると機器を傷める可能性があるので、燃料が完全に空になる前に停止される。液体酸素は液体水素よりも前に供給が停止される。液体酸素はより過激に反応する傾向があり、停止直後の加熱した金属部分に触れると爆発するかもしれないからである。ETはエンジン停止後に爆発ボルトで切り離され、大部分は大気圏内で消滅してわずかな部品がインド洋または太平洋に落下するが、どこに落ちるかは打上げプロファイルによって変わる。タンク内の配管はすべて密閉されており、圧力を解放するような装置は設けられていないため、ETは大気圏下層部で内圧によって破裂する。大気圏再突入時に表面の断熱材が焼失すると、内部に残っていた液体酸素や液体水素を熱から保護する手段がなくなるため、急膨張して爆発の大きな要因になる。このような手段によって、地上に大きな破片が落下するのを防いでいる。

ET分離直後は、軌道の近地点はまだ大気圏を離れてはいないので、そのままでは大気圏に再突入することになる。そのため軌道船は軌道操縦システム (Orbital Maneuvering System, OMS) を噴射し、近地点をより高い高度に設定してETと衝突するのを防止する。一部の飛行（すなわちISSミッションなど）では、打上げ能力を確保するためにOMSが、メイン・エンジンの燃焼後期に並行して使用された。投入時の軌道をこのように設定しているのは、ETを宇宙空間に放出せず大気圏内で廃棄するためと、もしOMSが点火しなかったり、何らかの理由で搭載室のドアが開かなくなるような事態が発生しても、このような軌道にしておけば自動的に地球に帰還できるから、という安全上の理由もある。

 軌道上
thumb|200px|軌道上のチャレンジャー号（1983年）
軌道に乗ると、シャトルは様々な、しばしば相互に関連した任務をこなす。1980年代から90年代にかけては、NASAとヨーロッパ宇宙機関が共同開発した宇宙実験室 (Spacelab) などを含む宇宙科学計画や多種多様な衛星や科学探査機の軌道投入に使用されてきた。90年代から2000年代にかけては衛星打上げの任務は減少し、計画の焦点はもっぱら宇宙ステーションの建設に移った。ほとんどの飛行は数日から2週間程度で終了するが、軌道滞在期間延長機器 (Extended Duration Orbiter) を搭載したり国際宇宙ステーションにドッキングすれば、滞在期間をさらに延長することもできる。

 大気圏再突入および着陸

シャトルの大気圏再突入の過程では、降着装置をおろすのと、対気速度計に使うピトー管 (air data probe) を展開する作業以外はすべてコンピューターが自動で行うが、もし何か緊急事態が発生した場合は手動で再突入することも可能である。滑走路への進入と着陸も自動操縦装置に任せることはできるが、通常は手動で行われる。

再突入の作業は、まず軌道船の飛行方向を反転させ、機体後部を進行方向に向けることから始まる。その姿勢でOMSロケットを進行方向に約3分間噴射し（逆噴射）、軌道周回速度を322km/hほど減速する。これにより、軌道の近地点を下げて大気圏上層部に入るようにする。逆噴射の間にかかる加速度は約0.1Gである。その後軌道船は反転して機首を下げ（地球から見ればひっくり返した姿勢になっていたので機首を上げる方向(ピッチ軸を時計回り)に180度回転）、機首を進行方向に向ける。逆噴射は、着陸地点のケネディ宇宙センターから見てほぼ地球の裏側の、インド洋上空の赤道付近で行われる。

thumb|140px|再突入時、機体表面の温度が1,500℃以上に上昇する状態を再現したシミュレーション画像
高度約120kmの熱圏下層部にさしかかる頃、機体にかかる空気抵抗が顕著になりはじめる。この時の速度はマッハ25（時速30,000km、秒速8.3km）ほどである。シャトルは40°ほどの迎角をとりつつ姿勢制御システムと動翼を併用して機体を制御し、長い航跡を引いて速度だけでなく熱も減少させながら次第に降下していく。空気抵抗が増加するにつれ、シャトルは宇宙船から次第に航空機としての性格を現すようになる。直進している間は、機体には機首を下げるかもしくは40°よりも高い迎角をとらせようとする力が働く。軌道船は途中で4回、70°以上の深いバンク角をとったS字飛行をする。この間迎角は40°を保ったままで、各Sターンは数分間行われる。この操作を行うことで、機体の運動エネルギーを上下方向ではなく左右方向に分散して減速する。このS字飛行が始まるのは熱負荷が最も強烈になる時間帯で、この間熱保護シールドは灼熱化し、加速度は最大となる。最後のターンが終わる頃には軌道船は完全に航空機(グライダー）となっており、機首を下げて機体を水平にし、着陸施設への進入作業が開始される。

軌道船の最大滑空比/揚抗比は速度によって相当に変化し、極超音速域では1:1、超音速域では2:1で、滑走路への進入と着陸を行う亜音速域では4.5:1にまで低下するhttp://klabs.org/DEI/Processor/shuttle/shuttle_tech_conf/1985008580.pdf。
大気圏下層部では、軌道船は毎秒50m（時速180km）という高い降下率を除けば通常のグライダーのように飛行する。この高い高い降下率は、しばしば「空飛ぶレンガ」「翼の生えたレンガ」と揶揄される。速度がマッハ3程度にまで低下したところで、機体の対気速度を検出するため、胴体前方下部の左右両側に設置された対気速度測定用のピトー管が展開される。

thumb|160px|着陸直前、一般の航空機と同じように降着装置をおろすアトランティス号
高度3,000m、滑走路端まで12kmに達したところで、進入および着陸操作が開始される。飛行士は空力ブレーキを作動させ、機体の速度を682km/hから着陸速度の346km/hにまで減速させる（一般的なジェット旅客機の着陸速度は260km/h程度である）。機体のノーズは滑走路手前ギリギリまでノーズダウンの状態であるが、着陸寸前でノーズアップの状態とされ機体下面の空気抵抗を利用してさらに減速が行われる。430km/hで降着装置がおろされ、343km/hあたりでタイヤが接地し着陸する。シャトルは通常航空機に対して重く、着陸時のタイヤへの加重は過酷で、ボーイング747の2-3倍の63.6トンにもなる。タイヤは16層構造で1本4000ドルで6回まで再使用できるが、実際は使い捨てである。空力ブレーキの作動を補助するために、後輪または前輪が接地したところで直径12mのドラグシュートが展開されるが、どちらの段階で開くかはシュートの展開モードの選定によって変わる。ドラグシュートは機体が110km/h以下になった段階で投棄される。

着陸後は、機体の表面温度が下がるまで数分間待ち、有毒な水素やヒドラジン、四酸化二窒素（姿勢制御システムや3台ある補助動力装置の燃料として使用される）、アンモニアが機体周囲から検出されないかを確認し終えるまで、軌道船は滑走路上で停止したままにされる。支援車両によってパージとベント用の配管が軌道船の燃料配管と貨物室への配管に取り付けられ、着陸後約45-60分かけて有害なガスが除去される。

以上の着陸行程は、基本的にグライダーとして動作するために、やり直しが行えない。そのため着陸地点の天候は厳重にチェックされ、気象予報によっては他の着陸ポイントに変更される。機体は航空機としては非常に重量があるため、通常の飛行場の滑走路では耐えられず、特別に強化された路面をもつ飛行場が選ばれた。通常はNASAシャトル着陸施設の長さ5.2kmの滑走路が主に使用されるが、カリフォルニア州のエドワーズ空軍基地も使用された。この他世界各地に予備の着陸地点が指定され、日本では嘉手納飛行場がその一つであった。

上記のような操縦特性とアプローチに習熟するため、NASAではガルフストリーム IIを改造したシャトル訓練機での訓練を行っていた。

 着陸施設
シャトルの着陸は、初期はカリフォルニア州エドワーズ空軍基地に、ケネディ宇宙センターの滑走路が整備された後は基本的にはケネディ宇宙センターで行われることが多かった。ケネディ宇宙センターの天候が不順な場合は回復するまで宇宙で待機したり、あるいはエドワーズ空軍基地やその他世界中に配置された代替基地に着陸することもできた。ただしケネディ以外の施設に着陸するということは、その後にシャトル輸送機でケープ・カナベラルまで運ばれて来なければならないことを意味した。代替着陸基地の一つのホワイト・サンズ空軍基地の滑走路はシャトルが着陸するには整備が不十分という問題もあった。STS-3ではコロンビア号がニューメキシコ州のホワイト・サンズ空軍基地に着陸したが、この滑走路は当時はまだ整備が行き届いておらず、細かい砂が機体に入り込んでその後の整備が大変だった。シャトルを空輸するためのクレーン設備も準備する必要があるなど問題があった。結局、同基地に着陸したのはこの1回だけである。

代替着陸施設は多数あるが、エドワーズ空軍基地とホワイト・サンズ空軍基地以外は使用されることはなかった。エドワーズ空軍基地についても、シャトルの大陸横断に掛かる多額のコストのため近年はできるだけ利用しない方針が採られており、日本人最後の乗務となったSTS-131の着陸時にも一時は使用が決定していたがhttp://www.cnn.co.jp/science/AIC201004200009.html、最終的にはケネディ宇宙センターへの着陸となった。

 飛行記録

主なシャトルの飛行記録は以下のとおりである。
thumb|1977年、シャトル計画の一環である進入着陸試験で、カリフォルニア州エドワーズ空軍基地のドライデン飛行研究センターに向けて初飛行するエンタープライズ号 (OV-101)
+ スペースシャトル主飛行記録 日時 軌道船 主なできごと/注記 1977年2月18日 エンタープライズ  シャトル輸送機に搭載されての初飛行 1977年8月12日 エンタープライズ 初の単独滑空飛行。ロジャース乾湖に着陸。 1977年10月12日 エンタープライズ 三度目の飛行。尾部保護カバーを取り除いての初飛行。ロジャース乾湖に着陸。 1977年10月26日 エンタープライズ エンタープライズ最後の滑空試験。エドワーズ空軍基地のコンクリート滑走路への初着陸。 1981年4月12日  コロンビア 宇宙空間への初飛行 (STS-1) 1982年11月11日 コロンビア 4名の飛行士を搭乗させての初の実用飛行 (STS-5) 1983年4月4日 チャレンジャー チャレンジャー初飛行 (STS-6) 1984年8月30日 ディスカバリー ディスカバリー初飛行 (STS-41-D) 1985年10月3日 アトランティス アトランティス初飛行 (STS-51-J) 1986年1月28日 チャレンジャー 発射73秒後に機体が爆発 (チャレンジャー号爆発事故) (STS-51-L)。7名の飛行士全員が死亡。 1988年9月29日 ディスカバリー チャレンジャー号事故後の初の再開飛行 (STS-26) 1989年5月4日 アトランティス シャトルを使用しての初の探査機発射（マゼラン、STS-30） 1990年4月24日 ディスカバリー ハッブル宇宙望遠鏡発射 (STS-31) 1992年5月7日 エンデバー エンデバー初飛行 (STS-49) 1996年11月19日 コロンビア 17日間と15時間にわたるシャトルの最長宇宙滞在記録 (STS-80) 2000年10月11日 ディスカバリー シャトル100回目の飛行 (STS-92) 2003年2月1日 コロンビア 大気圏再突入時に空中分解 (STS-107)。7名の飛行士全員が死亡。 2005年7月25日 ディスカバリー コロンビア号事故後の初の再開飛行 (STS-114) 2010年2月8日 エンデバー 最後の夜間発射 (STS-130) 2010年5月14日 アトランティス アトランティス 号の計画上での最後の飛行 (STS-132)（後にSTS-135が追加され、それが最後の飛行になった） 2011年2月24日 ディスカバリー ディスカバリー最後の飛行 (STS-133) 2011年4月29日 エンデバー エンデバー最後の飛行 (STS-134) 2011年7月8日 アトランティス アトランティスおよびスペースシャトル計画最後の飛行 (STS-135)
出典：NASA打上げマニフェスト、NASAスペースシャトル公文書記録

 事故

1986年1月28日、スペースシャトルチャレンジャー号が発射から73秒後に右側のSRBのOリングの故障が原因で空中分解し、搭乗していた7名の飛行士全員が犠牲になった。機体の最重要機器の一つであるOリングが、異常寒波が原因の低温により損傷した。現場の技術者は再三にわたり12℃以下の気温でのOリングの安全性は保証できないと警告したが、NASAの幹部はこれを無視した。

2003年2月1日、スペースシャトルコロンビア号が発射の際に主翼前縁の強化カーボン・カーボン断熱材が損傷したことにより、大気圏再突入時に空中分解した。地上管制室の技術者たちは損傷の広がりをより明確に把握できるよう、国防総省に対して三回にわたって高解像度の写真を撮影するよう要求し、NASAの熱保護システムの技術主任はコロンビアに搭乗している飛行士たちに耐熱タイルのダメージを調査させるべく船外活動の許可を求めた。NASAの幹部は国防総省の支援の動きに介入してこれを停止させ、船外活動の要求も拒否した"the Columbia Accident". century-of-flight.net。その結果、飛行士が自ら修理に赴くことや、発射準備作業中だったアトランティスで救援に向かうことの実現性は、ついにNASA幹部によって考慮されることはなかった。

 退役
2011年7月8日（日本時間9日未明）に打ち上げられたアトランティスのSTS-135をもって、30年あまりに及んだスペースシャトル計画を終了した。当初の予定では2011年2月26日の打ち上げが最後になる予定だったが、後に追加予算が認められて、非常時の救援ミッションのために待機していたアトランティスをISSの補給ミッションに転用する形で同年7月の打ち上げが認められた。

シャトル退役による宇宙開発計画の間隙を埋めるべく、飛行士や搭載物をISSに運ぶだけでなく、地球を離れて月や火星まで到達できるような宇宙船が現在開発中である。当初「有人開発船（Crew Exploration Vehicle）」と呼ばれていた計画概念は、その後オリオン宇宙船やコンステレーション計画へと発展した。しかし2010年にオバマ政権はコンステレーション計画の予算を打ち切り、今後は低軌道への衛星発射の事業は民間企業に委託することを提案した。次世代の宇宙船が登場するまでは、飛行士がISSに到達しまた帰還するためにはロシア連邦のソユーズ宇宙船か、または開発中のアメリカの民間商用宇宙船に頼る以外に手段がなくなる。オバマ大統領の提案はアメリカ合衆国議会によって承認されたが、次の宇宙船が開発されるまでの5年間にシャトルを延長して使用する可能性を含む対抗案も2010年に議会で検討された。しかし結局、シャトルの退役計画は覆されなかった。

退役後は、ディスカバリーはスミソニアン博物館の国立航空宇宙博物館別館、アトランティスはケネディ宇宙センターの見学者用施設、エンデバーはロサンゼルスのカリフォルニア科学センターにそれぞれ展示される。国立航空宇宙博物館別館に展示中のエンタープライズは、同館にディスカバリーが展示されることに伴い、ニューヨークのイントレピッド海上航空宇宙博物館に移されることになっている。2010年4月、タイム紙は「2010年に最も影響を与えなかった人々」のリストの中にスペースシャトルを挙げ、その理由を「シャトルは従来のロケットのように格好良くないから」としたLeast Influential People of 2010。

 民間商用宇宙船への交代
2008年12月23日、NASAはISSへの物資補給を民間に委ねる商業軌道輸送サービス (COTS) に関する契約を、スペースX社およびオービタル・サイエンシズ社と取り交わしたことを発表した"NASA Awards Space Station Commercial Resupply Services Contracts". NASA, December 23, 2008.。スペースXは2012年にファルコン9ロケットでドラゴン宇宙船を、オービタル・サイエンシズは2013年にアンタレスロケットでシグナス宇宙船を打ち上げ、スペースシャトルに代わってISSへの無人補給ミッションを果たした。

NASAは次いでISSへの有人飛行も民間に委ねるべく商業乗員輸送開発 (CCDev) 計画を開始し、2014年にスペースXのドラゴン2宇宙船とボーイングのCST-100宇宙船を選定した。しかし、有人宇宙船の開発はたびたび遅延を繰り返し、民間によるISSへの有人飛行が実現したのは、スペースシャトル退役から9年後の2020年5月の事であった。

 シャトル訓練機
thumb|C-11A シャトル訓練機（N947NA）
thumb|シャトル訓練機のコックピット。左席に訓練生が座る
シャトル訓練機 (STA) はシャトルの着陸訓練に使用されたアメリカ航空宇宙局の練習機である。グラマン ガルフストリーム IIをベースに4機が改造された。操縦特性が着陸進入時のオービタの挙動と合致するようになっており模擬的に着陸訓練を行うことが出来た。

 開発
外観は飛行訓練中の高い空気力学的荷重に耐えられるように改造されていた。操縦室の左席がオービタの制御と視界を忠実に再現していた。通常の飛行は右席のみで可能となっており、訓練空域までの移動などはこちらで操縦する。

 運用の歴史
4機のSTAが通常はテキサス州エルパソで飛行訓練を行い、ヒューストンで整備を行ったNASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing。STAは同様にフロリダ州のケネディ宇宙センターでも使用された。

 機体
 N944NA (sn144)
 N945NA (sn118)
 N946NA (sn146)
 N947NA (sn147)

 その他の用途
機体後部には数名分の座席が設置されておりT-38が使用できない・人数が多い場合（T-38は2名）に、STAはジョンソン宇宙センターとケネディ宇宙センター間の乗員輸送に使用された。

 ギャラリー

 架空のシャトル一覧
 映画『007 ムーンレイカー』（1979年）より、ムーンレイカー号。
 映画『スペースバンパイア』（1985年）より、チャーチル号。
 テレビドラマ『NASA〜未来から落ちてきた男〜』（1991年）より、フロンティア号。
映画『アルマゲドン』（1998年）より、インディペンデンス号、フリーダム号。
 映画『ゼロ・グラビティ』（2014年）より、エクスプローラー号。

 シャトル派生型打ち上げ機

thumb|200px|サターン V, スペースシャトル, アレス I, アレス V, と アレス IV.の比較
シャトル派生型打ち上げ機Shuttle-Derived Launch Vehicleまたは単純にシャトル派生機Shuttle-Derived Vehicle (SDV) はスペースシャトル計画で開発された技術を基にしたロケットで幅広い機種がこれまで提案されてきた。しかし2011年現在、どれも実用化には至っていない。SDVは既にNASAで複数回、提案されては廃案になるという事を繰り返してきた。1980年代末から1990年代初頭にNASAは公式に貨物専用のシャトル-Cを研究してきた。

2005年にNASAは月と火星の有人飛行計画のためにスペースシャトルの部材を大幅に流用したアレスIとアレスVの開発を決定した。同様に3番目のアレスIVも研究した。2011年4月の時点でNASAはスペースシャトルの代替としてのSDVはスペース・ローンチ・システムである。

 概念
thumb|160px|1978年にチオコール社で提案された直列型シャトル派生機の絵
SDVの概念はシャトル自体が飛行を開始した当時から提案された。SDVの概念には以下を含む:
 有翼のオービタを無人化された使い捨ての貨物ポッドに交換する（"側面搭載型" SDV）
 オービタを取り除いて外部燃料タンクの上部に上段と貨物部を備える（"直列型" SDV）
 かさばる貨物を打ち上げる為に大型の貨物コンテナを外部燃料タンクの後部に備える（後部貨物輸送機）
 固体燃料補助ロケット (SRB) を回収用有翼"フライバック"液体燃料補助ロケットに換装する。
 1本かそれ以上の本数の固体燃料補助ロケットに新開発の上段を開発して載せる。
 耐用回数の末期のオービタから主翼を除いてスペースシャトルの外部燃料タンクを軌道上に投入して組み合わせて宇宙ステーションとして利用する。
 2005年に明らかに前例のない1本の固体燃料ロケット（後に大幅に改良された"延長型"SRB）と新開発の2段目を使用するアレスIが発表された。

これらの案に共通するのは既存のスペースシャトルの構成要素を流用する事で開発費を抑え、より廉価に新型の重量物を軌道に投入する能力を持つ打ち上げシステムを開発しようという意図である。しかし、実際には個々の構成要素は新しい目的別には最適化されておらず、従来の構造体を流用する事によって補強が必要になるなど構造重量の増加の一因ともなり、最適化の障害となっている。有人飛行用としての高度な安全性を備え、再利用を前提としたシステムを使い捨てとして使用しようとした場合、過剰な安全装置等が貨物打ち上げには不要である場合も多い。その為、結局、新技術を盛り込んで最適化された構造の完全新規開発の機体と比較して無駄が多い事は否めず、生産、運用の過程で高くつく可能性が指摘されている。

 シャトルC

thumb|160px|シャトルCの夜間打ち上げの想像図
シャトルCはアメリカ航空宇宙局が提案したスペースシャトルの構成要素を流用した無人貨物打ち上げロケットである。外部燃料タンク (ET) と固体燃料補助ロケット (SRB) とメイン・エンジンを備えた貨物用モジュールを組み合わせて使用される予定だった。複数のシャトルCの概念が1984年から1995年にかけて提案された。

シャトルCの概念は理論的にはシャトル計画で開発された再利用技術によって重量物打ち上げロケットの開発費を減らす事が期待された。提案は複数回行われ、いずれも既存のシャトルの構造体や使用回数限度の迫ったメイン・エンジンや航法コンピュータを流用するというものだった。中にはコロンビア号やエンタープライズ号を1回限りの貨物打ち上げ機として使用する案もあった。チャレンジャー号の事故の前にNASAは年間14回の打ち上げを期待していた。チャレンジャー号の事故の後にはこの打ち上げ頻度は複数の理由により非現実的である事が明らかになった。シャトルCは無人であるので高い打ち上げ頻度でも整備費が安く安全性に関する要求水準が低いと考えられた。

2段階の開発が計画された。第一段階として貨物輸送機の形状と大きさが検討された。NASAによる研究は小型だが最も打ち上げ効率の良い機能的な輸送機を示した。

1990年代初頭、NASAの技術者は火星探査用の宇宙船を組み立てる為に地球周回軌道へ80トンの使い捨ての6機のセグメントを打ち上げる為にシャトルCの設計を含む有人火星飛行計画を立案した。代替案は4機のサターンVを使用する案だった。ブッシュ大統領が2010年にスペースシャトルの運用を終了すると発表した後、これらの提案された仕様は検討対象から外された。

 DIRECT

DIRECTはNASAのビジョン・フォー・スペース・エクスプロレーションで提案されたアレスIとアレスVの代替案として提案された。元のシャトル派生打ち上げ機では"ジュピター"と称され、より野心的な"プロジェクト2"で重量物打ち上げロケットのレビタリアン、軌道周回支援ステーションオリンピア、ガロン重量貨物宇宙船、宇宙ステーションアルゴとヘリオスと乗員貨物船アルテミスから構成され2011年に打ち上げる計画だった。
, DIRECTチームは69人のメンバーで構成されるとされ、 NASAの技術者、コンステレーション計画でNASAと契約した技術者とマネージャー62人から構成され、グループの刊行物によると少数のNASAには属さないメンバーもいる。

計画の名称である"DIRECT"はスペースシャトル計画のハードウェアと施設を"直接"移行する事によって最大限流用する哲学に由来する。

DIRECTには三つの派生機種があり2009年5月に最新の3.0版が発表された。2009年6月17日にワシントンDCで開催された有人宇宙飛行計画委員会の公聴会で明らかになった。

10月11日に2010年のNASAの権限法 (S. 3729) へのオバマ大統領による調印によってスペース・ローンチ・システムが義務化され、DIRECTチームは彼らの努力の成功を宣言した。彼らは新しい宇宙技術企業である: C-Star エアロスペース, LLC.へ組織変更したhttp://www.spacenews.com/policy/101011-obama-signs-nasa-bill.htmlhttp://www.launchcomplexmodels.com/Direct/documents/Direct-Team-Declares-Success-PR-101310.pdf。

 スペース・ローンチ・システム

thumb|left|160px|2011年2月のNASAのスペース・ローンチ・システム (SLS) の仕様
スペース・ローンチ・システム,またはSLSはNASAがコンステレーション計画の中止に伴いスペースシャトルの代替として開発中のシャトル派生型打ち上げシステムの一種である。

2010年NASA権限法によってアレスIとアレスVの機体設計を乗員と貨物輸送の両方に使用できる単体のロケットに一本化する構想である。より強力な機種に更新された。当初の打ち上げ能力は上段を除いたコアのみで構成され低軌道へ70から100トンの投入能力を備える。更に地球離脱段を上段に加えることで130トン以上の打ち上げ能力を獲得する見込みである。

 ジュピター

ジュピターシリーズは現在計画中のDIRECTスペースシャトル派生ロケットの一つである。NASAがコンステレーション計画のために開発していたアレスIとアレスVの代替として企図された。出来るだけスペースシャトルの構成要素や施設を流用する事が予定される。

 注記（出典および脚注）

 参考文献
 NSTS 1988 Reference manual
 How The Space Shuttle Works
 NASA Space Shuttle News Reference - 1981 (PDF document)
 Orbiter Vehicles
 Lecture Series on the space shuttle from MIT OpenCourseWare

 関連項目

 外部リンク

 NASA - Space Shuttle
 NASA - Human Space Flight Web Gallery
 JAXA - 宇宙ステーション・きぼう広報・情報センター - スペースシャトル

*
Category:宇宙船
Category:NASA
Category:アメリカ合衆国のロケット

The space shuttle was incorporated as part of the Space Transportation System (STS) or a space shuttle plan.

Question: what does sts stand for

システム ( Space Transportation System , STS ) あるいは スペース シャトル

The space shuttle was incorporated as part of the Space Transportation System (STS) or a space shuttle plan.

Question: what is another name for the sts

スペース シャトル 計画

Question: what was incorporated into the sts plan

"スペース シャトル は

Question: what is the name of the space transport system

what is the name of the space transport system

Question: what is the acronym for the space transportation system

what is the acronym for the space transportation system

The space shuttle was incorporated as part of the Space Transportation System ( STS ) or a space shuttle plan .

スペースシャトルは宇宙輸送システム (Space Transportation System, STS) あるいはスペースシャトル計画の一環としてもちいられた。

スペースシャトルは宇宙輸送システム (Space Transportation System, STS) あるいはスペースシャトル計画の一環としてもちいられた。

The space shuttle was incorporated as part of the Space Transportation System (STS) or a space shuttle plan.

The space shuttle was incorporated as part of the Space Transportation System (STS) or a space shuttle plan.

The Shuttle is a re-use spacecraft and consists of three parts of the orbitor vehicle (OV), the external fuel tank (ET) and the Solid Rocket Booster (SRB).

Length
Space Shuttle: 184 feet
Orbiter: 122 feet
Height
Orbiter on runway: 57 feet
Wingspan
78 feet
Liftoff Weight*
4.5 million pounds
Orbit
115 to 400 statute miles
Velocity: 17,321 mph
* weight will vary depending on payloads and on board consumables.
+ More NASA Facts...
|Shuttle Basics|
|
The Space Shuttle is the world's first reusable spacecraft, and the first spacecraft in history that can carry large satellites both to and from orbit. The Shuttle launches like a rocket, maneuvers in Earth orbit like a spacecraft and lands like an airplane. Each of the three Space Shuttle orbiters now in operation -- Discovery, Atlantis and Endeavour -- is designed to fly at least 100 missions. So far, altogether they have flown a combined total of less than one-fourth of that.
|
Image left: The components of the Space Shuttle system: Orbiter, External Tank, and Solid Rocket Boosters. Photo Credit: NASA
Columbia was the first Space Shuttle orbiter to be delivered to NASA's Kennedy Space Center, Fla., in March 1979. Columbia and the STS-107 crew were lost Feb. 1, 2003, during re-entry. The Orbiter Challenger was delivered to KSC in July 1982 and was destroyed in an explosion during ascent in January 1986. Discovery was delivered in November 1983. Atlantis was delivered in April 1985. Endeavour was built as a replacement following the Challenger accident and was delivered to Florida in May 1991. An early Space Shuttle Orbiter, the Enterprise, never flew in space but was used for approach and landing tests at the Dryden Flight Research Center and several launch pad studies in the late 1970s.
The Space Shuttle consists of three major components: the Orbiter which houses the crew; a large External Tank that holds fuel for the main engines; and two Solid Rocket Boosters which provide most of the Shuttle's lift during the first two minutes of flight. All of the components are reused except for the external fuel tank, which burns up in the atmosphere after each launch.
The longest the Shuttle has stayed in orbit on any single mission is 17.5 days on mission STS-80 in November 1996. Normally, missions may be planned for anywhere from five to 16 days in duration. The smallest crew ever to fly on the Shuttle numbered two people on the first few missions. The largest crew numbered eight people. Normally, crews may range in size from five to seven people. The Shuttle is designed to reach orbits ranging from about 185 kilometers to 643 kilometers (115 statute miles to 400 statute miles) high.
The Shuttle has the most reliable launch record of any rocket now in operation. Since 1981, it has boosted more than 1.36 million kilograms (3 million pounds) of cargo into orbit. More than than 600 crew members have flown on its missions. Although it has been in operation for almost 20 years, the Shuttle has continually evolved and is significantly different today than when it first was launched. NASA has made literally thousands of major and minor modifications to the original design that have made it safer, more reliable and more capable today than ever before.
Since 1992 alone, NASA has made engine and system improvements that are estimated to have tripled the safety of flying the Space Shuttle, and the number of problems experienced while a Space Shuttle is in flight has decreased by 70 percent. During the same period, the cost of operating the Shuttle has decreased by one and a quarter billion dollars annually -- a reduction of more than 40 percent. At the same time, because of weight reductions and other improvements, the cargo the Shuttle can carry has increased by 7.3 metric tons (8 tons.)
In managing and operating the Space Shuttle, NASA holds the safety of the crew as its highest priority.

Question: what is the et in the shuttle

The Shuttle is a re-use spacecraft and consists of three parts of the orbitor vehicle (OV), the external fuel tank (ET) and the Solid Rocket Booster (SRB).

Question: how many parts does the shuttle have

Question: what does srb stand for

Question: what is the solid rocket Booster called

Question: what does ov stand for in the shuttle

Question: what are the parts of the shuttle called

what are the parts of the shuttle called

Question: what is the name of the re-use spacecraft

what is the name of the re-use spacecraft

The Shuttle is a re - use spacecraft and consists of three parts of the orbitor vehicle ( OV ) , the external fuel tank ( ET ) and the Solid Rocket Booster ( SRB ) .

シャトルは再使用型宇宙往還機であり、軌道船 (Orbitor Vehicle, OV)、外部燃料タンク (External Tank, ET)、固体燃料補助ロケット (Solid Rocket Booster, SRB) の三つの部分によって構成されているShuttle Basics. NASA.。

シャトルは再使用型宇宙往還機であり、軌道船 (Orbitor Vehicle, OV)、外部燃料タンク (External Tank, ET)、固体燃料補助ロケット (Solid Rocket Booster, SRB) の三つの部分によって構成されているShuttle Basics. NASA.。

The Shuttle is a re-use spacecraft and consists of three parts of the orbitor vehicle (OV), the external fuel tank (ET) and the Solid Rocket Booster (SRB).

The Shuttle is a re-use spacecraft and consists of three parts of the orbitor vehicle (OV), the external fuel tank (ET) and the Solid Rocket Booster (SRB).

The shuttle’s composition at the time of launch was mostly colored in white when the external fuel tank (ET) was first launched (Columbia), and the external fuel tank was also painted in white.

In addition to the above, using two kinds of solid rocket for artificial satellite launch, known as PAM and IUS, developed for STS, we could also carry the equipment to an even higher orbit.

Question: what color was the external fuel tank of the columbia space shuttle

The shuttle’s composition at the time of launch was mostly colored in white when the external fuel tank (ET) was first launched (Columbia), and the external fuel tank was also painted in white.

In addition to the above, using two kinds of solid rocket for artificial satellite launch, known as PAM and IUS, developed for STS, we could also carry the equipment to an even higher orbit.

what color was the external fuel tank of the columbia space shuttle

Question: what were the names of the two types of solid rocket used to launch the space shuttle

PAM と IUS と 呼ば れる 人工 衛星 打上げ 用 の 2 種類 の 固体 ロケット を 用いれ ば 、 搭載

what were the names of the two types of solid rocket used to launch the space shuttle

Question: what was the name of the first space shuttle launched

what was the name of the first space shuttle launched

Question: when was the space shuttle first launched

when was the space shuttle first launched

Question: what was used to insulate the external tank of the columbia space shuttle

what was used to insulate the external tank of the columbia space shuttle

Question: how many parts were in the orbiter vehicle of the space shuttle

how many parts were in the orbiter vehicle of the space shuttle

The shuttle ’s composition at the time of launch was mostly colored in white when the external fuel tank ( ET ) was first launched ( Columbia ) , and the external fuel tank was also painted in white .

" NASA Takes Delivery of 100th Space Shuttle External Tank " .NASA , August 16 , 1999 . Quote : " ... orange spray - on foam used to insulate ... ""Media Invited To See Shuttle External Fuel Tank Ship From Michoud" . NASA , December 28 , 2004 . Quote : " The giant , rust - colored external tank ... " Two white and thin solid fuel assistant rocket ( Solid Rocket Boosters , SRB ) consisted of three parts of the orbiter Vehicle ( OV ) of the astronauts and cargo .

In addition to the above , using two kinds of solid rocket for artificial satellite launch , known as PAM and IUS , developed for STS , we could also carry the equipment to an even higher orbit .

発射時のシャトルの構成は、おおまかに オレンジ色の外部燃料タンク (External Tank, ET)最初の打ち上げ（コロンビア号）の際は、外部燃料タンクも白色に塗装されていた。 "NASA Takes Delivery of 100th Space Shuttle External Tank" . NASA, August 16, 1999. Quote: "…orange spray-on foam used to insulate…""Media Invited To See Shuttle External Fuel Tank Ship From Michoud". NASA, December 28, 2004. Quote: "The gigantic, rust-colored external tank…" 2本の白色で細長い固体燃料補助ロケット (Solid Rocket Boosters, SRB) 宇宙飛行士と貨物を搭載する軌道船 (Orbiter Vehicle，OV) の三つの部分から構成されていた。 なお、上記に加えて、STSのために開発された、PAMとIUSと呼ばれる人工衛星打上げ用の2種類の固体ロケットを用いれば、搭載物をさらに高い軌道に運ぶこともできた。

発射時のシャトルの構成は、おおまかに オレンジ色の外部燃料タンク (External Tank, ET)最初の打ち上げ（コロンビア号）の際は、外部燃料タンクも白色に塗装されていた。

なお、上記に加えて、STSのために開発された、PAMとIUSと呼ばれる人工衛星打上げ用の2種類の固体ロケットを用いれば、搭載物をさらに高い軌道に運ぶこともできた。

The shuttle’s composition at the time of launch was mostly colored in white when the external fuel tank (ET) was first launched (Columbia), and the external fuel tank was also painted in white. "NASA Takes Delivery of 100th Space Shuttle External Tank" .NASA, August 16, 1999. Quote: "...orange spray-on foam used to insulate...""Media Invited To See Shuttle External Fuel Tank Ship From Michoud".NASA, December 28, 2004. Quote: "The giant, rust-colored external tank..." Two white and thin solid fuel assistant rocket (Solid Rocket Boosters, SRB) consisted of three parts of the orbiter Vehicle (OV) of the astronauts and cargo. In addition to the above, using two kinds of solid rocket for artificial satellite launch, known as PAM and IUS, developed for STS, we could also carry the equipment to an even higher orbit.

The shuttle’s composition at the time of launch was mostly colored in white when the external fuel tank (ET) was first launched (Columbia), and the external fuel tank was also painted in white.

In addition to the above, using two kinds of solid rocket for artificial satellite launch, known as PAM and IUS, developed for STS, we could also carry the equipment to an even higher orbit.

The exterior shell is made of 13mm thick steel.

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Question: how much fuel does solid fuel aligned rockets generate

The exterior shell is made of 13mm thick steel.

how much fuel does solid fuel aligned rockets generate

Question: what is the exterior shell of the rocket made of

what is the exterior shell of the rocket made of

Question: how high was the rocket when it was separated

how high was the rocket when it was separated

Question: how thick is the exterior shell of the nasa rocket

how thick is the exterior shell of the nasa rocket

Question: when was the rocket separated from the earth

when was the rocket separated from the earth

Question: what does srbs stand for

"固体 燃料 補助 ロケット 固体 燃料 補助 ロケット ( SRB ) は 2 基 合計 で 発射 時 に 必要 と さ れる 推力 の 83 % 、 約 1,250万 ニュートン （ 1,276.8 トン ） を 発生 し Columbia Accident Investigation Board Report , Vol II , Appendix D . 7 . NASA , October 2003 . 、 打上げ から 2 分 後 、 高度 約 15万 フィート ( 46 km ) に 達し た ところ で 切り 離さ れ 、 パラシュート で 海 に

Question: how much fuel does solid fuel assured rockets generate

how much fuel does solid fuel assured rockets generate

Solid Fuel Assisted Rockets Solid Fuel Assisted Rockets ( SRBs ) generate approximately 83 % of the drive needed at the launch at 2 bits total , about 1.2 million newton ( 1,276.8 tons ) and Columbia Accident Investigation Board Report , Vol II , Appendix D.7 . NASA , October 2003 , two minutes after the launch , is separated at a height of approximately 150,000 feet ( 46 km ) and watered and recovered in the sea with a parasut .

The exterior shell is made of 13 mm thick steel .

固体燃料補助ロケット 固体燃料補助ロケット (SRB) は2基合計で発射時に必要とされる推力の83%、約1,250万ニュートン（1,276.8トン）を発生しColumbia Accident Investigation Board Report, Vol II, Appendix D.7 . NASA, October 2003.、打上げから2分後、高度約15万フィート(46km)に達したところで切り離され、パラシュートで海に着水して回収される。 外殻は厚さ13mmの鋼鉄でできている。

固体燃料補助ロケット 固体燃料補助ロケット (SRB) は2基合計で発射時に必要とされる推力の83%、約1,250万ニュートン（1,276.8トン）を発生しColumbia Accident Investigation Board Report, Vol II, Appendix D.7 . NASA, October 2003.、打上げから2分後、高度約15万フィート(46km)に達したところで切り離され、パラシュートで海に着水して回収される。

外殻は厚さ13mmの鋼鉄でできている。

Solid Fuel Assisted Rockets Solid Fuel Assisted Rockets (SRBs) generate approximately 83% of the drive needed at the launch at 2 bits total, about 1.2 million newton (1,276.8 tons) and Columbia Accident Investigation Board Report, Vol II, Appendix D.7 . NASA, October 2003, two minutes after the launch, is separated at a height of approximately 150,000 feet (46 km) and watered and recovered in the sea with a parasut. The exterior shell is made of 13mm thick steel.

The exterior shell is made of 13mm thick steel.

[L109] Various analyses were performed in support of the investigation of the STS 51-L incident. Included in this appendix are the following analysis summaries:
A. Structural Analyses of SRM Segments, Field joints, and Seals
B. SRM Temperatures, Prelaunch and Launch
C. Assessment of joint Leakage from Near Ignition to 58 Seconds
D. Flow and Thermal Analyses
Section C considers the events between the near-ignition time frame when the smoke puff was observed, and the observed hot plume after 58 seconds.
A. Structural Analyses for SRM Segment, Field joints, and Seals
1. Organization/Responsibilities
This appendix deals with all analyses relative to the SRM field joint and seals in terms of structural capability, mating characteristics, O-ring seal gap opening, O-ring response, and fracture mechanics. Additional loads and analyses are presented in the Systems Working Group Final Report.
Various MSFC and contractor personnel contributed to the successful completion of Appendix B. The overloads reconstruction tasks were conducted by MSFC, JSC, and Rockwell Space Division and are reported in the Systems Working Group Final Report.
2. Method of Investigation
The method followed for these analyses was: (1) determine all events important to joint/seal analysis, (2) determine or acquire natural and induced environments for these events, (3) develop appropriate structural models to characterize the hardware for the event, (4) perform static, dynamic, and capability analysis using models and event environments, and (5) evaluate results in terms of findings and conclusions. Figure B.1 is a list of the events analyzed starting with SRM segment mating and ending with the max Q flight event. Figure B.2 breaks down further the overall approach while Figure B.3 depicts the overall flow including special 51-L geometric dimensions. The SRM joint Deflection Reference Test (MTI 103) was used to fine tune and verify the various models. Because of the complexity of the structural configuration and the stringent prediction accuracy requirements, several different models, parameter effects, etc., were required in order to meet the objectives of the investigation. Figure B.4 is a listing of the major models developed and used.
The environments were determined from telemetered data, flight event reconstruction, special analyses, and measured natural environments. The flight event reconstruction determines all external loads environments and is discussed in the Systems Working Group Final Report.
Simplification of the overall analysis was achieved by four special studies which allowed reducing the model and approach to a quasi-static or static analysis, using simplified pin/hole constraints, and propellant as mass only. These analyses were: (1) dynamic effects of bending and shell modes on field joint response, (2) effect of elastic propellant on field joint response, (3) nonlinear iteration of joint, pin boundary to check the simplified pin/hole constraint assumptions, and (4) O-ring response analysis. The results of these analyses will be discussed along with the overall field joint analysis results.
3. Terms and Acronyms
The terms and acronyms used in this appendix are identified in the main body of the report.
4. Narrative Description
This section of the report will discuss the assumptions, results, and findings that contribute to the field joint response. Special studies conducted to verify assumptions and determine unique effects will each be discussed. Next, the basic joint response will be formulated and the combined behavior will be determined. Special capability studies will also be discussed. Finally, the findings and conclusions will be stated.
4.1 Dynamic Effect of Bending and Shell Modes on Field Joint Opening
Bending and shell modes were developed for the 360° one-diameter forward and aft segment models including models of the clevis, tang, and pins. Figure B.5 shows the finite element models for the clevis, tang, aft ET attach ring and the forward and aft one-diameter length segments. Approximately 150 modes were calculated for a frequency range of 7 Hz thru 7,000 Hz. The dynamic response was run for the SRM ignition time portion of the lift-off transient and compared to a time point static analysis. The difference in shell deflection and field joint gap deflection between the static and dynamic analyses was negligible. As a result of these analyses it was concluded that a quasi-static analysis was adequate for determining gap opening as a function of time to the various induced and internal loads acting on the SRM case and joints.
4.2 The Effect of Elastic Propellant on Field joint Response
(a) Model Description
The SRB aft field joint was modeled using EAUSPAR finite element code. The math model defines a total of 220 nodes in a symmetric 4° segment of the motor. The propellant grain is [L110] modeled with 80 Solid elements while the steel motor case is represented by 19 quadrilateral plate elements. The model includes 116 inches of the motor length such that the effects of the end boundary conditions are not significant near the field joint. Figure B.6 gives a graphic representation of the propellant/case model. Note in particular joint 91, where the MACRO and plate elements join. The model is used to determine the static and dynamic effects of the elastic propellant on field joint response.
(b) Applied Loading
The pressure time history function used in the present analysis is shown in Figure B.7. The maximum pressure is 768 psi at t = 0.652 seconds. The pressure time trace defines the internal pressure applied to the inside bore of the propellant grain. This pressure function was reconstructed from hoop strain gage data (TWR 13852 and TWR 14914) taking into account the propellant radial stiffness effects and ballistic model results. For the purpose of understanding the influence of the propellant elasticity and damping characteristics of field joint response, a second analysis was performed using the same pressure function applied directly to the steel motor case model without the propellant. The math model is only capable of predicting the response to zero harmonic loading; therefore, no strut loads were considered. Previous analysis has shown that the case response is dominated by the internal pressure in the motor.
(c) Propellant Damping
Propellant damping is nonlinear; the function is dependent upon temperature, strain, and frequency. Figure B.8 shows the propellant damping properties as a function of temperature and frequency. The data points are given in terms of equivalent linear damping where percent damping equals one-half of the loss modulus/storage modulus ratio. Notice, as temperature drops percent damping decreases with less dependency on frequency. In general, propellant damping is high, greater than 20 percent for all cases. For the analysis reported herein, 30 percent damping is assumed for the propellant and 1 percent damping for the structural shell. The effects of propellant structural damping on field joint response are clearly depicted in Figures B.9 through B.11. Acceleration, velocity, and displacement responses of Joint 91 are plotted with and without propellant. Joint 91 response is typical of all positions along the case shell. As evidenced from these plots, propellant damping filters out much of the high frequency noise in the velocity and acceleration response, but has virtually no effect on the displacement response (amplitude or time).
(d) Static and Dynamic Response Calculations
Dynamic response analysis of the field joint was performed using the method of modal superposition. A total of 250 vibration modes were used in the summation, and response quantities were computed at several locations on the motor case. Figure B.11 shows the time history for the radial component of displacement for Joint 91 on the shell. The maximum value of the shell displacement for this analysis was 0.357 inch. When the same value for the maximum pressure was applied statically, the radial deflection was also 0.357 inch. Similar results were found for other response quantities. Thus for this particular model and load time history, no real difference was noted between the dynamic and static response when the dynamic load was applied quasi-statically.
Four different model configurations were analyzed for a static pressure load of 768 psi. The results of these cases are summarized below in Table B.1. Notice, propellant elasticity (assumed to be constant at 800 psi) has minimal effect on shell radial growth and primary gap opening. Compare Case 1 to 2, and Case 3 to 4.
(e) Findings
The findings apparent from this investigation related to the structural behavior of the SRM aft field joint are:
(1) The elasticity and damping effects of the solid propellant on the field joint response are small. The structural behavior of the SRM is dominated by the steel motor case.
(2) The dynamics of the problem do not yield results that differ from a quasi-static analysis using dynamic loads. The displacement response virtually tracks the applied load, with almost no time lag.
4.3 Nonlinear Iteration of joint, Pin, etc.
Using ANSYS, Thiokol built a finite element model of the web, pin, tang, and a length of the membrane. ANSYS allows the assumption of gaps between the pin and holes, introduction of friction forces, allowing the elements to contact and move or lock as a function of the applied loads. The analysis was run for the ignition pressure buildup case allowing the model to iterate at each step and acquire a position of the pin relative to the hole elements. This case was compared to an analysis where the pin was frozen (locked) in the tang and clevis and ran as a linear analysis. Figure B.12 shows the model and the radial deflection for the linear run. Also shown is the displacement as various colors. Displacements are shown overall and locally in the clevis/tang. Figure B.13 shows the force vectors and displacements acting on the pin with the clevis contact shown at the top and the tang at the bottom. Internal displacement amplitude distribution is shown by the various colors. The color sequence from minimum to maximum displacement is given as dark blue, light blue, green, yellow, and red. Figure B.14 and Figure B.15 are the same plots for the nonlinear analysis. The pin hole contact locked at 200 psia internal pressure. The position and resulting forces and stress field are very close to those assumed for the linear case. This verifies that the engineering assumption made to arrive at gap openings was reasonable and adequate.
In addition, the gap opening was compared to the results obtained by the MSFC analysis. Overall gap opening was 0.034 inch MSFC vs. 0. 031 inch Thiokol. The MSFC model used 3D solid elements while the Thiokol model was axisymmetric. The split between axial force and pressure force (radial) was 15 percent and 85 percent respectively for both models. The nonlinear model untuned gave a gap of 0.040 inch and a 10 to 90 percent split between axial and pressure forces. As a result of this study, a linear locket pin model was used for all gap response analysis. Errors introduced by this assumption should not exceed two mils at maximum gap opening.
Case
Model Configuration
Shell Radial Growth (inch)
Primary Gap Opening (inch)
.
1
Pressure Load Acting Directly on the Steel Motor Case, Propellant Included, No Axial Effects Considered
2
Pressure Load Acting Directly on the Steel Motor Case, No Propellant, No Axial Effects Considered
3
Pressure Load Acting on the Inside Bore of the Propellant Grain, Axial Effects Considered
4
Pressure Load Acting on the Steel Motor Case, No Propellant, Axial Effects Considered
[L111] 4.4 O-Ring Response Analysis
An O-ring type sea] has two mechanisms for sealing in an application like the SRM field joints where joint rotation (gap opening) occurs as pressure increases: (1) the resiliency of the seal is sufficient to allow the seal to expand and track the gap opening maintaining a sealing surface, (2) free-floating seal translated into a sealing position by pressure forces in particular flow across the seal through the gap. In practice it is highly desirable to depend on the resiliency and not the free-floating pressure actuated approach. This section will analyze both cases.
(a) Model Assumptions
The O-ring seal was modeled with wedged-shaped solid elements using the finite element code SPAR. The model is limited to a 0.1 segment of the 6-foot-radius O-ring with the proper axisymmetric boundary conditions, and the seal material is characterized as linear elastic. Actually, the Viton seal material is highly nonlinear and has a large time dependency. Typical O-ring relaxation and resiliency data are plotted in Figure B.16. Nonlinear stiffness due to large strains and corresponding deflections and nonlinear material behavior are not defined in the present seal model. An axisymmetric model will not allow various sectional seal initial states on circumferential cross talk. By its nature this approach assumes the same behavior for the total 360° of the seal. Although linear elastic analysis has major limitations for the analysis of the highly nonlinear O-ring, an important understanding of the seal dynamic response is obtained from the results.
(b) Enforced Displacements
A two-dimensional slice of the O-ring cross section positioned inside of the O-ring groove is shown in Figure B.17. The O-ring displacement shapes for three different states of compression are also given. A nominal diameter O-ring (0. 280 inch) is compressed in a maximum depth groove (0.216 inch). A tang-clevis contact boundary, causes a steel-to-steel compression of the O-ring of 0.064 inch, and the O-ring seals on the sides of the O-ring groove. In the fully compressed state, flow around the O-ring is restricted. As the gap opens during the SRM ignition transient, O-ring resiliency and the tang-clevis/O-ring flow boundary characterize the seal response.
(c) Response of Fully Compressed O-Ring
The transient response of the seal after steel-to-steel compression of the tang and clevis was investigated using an uncoupled linear analysis scheme employing superposition of vibration mode shapes. Thirty modes of the O-ring cross section were used from 2,000 to 8,000 Hz. Pressure was applied to the O-ring cross section along the top 180 degrees. The sides of the O-ring were considered sealed. A linear representation of the response is presented in Figure B.18 for release of the fully compressed O-ring. The response plot with a pressure of 936 psi shows that the pressure drives the seal farther into the groove locking the O-ring in this position. For this case, tang-clevis gap sealing cannot occur. The linear assumptions show the seal being driven through the bottom of the groove while in fact this is not possible. The important and correct result is that the seal does not respond to any gap opening and is compressed into the groove by pressure. Additional response plots for the initial steel-to-steel compression of the O-ring are presented for pressures of 468 and 374 psi. The overall trend is to force the seal into the groove. These plots show a slight initial tendency to respond up. This is due to the natural tendency for the top enforced flat on the O-ring to return to a circular configuration. The exact transition pressure between holding the seal in and allowing the sides to unseal is hard to obtain realistically from this analysis. A pressure less than the transition pressure would allow pressure to get under the seal to allow it to respond as the pressure increases. The exact determination of the transition pressure requires a detailed nonlinear analysis taking into account nonlinear properties which vary with time and temperature. The O-ring steel-to-steel compression response plots have a constant pressure load over the top 1800, which would not be the case the instant the seal sides unseal. The plots shown in Figure B.18 investigate the tendency of the 0ring to lock in the groove. Note, seal damping is assumed to be equal for all pressure states given at 2 percent.
(d) Response of O-ring Compressed 0.046 Inch
As in the previous response investigation, linear elastic uncoupled modal transient analysis techniques were used. The initially compressed 0.046-inch seal was modeled with model superposition. This study assumes the tang-clevis combination to instantaneously be separated by 0.048 inch (Figure B.19). Sonic conditions at 400 psi were assumed at the top of the O-ring and the pressure distribution around the O-ring was calculated from area ratios. The downstream side of the O-ring was assumed to have separated flow, and the middle bottom of the O-ring was assumed sealed on the bottom of the groove. The flow velocity between the tang and clevis was 1,038 ft/sec and the velocity over the 0ring top was 2,481 ft/sec. The one-dimensional flow analysis assumed has limitations that have not been test verified. The degree that the Bernoulli induced forces are present for this application are not known. Tests indicate strongly that pressure can actuate seals if initial seal compression is not too large. These test results do not, however, quantitatively verify assumptions made. The first response plot, Figure B.20, shows the linear elastic response of the 0.046-inch compressed O-ring without pressure. Figure B.21 shows the O-ring response of the pressure loading defined by the 0.048-inch tang-clevis separation. This pressure loading state speeds up the response. Although the linear elastic analysis shows very fast response, the nature of the material suggests this is not the case. These plots demonstrate that the dynamic inertia forces are very small.
If in fact the material is slow to restore to its original undeformed shape, it becomes necessary to depend on other sealing phenomenon for quick reaction. An analysis was then done allowing the seal to be free floating to see if it would translate into a sealed position. This was done using the mode shapes of a free O-ring cross section. The first free cross section mode was 4,000 Hz. The response plot given in Figure B.22 shows that the 0ring is capable of free floating into position at a faster rate than the tang/clevis gap opens. Note that this type of pressure sealing is independent of flow material response.
(e) Seal Groove Tolerance Effects
The analyses performed in subsections (c) and (d) showed the criticality of three factors: (1) groove dimensions, (2) seal dimensions, and (3) initial gap size (O-ring initial compression) on O-ring sealing capability at low resiliency. This section analyzes the various gap and O-ring dimensional variations arriving at initial minimum gap characteristics where seal touches groove sides. This position in combination with low resiliency (low temperature) greatly retards the ability to reach a sealing position.
Figure B.23 is the result of combining a nominal O-ring diameter with an average groove assuming 2.5 angle on sides. At a gap of 0.004 inch the seal touches that groove sides. Figure B.24 shows the condition for a maximum O-ring diameter and a minimum groove. In this case the seal touches the groove sides for a gap of 0.032 inch. If the combination of a nominal C-ring and maximum groove is assumed, the O-ring will never touch the groove sides (Figure B.25).
Taking the measured groove dimensions for the right. SRM aft field joint primary (0.2105 depth, 0.307 width, side angle 1), secondary (0. 2 101 depth, 0. 306 width, side angle 1) and the range of O-ring variations possible (not measured) leads to the following results:
.
Measurements of the other field joint gaps showed them to be within ± 0.003 inch of the above groove dimensions. The expected gap at which wall-to-wall contact is made would therefore range [L112] from 0 to 0.025 inch. The critical dimensions determining whether a seal is in an undesirable initial condition are: (1) O-ring diameter, and (2) initial gap characteristics (O-ring squeeze).
(f) Findings
(1) Initial forces necessary to compress O-ring steel-to-steel (0,064 inch).
Fluoroelastomer, psi
Temperature, °F
Compression Force, lbs/in
.
1,240
70
52.8
1,970
40
83.9
3,170
25
135.0
16,400
10
698.0
Note: The temperature/modulus relationship is based on early test data. (Fluoroelastomer = Viton-E/fluorile)
(2) Steel-to-steel compression of the O-ring results in contact and sealing on the sides of the 0-ring groove. This side groove sealing can couple with pressure forces to hold the seal in a fully compressed state making pressure sealing impossible.
(3) If the seal is not sealed on the sides of the groove (as in the steel-to-steel case), the Bernoulli induced flow forces are sufficient (based on one-dimensional flow assumptions of this analysis) to rigid body translate the O-ring to the sealed position faster than the gap opening transient. This rigid body translation of the seal to a sealing position is known as a free-floating seal. It should be noted that according to the Parker Seal Handbook, it is not recommended that free-floating seals be used for pressures over 200 psi. Obviously, a free-floating actuated seal is highly susceptible to flow induced oscillations.
(4) Although linear elastic theory predicts a very fast snapback of the seal from the compressed state, the extreme time dependency of the Viton material makes response from a highly compressed state too slow to follow the gap opening transient. The seal must depend on a free-floating translation for quick sealing of large gaps which is not the most desirable and reliable approach.
4.5 SRM Field Gap Response for Various Flight Events
The SRM field gap opening as a function of internal and external forces has been analyzed for all flight events in terms of delta gap opening (gap change from any initial state). As stated previously, several finite element models were developed and analyzed because of the complexity of the problem. In generally they generically match the tang and clevis portion shown previously on Figure B.5. The resulting analysis using these models can be classified in two broad categories: (1) static analysis of a narrow width 2 to 4 model of the clevis tang, pin, and segment membrane of 50-inch to 70-inch length depending on the individual choice, and (2) quasi-static analysis using the 360, and one-diameter length segment models with and without the ET attachment ring. The static analysis model, Category 1, is capable of line load and internal pressure only, while the quasi-static model, Category 2, can handle simultaneously pressure, line, and external forces. The quasi-static analysis serves as the base model since external forces can be applied.
(a) Static Analysis of Gap Response
As stated, several static models have been developed. The models were adjusted slightly so that they match the 103 joint Deflection Test conducted at Thiokol. All adjustments were within the variation expected for materials, properties, etc. Figure B.26 is a plot of the strain data versus pressure for various gage locations used for correlating model simulated strains. Figure B.27 shows the gap opening measured at the midpoint of the ridge between the primary and secondary O-ring measured circumferential. Figure B.28 shows gap opening for several pressures versus linear voltage displacement transducer (LVDT) locations or tang/clevis/pin location. Agreement between model and test was achieved within 2 mils at maximum pressure (gap opening) using the ignition pressure buildup at lift-off. A static analysis of the primary and secondary seal aft field joint was run assuming the primary seal does seal, then assuming it does not seal and pressure gets between the primary and secondary seal. Finally, the effect of the alignment gap was analyzed. Table B.2 summarizes the results presented in Figures B.29 through B.34.
.
Primary Seal Gap
Secondary Seal Gap
The effective peak pressure on the aft field joint of 7 70 psi was used in this analysis. Strut loads were not considered in this simplified model. Observations indicate that the basic model matches the test results well. That is, if the primary seal leaks the gap opening increases due to pressure on the secondary seal and the larger clevis tang area, and the area around the alignment slot has an increased gap opening of at least 2 or 3 mils.
(b) Quasi-Static Gap Opening Analysis for Flight Events
The model used for this analysis was shown in Figure B.5 and has the capability of handling external and internal loads as a function of time. These loads were reconstructed using the procedure shown on Figures B.35 and B.36 developing a time-dependent set of functions. Figure B.37 is a time response of two of these reconstructed loads of the aft field joint. Shown in this figure are the key lift-off events: the axial load and the pitch plane bending moment. The major load is the internal pressure of the SRM. Clearly evident is the structural elastic body ring out when the SSME buildup introduced energy is released.
Table B.3 lists the design limit load and the results of the 51-L loads reconstruction for all field joints. Since all field joints are designed to the highest axial load (forward field joint) and the gap opening is a function of this axial load in conjunction with the equivalent internal pressure and external loads, the forward field joint should experience the highest load and field joint seal gap opening. Effects of all of these combined loads will be evident in the data that follows.
Figures B.40 through B.52 are the delta gap response plots for all the flight events/SRM for-ward and aft field joints for 51 -L. For reference, the SRB attach member (strut) identification and the Shuttle Coordinate System shown in Figures B.38 and B.39, respectively. Both a time response for the maximum circumferential is shown in conjunction with the peak plotted for the total 360° around the joint. The highest delta gap opening occurs for the forward field joint primary seal (37 mils) at lift-off while the aft joint peaks near 28 mils. All joint delta gap openings peak at lift-off. The primary seal delta gap opening is approximately a factor two higher than the secondary seal gap. The mid-field joint gap (not shown) is slightly lower than the forward field joint.
The aft field joint, because of the external loads applied through the struts, shows smaller dynamic activity (Figure B.40). Here the approximately 3 Hz vehicle mode produces a small delta gap opening of ± 2 mills. A max/min for any aft joint location is plotted on Figure B.41 and shows about a 3-mil variation circumferential. Max Q gap openings are approximately 30 percent less than lift-off; however, the external loads influence is greater.
The aft field joint was also used to check the effect of installation alignment slots on the seal gap opening. Figures B.33 through B.55 show this effect. Approximately 2-mil gap opening increase is observed. A blowup of this area is shown in Figure B.54 with the increase and the influence of the alignment slots on gap opening. The gap opening at the alignment slots is also apparent in the MTI 103 joint Deflection Test data (Figure B.55).
Delta field gap opening is a very dynamic phenomenon at SRB ignition with a total travel of 20 to 30 mils within 600 milliseconds. All other transient gap conditions occur at a much slower rate.
4.6 SRM Structural Capability
In this section, the structural capability of the SRM is assessed in terms of: (a) mating loads and resulting stresses on the [L113] aft field joint. (b) effect of partial joint failure, and (c) effect of local membrane holes.
(a) Mating Analysis
The analysis to assess the stresses and deflections during mating -if the aft center segment to aft segment is scheduled to be completed by April 13. As part of this analysis we have:
(1) Determined the point load required to deflect the suspended segment the width of the chamfer on the tang (0.2 inch).
(2) Set the tang on the inner leg of the clevis at one point and allowed the segment to rotate until the tang on the opposite side 180 degrees away was almost resting on the outer leg of the clevis.
(3) Determined the loads required to mate on ovalized tang and clevis.
(4) Determined the maximum bending stress that resulted from use of the SRB circumference alignment tool during mating of :he right aft center segment of 51-L.
The radial load required to deflect the suspended segment 0.2 inch as in (1) above is only a low 75 pounds because the overhead crane cable is over 640 inches long. The axial load applied to the clevis inner leg by the loading in item (2) above is 8,000 pounds. The loading required to mate an ovalized tang and clevis is currently in progress. Initial indicators are that the peak radial force per circumferential inch required to mate the two segments is less than 100 pounds per inch. The maximum load applied radially inward using the alignment tool was 4,000 pounds. Calculations using Flexible Circular Ring Analysis resulted in a bending stress of 59,656 psi. Safety factors for the bending stress are 3.02 on yield and 3.2 7 on ultimate. The analysis is conservative because the stiffness of the propellant is not considered. The mating scenario has also been analyzed from a fracture mechanics viewpoint (Section 4.7). The conclusion reached from the stress and fracture mechanics analysis is that the mating loads caused no damage to the joint.
(b) Effect of Partial joint Failure
As a part of the investigation, the length of the aft field joint failure around the circumference that could be failed without an overall joint failure was determined for various loadings and temperatures of the unfailed joint.
In Figure B.56 the circumferential opening (W) is the sum of all joint lengths around the circumference of the field joint. In this figure the maximum failed joint length is 315 inches with .he unfailed joint at room temperature. The example shows this failed length decreases to 273 inches when the temperature is raised 800°F. In Figures B.57 and B.58 the failed joint length is continuous. Figure B.57 considers both the axial load and the vehicle bending moment in determining the allowable failed joint length. The length is 158 inches with the remainder of the joint at room temperature. Figure B.58 considers the axial load and the internal pressure-induced bending moment. For this condition, the maximum failed length is 131.9 inches with the remainder of the joint at room temperature. Both Figures B.57 and B.58 show the failed joint length as a function of the temperature of the remainder of the joint.
(c) Effect of Local Membrane Holes
As a part of the failure investigation, various membrane failures were examined. A round hole was examined with a case internal pressure of 600 psi. For this hole and pressure the stress at the hole edge is 240 ksi which is 23 percent above the D6AC design strength and overall failure of the shell wall would result. At 600 psi internal pressure, elliptical holes with the dimensions shown in Figure B.59 would survive and at 975 psi the elliptical hole shown in Figure B.60 would survive. If the elliptical hole grows in the circumferential direction to a ratio of the major to minor axis of an ellipse (B/A) of 8.0 as shown in Figure B.61, the maximum allowable internal pressure is 227 psi. The analyses support the conclusion that the anomaly was not associated with a burn through in the SRM case membrane.
4.7. Fracture Mechanics
The fracture mechanics analysis that has been completed to date assesses the stresses and deflections during mating of the aft center segment to aft segment. The O-ring groove area has been analyzed to determine the criticality of a flaw in the highest stressed area.
If the clevis was damaged by the mating forces, the weakest location, which would be damaged first, is at the secondary O-ring groove. A fracture mechanics analysis was performed using a conservative 1,000 pound radial load shown in Figure B.62 combined with an 8,000 pound axial load shown in Figure B.63 and the 936 psi internal pressure shown in Figure B.64. The stresses at the O-ring groove for each of these loadings is also shown in Figures B.62 through B.64. Figure B.65 shows the O-ring groove model along with the stress variation for the internal pressure loading case. The load history used for the fracture mechanics analysis is shown in Table B.4. Table B.5 shows the initial flaw size and the final size after 100 launch cycles for two flaw shapes. The flaw size analyzed was larger than that typically found by Magnetic Particle Inspection. The flaw growth is less than 2 percent in 100 launch cycles. The conclusion reached from the fracture mechanics analysis is that mating loads caused no damage to the joint, and that the failure was not the result of growth of an undetected flaw.
4.8. SRM Case Pressure Rounding Discussion
Twelve-foot-diameter shells one-half-inch thick containing propellant weighing over 300,000 pounds are assembled using a tang/clevis joint arrangement. The axial loads are transferred through 177 pins in double shear. The joint has some assembly clearance which is reduced after assembly by installing shims. Some radial clearance remains after shimming. Because of the weight of the motor segment and the high ratio of diameter to shell thickness (D/t) ratio, the assembled joint clearances can vary form one side of the clevis to the other. This clearance was 0.020 inch for the STS 51 -L right SRM aft flight joint. If both parts (tang/clevis) were round and concentric, then a radial clearance of 0.004 inch would be predicted. The pin/hole stop is sufficient to allow this configuration to vary to both sides of the clevis, even though they are on a 2-degree angle.
The aft segment is supported on the four posts on the MLP. The posts are in a rectangular configuration. This nonaxisymmetric loading influence appears to show up on the aft segment clevis field joint as an out-of-round condition. The aft center segment is also out-of-round due to the two-point lifting which causes a water bucket effect.
For the STS 51-L right SRM aft field joint a condition of zero gap (metal-to-metal) minimum to a maximum of 0.020 inch is assumed to exist between the sealing diameters. This represents the maximum and minimum sealing gap at stacking conditions. Some gap equalization will occur due to O-ring compressive forces. This analysis is forthcoming.
The aft center segment tang is lowered into the clevis of the aft segment. The joint is pinned as pin holes line up. Finally, the total load is transferred from the crane to the pins. It is estimated that the line loads are sufficient to create friction loads that will hold the tang/clevis in the as assembled position. The radial loads are estimated to be relatively low so that no repositioning of tang/clevis occurs during the remainder of the stacking.
The SSME firing prior to ignition of the SRM relieves the compression pin load so that some opportunity does exist that will allow adjustment of the sealing gap. Considerable gap adjustment will again occur as pressure builds up in the SRM case after ignition. The out-of- round joint and case membrane tends toward a round condition as this represents a minimum energy condition. Some simple estimates indicate that this may happen at about 200 psi. There are, however, no tendencies to cause the cases to be concentric. STS 51-L right SRM aft field joint would have an annulus of 0.004 inch assuming a round and concentric condition. Therefore, 0.008 inch would be estimated if the case is against one side (metal-to- metal) and the full, nonconcentric gap on the other side (180). The time for this readjustment to occur is very fast; friction is overcome. A snap-action, maybe on the order of 0.008 seconds, occurs. This means that the initial [L114] gap of zero will suddenly become 0. 008 inch. The seal gap of 0. 020 inch could suddenly become 0.008 inch for a change of 0.012 inch.
In the meantime, the SRM pressurization will continue to open the seal gap according to the linear function shown in Figure B.66. These data were determined from the Test MTI 103 which showed a gap opening of 0.034 inch at 1,004 psi. This measurement was taken at the midpoint of the land between the O-ring load. Analysis to transfer this measurement to the primary sealing gap and the secondary sealing gives 0.035 inch and 0.022 inch, respectively. The pressure time function of the joint is shown in Figure B.67.
The effective O-ring squeeze is established as follows:
.
Norm Dia
Tol
Stretch
Set (14 1/2 %)
Temp (length)
(dia)
Eff Dia
Table B.1 has been developed based on an initial gap of 0. 020 inch. It is assumed that when the chamber pressure reaches 200 psi that the compressive pin loads are relieved so that they can assume a new position. The loads that are involved in joint mating are different than those at launch. Therefore, the tang/clevis relative changes can occur during changeover from compression prior to SRM ignition to tension which is developed by pressure blowoff loads. This changeover is estimated to occur at about 200 psi. With these assumptions, Table B.6 was developed for total gap opening and O-ring squeeze at various time steps. Table B.7 was developed with similar assumptions, except that an initial seal gap of zero or (metal - to- metal) was assumed.
Figure B.68 shows the gap results as a function of time for the primary and secondary O-rings. Initial assumption of zero gap and 0.020 inch are shown. The O-ring resiliency response is aggravated by the sudden increase of gap of 0.008 inch due to rounding. At 25°F, the O-ring could not follow the case movement due to pressure, let alone the sudden jump. The case rounding starting at the max gap of 0.020 inch helps to reduce gap opening motion when the 0.020-inch initial gap goes to 0.008 inch. This could cause the seal to open and then recontact the sealing surface so that if pressure were to arrive at the O-ring it could seal.
4.9 Total Gap Opening Versus Flight Time
This section of the report combines the various analysis results to produce an overall SRM field gap opening versus time (flight events designated) for 51-L. (Figures B.69 through B.75 for all field joints right and left SRMs.) Figures B.69 and B.73 are the overall gaps for right and left SRMs while the other figures give a more detailed view of the transient, rounding, and roll maneuver areas. The approach taken has been to add the delta gap to the geometric gap and gap change due to rounding. Indicated on the graph for the lift-off peak and max Q peak are the gap opening range for the design limit load envelope. The rounding effect assumes that the 20-mil worst gap opening is combined with a metal-to-metal condition. When the case rounds the two become essentially equal. Other cases are possible; however, the data presented should be representative.
The principal gap occurs during the first 600 milliseconds of SRM burn going to a peak between 36 and 45 mils. Only the aft field joint gap responds at any significant level to the dynamics. This response is mainly due to strut loads. The bending moment effect is present but not as pronounced. The roll maneuver adds approximately 2 mils. At approximately 20 seconds the gap starts reducing due to internal SRB pressure reduction. Since pressure dominates the SRM loads and gap opening the dynamics (mainly rigid body motion) adds only about 3 mils opening to the pressure and line load-induced opening. These dynamics are due to the wind shear and gust-induced vehicle response. These results match the joint rotation test data differing only by the external loads influence and tolerance stack ups.
Combining the results of the gap opening data with the seal response data indicates strongly that the critical gap opening and seal blowby time occurs within the first 600 milliseconds of SRM burn with some delta motion occurring through the first two seconds. If the O-ring seals normally at lift-off there is no additional situation occurring which should cause the seal to leak since the internal pressure reduces, therefore the gap reduces during max Q. If the seal is degraded or if the joint gets heated, then the gust-induced response could possibly start a leak. In particular, heating the joint would lead to amplification of the gap opening due to dynamics. If the seal was cold (low resiliency) and a metal-to-metal condition existed (seal pressed against groove), it is unlikely that the seal would have pressure actuated and sealed early. The forward and midfield joints have 6 to 10 mils more opening of the primary seal gap at lift-off and max Q. All things being equal these joints should leak first.
In conclusion, the gap openings presented for the primary and secondary O-rings are representative of the situation that occurred on 51-L. The accuracy of the results should be well within ± 10 percent and should be adequate.
5. Findings and Conclusions
a. SRM field joint seal gap opening is primarily due to internal pressure. This pressure bulges the case (radial) and lengthens the case longitudinally (line load). Thrust also adds to line load. Approximately 85 percent of the gap opening is due to the radial bulging for the case where no external forces are applied.
b. External loads and dynamics have a small effect on seal gap opening.
c. The forward field joint has the highest load and seal gap opening with the aft field joint the lowest.
d. The majority of the field joint seal gap opening occurs the first 600 milliseconds after SRM ignition due to the fast buildup of internal ballistics pressure.
e. The largest field joint seal gap opening occurs during the lift-off transients. The roll maneuver has a comparable value.
f. Stackup tolerances and case loading can add up to 8 mils to the dynamic seal gap opening that occurs from pressure and loads.
g. There is a vehicle structural dynamic oscillation during the lift-off transient of approximately 3 Hz. This oscillation adds ± 2 mils to the seal gap opening.
h. Modeling the solid propellant elasticity and damping effects had a small influence on seal gap opening (less than 2 mils maximum).
i. SRM lateral bending and shell modes had negligible effect on seal gap opening predictions.
j. Steel-to-steel compression of the O-ring results in contact and sealing on the sides of the O-ring groove. This side groove sealing can couple with pressure forces to hold the seal in a fully compressed state making pressure sealing improbable.
k. If the O-ring is not sealed on the groove sides, the Bernoulli induced flow forces can rigid body translate the O-ring to the sealed position faster than the gap opening transient.
1. Although linear elastic theory predicts a very fast snapback of the seal from the compressed state, the extreme time temperature dependency of the Viton material makes response from a highly compressed state too slow to follow the gap opening transient. In this state the seal must depend on a free-floating translation for quick sealing of large gaps.
m. The mating and fracture mechanics analysis results show the mating loads and resulting stresses during assembly of the aft center segment to the aft segment were low, causing no damage to the joint.
n. Motor case structure integrity is maintained for a family of membrane burn through cases that result in elliptical hole [L115] formation. if the elliptical hole grows in the circumferential direction to a B/A of 8.0, where B/A is the ratio of major to minor axis of ellipse the maximum allowable internal pressure is 227 psi.
o. The max Q region of flight had active response due to wind gust. The Q was slightly outside the flight experience envelope as was the SRB gimbal duty cycle.
p. If the seal is degraded or if the joint gets heated, then the gust-induced response could possibly start a leak. In particular, heating the joint would lead to amplification of the gap opening due to dynamics.
B. SRM Temperatures
The STS 51-L vehicle experienced the coldest ambient temperatures at launch of any of the Space Shuttle launches to date. Since many of the SRM failure scenarios relate to cold temperatures at least in part, and because some abnormally low infrared temperature measurements were made on the right-hand SRM aft segment prior to launch, there has been a significant effort to assess the pre-launch thermal environments.
The SRM case temperatures are normally predicted to be within a few degrees ( ± 5) of the local ambient temperature except when ambient conditions are changing rapidly or when the motor case is exposed to sunlight. The infrared measurements made on the right SRM were 15° to 17°F below the ambient temperature and were clearly thought to be indicative of an anomalous condition, such as a cryogenic leak, unless the measurements could be found to be in error.
Several studies and tests were conducted to assess the infrared instrument used to take the measurements. In addition an analysis was performed to assess the possibility that normal launch pad or vehicle systems could locally depress the ambient temperature near the SRM enough to be consistent with the infrared measurements. Pre-launch film coverage and other on-pad observations were also reviewed to assess any possibility of cryogenic leaks.
It was concluded from this effort that the infrared measurements were in fact in error, that the local environment on the pad in the vicinity of the SRM was somewhat colder than the ground level ambient temperature, but that there was no indication from these data to suggest a cryogenic leak or anomalous venting on the pad. It was further concluded that the best estimate for the right SRM aft field joint temperature at the coldest location at the time of launch was 28°F with an uncertainty of ± 5°F and that the minimum temperature on the other five joints was approximately the same within 3°F.
1. Ambient Environment
The ground level temperature was falling throughout the preceding night with a value of 30'F at midnight (00:00 E.S.T.) and a minimum of 24°F just before dawn at 07:00 E.S.T. After dawn the ground level ambient temperature increased to 36°F at the time of launch (11:38 E.S.T.). The winds were generally out of the northwest (azimuth of 260° to 310°) at 5 to 9 knots throughout the launch operations. The relative humidity was from 45 to 65 percent with a corresponding dew-point temperature of 16°F to 18°F during the preceding night, and had fallen to 30 percent with a dew-point temperature of 11°F at the time of launch. These data are summarized in Table B.8.
At 00:00 E.S.T (prior to Load)
At 07:00 E.S.T (minimum temp)
At 11:38 E.S.T (time of launch)
.
Ambient Temp (Ground Lvl)
Relative Humidity
Dew Point Temperature
Wind Direction
.
.
Wind Speed
.
.
The ambient temperature measurement for the Launch Pad 39B is a ground level measurement located at camera site 3, approximately 1,000 feet from the pad. The ambient temperature on the launch pad is not measured, however it can be inferred from surrounding meteorological tower data. At the time of launch the surrounding towers showed a significant temperature stratification versus altitude with the air temperature at the SRM aft joint elevation (120-foot level) about 5°F colder than the ground level ambient. This is shown in Figure B.76, and summarized in Table B.9.
.
Ambient Temp At Pad Level
Ambient Temp At Ground Level
Pad - Ground Difference
2. Observations and Measurements
Data pertaining to conditions on the launch pad consist of reported observations from the Ice/Frost Team inspections, infrared measurements of the vehicle and launch pad, and photographic and video coverage of pre-launch operations. These data were reviewed in an attempt to obtain a consistent evaluation of the pad environmental conditions, and to assess the possibility of any anomalous cryogenic systems which could affect the local temperatures.
The Ice/Frost Team observations are well documented by KSC in reports and photographic files which are available upon request. The following is a summary and an assessment of the SRM environment reports.
a. Water Trough Observations
During the night ice had formed in the overpressure water troughs which are located at the base of each of the SRBs. The troughs had been filled with an antifreeze mixture which should have protected against freezing down to temperatures of 16°F. Since the measured ambient temperature was near 24°F, there was some concern that the freezing in the troughs was an indication of additional cooling of the local ambient on the pad. The reports concerning the ice in the troughs read as follows:
During the first assessment made by the Ice/Frost Team at 01:30 E.S.T., "the secondary overpressure water troughs were found to have sheet ice estimated to be 1/2-inch thick. Eight left-hand troughs and 12 right-hand troughs had ice. The ice density was approximately 25 pounds per cubic foot No ice was observed in the primary water troughs until the T-3 hour subsequent inspection was made."
During the subsequent T-3 hour assessment which began at 07:00 E.S.T., the Ice/Frost Team observed "Ice in the troughs had thickened and was found to be solid. All secondary troughs except the northern most one now had ice. The two inboard primary hole troughs were also forming ice. An attempt to use a waterjet from a Firex hose to break up the ice was unsuccessful. . . . Approximately 95 percent of the ice was removed. . . ."
A final assessment was made between 10:30 and 10:55 E.S.T. and the observations from this inspection were "the water troughs were checked and found to be forming ice, which was again removed. . . . During the third inspection. only the two south primary troughs remained ice free."
The first inspection was conducted after a stop flow of the ET cryogenic loading operations. At the time of the inspection there was little or no propellant in the ET, although the pad and Orbiter systems had been cooled down, as well as the ET aft LH2 dome. At this time the Ice/Frost Team estimated that there was 1/2 inch of ice in some of the troughs. Calculations indicate the.....
Figure B.1. List of 51-L Assembly/Flight Events Analyzed.
Figure B.2. 51-L SRM Field Joint Analysis Approach.
Figure B.3. Overall Analysis/Test Flow Plan..
Figure B.4. Field Joint Static and Dynamic Models.
Figure B.5. Finite Element Models for the Clevis, Tang, Aft ET Attach Ring, and the Aft 1 Diameter Length Segments.
Figure B.6. Propellant/Case Model Definition.
Figure B.7. SRB Internal Pressure Build-Up Function (Aft Field Joint).
Figure B.8. Propellant Damping Properties (Ref- TWR-11779).
Figure B.9. Joint 91 Acceleration Response, With and Without Propellant.
Figure B.10. Joint 91 Velocity Response, With and Without Propellant.
Figure B.11. Joint 91 Displacement Response, With and Without Propellant.
Figure B.12. ANSYS Model of Tang/Clevis Segment, Linear Run.
Figure B.13. ANSYS Model of Pin, Linear Run.
Figure B.14. ANSYS Model of Tang/Clevis Segment, Nonlinear Run.
Figure B.15. ANSYS Model of Pin, Nonlinear.
Figure B.16. O-Ring Material Property Characteristics.
Figure B.17. O-Ring Displacement Shapes for Various States of Compression.
Figure B.18. Response of Steel to Steel Compressed O-Ring at Different Flow Pressures, Linear Elastic Response.
Figure B.19. Compressed O-Ring .046 Inch Model.
Figure B.20. Response of O-Ring Compressed to .046 IN. (Linear Elastic Response, No Pressure).
Figure B.21. Response of O-Ring Compressed to .046 IN. (Linear Elastic Response With Pressure).
Figure B.22. Response of Free-Floating O-Ring Compressed to .046 IN. (Linear Elastic With Pressure).
Figure B.23. O-Ring Compression Required to Seal Groove Sides.
Figure B.24. O-Ring Compression Required to Seal Groove Sides.
Figure B.25. Steel to Steel Compression.
Figure B.26. Strain Gage Data versus Pressure versus Gage Location, Referee Test (MTI 103).
Figure B.27. Gap Opening Between the Primary and Secondary O-Ring Referee Test Data (MTI 103).
Figure B.28. Gap Opening Versus Pressure versus LVDT Positions, Referee Test Data (MTI 103).
Figure B.29. Primary Gap Opening Versus Time, Primary Seal.
Figure B.30. Secondary Gap Opening Versus Time, Primary Seals.
Figure B.31. Primary Gap Opening Versus Time, Primary Leaks.
Figure B.32. Secondary Gap Opening Versus Time, Primary Leaks.
Figure B.33. Seal Gap Opening Versus Pressure, Referee Test (MTI 103).
Figure B.34. Primary Gap Opening at Alignment Slot Versus Time, Primary Seals.
Figure B.35. Reconstruction of Flight Liftoff Loads.
Figure B.36. Reconstruction of In-Flight Loads.
Figure B.37. Time Response of Reconstructed Loads of the Aft Field Joint.
Figure B.38. SRB Attach Member Load Locations.
Figure B.39. Shuttle Coordinate System View Looking Forward.
Figure B.40. Delta Gap Opening Transient, Aft Field Joint.
Figure B.41. Max-Min Summary of Delta Gap Opening, Aft Field Joint.
Figure B.42. Delta Gap Opening Transient, Forward Field Joint.
Figure B.43. Delta Gap Opening at t= 8.15 Sec., Forward Field Joint.
Figure B.44. Delta Gap Opening During Roll Maneuver, Forward Field Joint.
Figure B.45. Delta Gap Opening at t = 36.67 Sec., Forward Field Joint.
Figure B.46. Delta Gap Opening at t = 63.67 Sec., Forward Field Joint.
Figure B.47. Delta Gap Opening During SSME Thrust Buildup, Aft Field Joint.
Figure B.48. Delta Gap Opening Transient, Aft Field Joint.
Figure B.49. Delta Gap Opening at t = 8.15 Sec., Aft Field Joint.
Figure B.50. Delta Gap Opening During Roll Maneuver, Aft Field Joint.
Figure B.51. Delta Gap Opening at t = 36.67 Sec., Aft Field Joint.
Figure B.52. Delta Gap Opening at t = 63.67 Sec., Aft Field Joint.
Figure B.53. Delta Gap Opening at P = 936 PSI, No Strut Loads.
Figure B.54. Delta Gap Opening, Alignment Slot Region Detail.
Figure B.55. Referee Test Data (MTI 103) Displacements at Alignment Pin Slots.
Figure B.56. Axial Load only Effect on Circumferential Opening.
Figure B.57. Axial Load and Vehicle Bending Moment Effect on Circumferential Opening.
Figure B.58. Axial Load and Internal Pressure Induced Bending Moment Effect on Circumferential Opening.
Figure B.59. Elliptical Hole Formation, Motor Case Pressure - 600 PSI.
Figure B.60. Elliptical Hole Formation, Motor Case Pressure - 975 PSI.
Figure B.61. Elliptical Hole Formation, B/A = 8.00
Figure B.62. Clevis Joint Stress Distribution, 1000 Pound Radial Load.
Figure B.63. Clevis Joint Stress Distribution, 8000 Pound Axial Load.
Figure B.64. Clevis Joint Stress Distribution, 936 PSI Flight Load.
Figure B.65. O-Ring Groove Model, Stress Variation for Internal Pressure.
Figure B.66. Seal Gap Opening Versus Pressure, Referee Test (MTI 103).
Figure B.67. Pressure vs. Time, 51-L Right SRM Aft Field Joint.
Figure B.68. STS 51-L Right SRM Aft Field Joint, Rounding at 200 PSI.
Figure B.69. Total Gap Opening For All Flight Events, STS 51-L Right SRM Field Joints.
Figure B.70. Initial Gap and Rounding Effect Right SRM.
Figure B.71. SSME Build-up and Lift Off Transient.
Figure B.72. Roll Maneuver Delta Gap Opening.
Figure B.73. Total Gap Opening Left SRM Field Joints.
Figure B.74. Initial Gap and Rounding Effect Left SRM.
Figure B.75. SSME Buildup and Lift Off Transient.
Table B.3. SRM Field Joint Loads Comparison.
Table B.4. Load History Used in Fracture Mechanics Analysis.
Table B.5. Fracture Mechanic Results.
Table B.6. O-Ring Squeeze 51-L Right-Hand Motor, Aft Field Joint. File O-Ring N1A.
Table B.7. O-Ring Squeeze 51-L Right-Hand Motor, Aft Field Joint. File O-Ring 1B [?18].
[L166]....ice could easily be forming in these troughs due to cooling from the night sky alone. The troughs are recessed somewhat below pad level and are thus protected from wind-driven convection. Because of the horizontal orientation, the water/anti-freeze mixture will experience maximum radiative cooling to the night sky. The underside of the troughs is also protected from convection because of the position high in the Mobile Launch Platform (MLP) exhaust hole.
Calculations of the heat balance for the water troughs was performed for ambient temperatures beginning on the previous evening. These calculations predict the equilibrium temperature that the mixture would reach if allowed to completely freeze and come to steady state, as well as the freezing rate in inches per hour if held at the 16°F freezing temperature reported for the mixture. A summary of these calculations is presented in Table B.10 and shows that the mixture could have begun to freeze as early as 07:00 p.m. the previous evening. This probably did not occur, however, since the troughs were being filled with the antifreeze mixture near that time. Note that the "equilibrium mixture temperature" is a calculated lower limit and not the expected mixture temperature.
Date - Time
.
1/27 19:00 E.S.T.
1/27 21:00 E.S.T.
1/28 00:00 E.S.T.
1/28 07:00 E.S.T.
The freezing rates shown would predict about 0.3 inch of 25 lb ft2 density ice at the time of the first inspection which is considered to be within range of the observations. Additional cooling due to the evaporation, or a higher than expected freezing temperature of the mixture could explain the difference between the calculation and the observed 0.5-inch thickness. (Tests at KSC indicate that the water-antifreeze solution may have stratified and that the top layer could have been freezing at temperatures between 18°F and 22°F.)
b. Left SRB Icing Observations
An additional observation made by the Ice/Frost Team which relates to the SRM environment was that the left SRB aft skirt and aft motor segment had formed patches of sheet ice due to water spray from the Fixed Service Structure (FSS). The occurrence of ice on the left SRB is used to explain the infrared measurements made at this location. The spray of water from the FSS and subsequent freezing on the SRM surface would provide a 32°F surface temperature as viewed by the infrared instrument, even though the pad ambient temperature was near 22°F.
c. Infrared Measurements
The Ice/Frost Team routinely use an infrared pyrometer to aid in the assessment of icing conditions on the ET. This instrument provides for remote sensing of the tank surface temperature based on the infrared energy omitted from the surface. Measurements are also made of selected SRB and Orbiter surfaces as well as the launch pad structure. During the STS 51-L inspection an Omega Scope 2000A infrared gun was used to make the measurements. The gun had been in use for the previous two inspections including the STS 51-L abort and STS 61-C.
The initial readings obtained during the STS 51-L inspection revealed several measurements which were abnormally low, including a 7°F reading of the right SRB skirt, and a 9°F reading on the right aft motor segment. At the time of these readings, the ambient temperature was at 22°F at pad level and had been gradually falling several hours as shown in Figure B.76. Exposure to the cold night sky and the cold ET surface could cause the SRM case to be somewhat below ambient in this transient, but not as low as the 9°F obtained from the infrared measurement.
d. Infrared Pyrometer Testing
Testing of the infrared pyrometer used by KSC, and testing of the same type (model) but of a different unit, were performed by both MSFC and JSC. All three investigations revealed a substantial transient error or bias in the instrument when it is exposed to a change in temperature similar to the transition from room temperature to the 22°F ambient as experienced for STS 51-L. The transient error for the instrument tested by KSC is shown in Figure B.77. A similar but slightly higher error (-18°F) was observed on the instrument tested by MSFC, and the tests at JSC show an even greater error on their instrument especially when exposed to wind at the lower temperature. All three tests showed good measurement accuracy once the instrument had stabilized at the exposed environment.
In addition to the transient bias of the instrument discussed above, an error due to surface emissivity was also identified, both analytically and by test. The error occurs because the instrument is set for a surface emissivity of 1.0 whereas the actual SRM surface emissivity is from 0.87 to 0.95 depending on surface contamination. The right SRM is especially susceptible to the error since the reflected background radiation is primarily from the night sky, which emits little radiation in the wavelength at which the instrument operates. The estimated corrections to the right SRM for emissivity is + 3°F with an uncertainty of + 2.0°F. Surfaces which are wetted, or which have ice or frost formations would have little error expected because both ice and water have an emissivity near 1.0 which eliminates the error. The ET, the water troughs, and the areas of the left SRM which had ice formations are in this category.
e. Corrected Infrared Measurements
The infrared measurements were corrected based on the transient error reproduced by the NASA-KSC tests on the actual instrument and on the timeline of observations made by the Ice/Frost Team. The additional errors due to emissivity and general instrument inaccuracy were also added and the uncertainties in each were root sum squared (RSS) to obtain an overall uncertainty. The error budget for the right and left SRM measurement locations is presented in Table B.11 below.
.
.
Transient Error
Emissivity Error
Instrument Error
Total
The left and right SRM measurements were corrected based on the total error shown in Table B.11 and compared to our predicted temperature for the same location (Table B.12). Also included in Table B.12 are three other selected measurements made on the MLP level and corrected in the same manner. The range accounts for the uncertainty in both the corrected measurement and the analysis.
.
.
Right SRM
Left SRM
ET Aft Dome
ET at Right SRM
Water Troughs
[L167] The right SRM shows as good agreement as is normally expected between measured and predicted data. The left SRM prediction does not account for the water and ice buildup observed which would explain the corrected measurements. The two ET measurements agree well with predictions, and the water trough measurements agree with the reported antifreeze freezing temperature of 16°F. The low value of the water trough prediction is for a transient cool-down of a frozen mixture, whereas the high value is for a partially frozen mixture and was thus set at the reported mixture freezing temperature.
f. Photographic and Video Analysis
Both photographic and video coverage of pre-launch operations were reviewed to assess the possible transport of the SSME seal drain vapors toward the right SRM as well as detect any indication of a cryogenic leak or venting on the pad. The vapors from the engine drain lines were observed to be blowing toward the right SRM at times, and were generally consistent with a launch pad flow analysis discussed later. No other anomalous vapors were noted.
3. Analysis of Local Ambient Temperature Depression
The thermal analysis which is performed to predict the pre-launch SRM temperatures normally does not include any depression of local ambient temperature due to the effect of pad or vehicle cryogenic or subcooled venting, or due to convective cooling from the ET. (Radiation to the cold ET surface, however, is included.) In view of the low infrared measurements made on the right SRM and the ET (uncorrected) a study was initiated to quantify the possible depression of the ambient temperature on the launch pad.
a. Pad and Vehicle System Purges and Vents
Cryogenic or subcooled gases are routinely vented in the vicinity of the Shuttle vehicle including the oxygen vent at the top of the fixed surface structure, the SSME seal drains which vent at the base of the Orbiter, and several inerting purges which vent at various locations. The vehicle and ground purges were reviewed and none were found to be of a magnitude or in a position to affect the environment near the right SRM. The oxygen venting at the top of the fixed surface structure similarly is not in a position to affect the right SRM although it could influence the ET upper surface temperatures. The SSME drain seals, however, do provide a large flow rate (approximately 1.1 lbm/sec of 250°F gas from each of the three engines). They are in a location where southwest winds may provide a transport of gas to the vicinity of the right SRM. This was selected for further study.
b. ET Convective Cooling
Convective cooling from the ET surface as it affects the SRM temperatures is normally assumed to be minor since the ET insulation system provides very good insulating capability and thus limits the heat loss to the surrounding air. In addition, any natural or forced thermal boundary layer which may form on the tank surface would tend to mix and rapidly dissipate in the continually changing wind conditions before interacting with the SRM surface. Conditions which would be favorable to maintaining a cold local environment around the SRM would be little or no wind, which would also limit the effect on the SRM. In view of the low measurements on the right SRM, however, the potential for convective cooling from the ET was selected for further study in an effort to at least establish bounds on the possible effect.
Two efforts were initiated by MSFC to further study the potential for a local ambient temperature depression. The first was a three-dimensional flow and heat transfer analysis of the Shuttle and Launch complex to assess the interactions of the wind with the pad structure and vehicle, convective cooling from the ET, and the possible transport of the SSME drain seal gases to the SRM. This was performed by CHAM of North America Incorporated. The second effort was part of a general independent SRM thermal analysis performed by SRS Technologies which was intended as an audit of the MSFC analysis. The SRS effort assessed the potential effect of ET convective cooling on the SRM in the absence of any wind.
(1) Launch Pad Flow and Thermal Analysis
The three-dimensional thermal and flow numerical model included the launch pad (mound), the MLP, the support structures, and the Shuttle vehicle. These features were modeled in an attempt to determine the wind flow around the vehicle, and also to assess any wind induced vertical flow either through the SSME exhaust duct or near the Shuttle vehicle. The wind conditions in the launch area during the STS 51-L launch operations were used, with the wind directions of 260°and 300° considered. The analysis has shown that the maximum cooling near the right SRM is about 4.5°F for the 260° wind direction, and about 3°F for the 300° wind direction.
(2) ET Natural Convection Analysis
The purpose of this analysis was to determine the temperature and velocity in a boundary layer that could possibly have developed on the ET and intersected the SRM. The boundary layer analysis performed considered the flow to be initially laminar and then transitioning to turbulent flow with a simultaneous heat balance solution at the ET surface. The results of the analysis predicted that at the SRM surface, 12 inches from the ET, the velocity would be 1.73 ft/sec and the temperature 5.5°F below ambient. This is at the station of maximum boundary layer thickness which occurs at the base of the ET near the SRB attach point.
The boundary layer was allowed to grow over the entire height of the hydrogen tank undisturbed by wind, thus producing the maximum thickness, velocity, and cooling at the SRM surface, which should be considered a worst-case bounding of the problem. The occurrence of this type of boundary layer would also result in colder than normal ET surface temperatures, calculated to be below -25°F at the point of maximum boundary layer thickness. There was no indication that this had occurred either from the infrared measurements or from observations, such as excessive frost buildup.
c. Summary of Local Temperature Depression
The three-dimensional thermal and flow analysis and the ET natural convection analysis indicate that an ambient temperature depression on the order of 3°F to 5°F could be possible in the vicinity of the right SRM for the conditions which existed during STS 51-L. These results were subsequently used to adjust the depressions in the aft field joint predicted temperatures discussed in subsection C-5.e.
4. SRM Thermal Analysis
a. SRM Thermal Model Description
Thermal analysis of the SRM case is performed using a two-dimensional model of a SRM segment which can be configured to represent either the left or right aft or mid-segments. The model includes the motor case, insulation, and propellant and is constructed in eight circumferential nodes (around the case) and nine radial nodes from the steel case inward through the propellant. The external environment simulation includes convection based on an average wind speed, solar radiation, and radiation to the ambient, the ET, the Orbiter, and the sky. Sky temperature is calculated from ambient temperature using a simplified algorithm, and solar radiation is simulated by an algorithm as a function of date, time, and position on the SRM case. The analysis is transient and typically encompasses a 150-hour period (6.25 days) prior to launch. The hourly observed ambient temperatures for the mission of interest are used as input. These temperatures are corrected for any difference due to elevation as discussed previously.
b. SRM Thermal Analysis for STS 51-L
Three cases were run for each SRM. These included a nominal aft segment, a nominal mid-segment, and a cold case aft segment. The cold case used extreme values for various assumed parameters in an attempt to establish a lower bound on the predicted joint temperature. These included minimum wind, an extreme cold sky temperature (approximately - 30°F), and maximum radiative cooling from the ET.
A summary of the cases run for STS 51-L is presented in Table B.13. The table shows the range of temperatures around the motor case circumference for each of the three cases for both the left [L168] and right SRM. As seen from these data the minimum predicted temperature on the right aft SRM at the time or the inspection was 20°F for the cold case. The nominal case which is based on more realistic assumptions predicts a minimum of 230F. Similarly, for the time of launch the minimum predicted temperature on the right aft SRM is 24°F for the cold case and 28°F for nominal.
Also evident from these data is that the minimum temperature at each of the joints for both right and left SRMs is approximately the same within the uncertainty band. Only those areas in the sun are significantly warmer. We would expect the right booster to be somewhat cooler than the left due to possible ET convective cooling, but only by 2°F to 3°F.
.
.
A: 07:00) E.S.T. Inspection)
Right SRM
.
Left SRM
At 11:30 E.S.T. (Launch)
Right SRM
.
Left SRM
Note: Range shown is around the SRM circumference.
Figure B.78 presents the ambient temperature versus time for the 2.5 days prior to STS 51-L launch. Figure B.79 shows the average motor case temperature and mean propellant bulk temperature for the aft right segment over the same timeframe. Figure B.80 shows the eight circumferential segment temperatures for the aft right SRM segment during the 16 hours prior to launch and Figure B.81 shows the same data for the aft left SRM segment. A complete report on the thermal analysis performed for STS 51-L including all the output data is contained in the SRM Working Group Files.
c. Independent SRM Thermal Analysis
MSFC contracted with SRS Technologies to build an independent thermal model of the SRM and Shuttle configuration providing an independent predicting of SRM joint temperatures for the STS 51-L mission, and an audit and verification of the MSFC SRM thermal model. The SRS effort also included an analysis of ET convective cooling.
A summary of the SRS cold and nominal analysis of the aft right segment as compared to the MSFC analysis indicates good agreement between the two models (Table B.14). Note that the MSFC model predicts lower cold case temperatures than the SRS model. This has been attributed to the finer nodal breakdown and more the detailed radiation coupling of the SRS analysis. The complete SRS final report is contained in the SRM Working Group Files.
.
.
07:00 E.S.T.
Cold Case
(inspection)
Nominal
11:30 E.S.T.
Cold Case
(time of launch)
Nominal
d. ET to SRB Attach Strut Thermal Analysis
A thermal analysis of the possible heat loss due to conduction through the aft attachment struts between the ET and the SRB was conducted using a detailed strut thermal model and a simplified SRM three-dimensional case section model. The attach strut model established the heat load on the SRM which was then applied to the simplified SRM case action model. Results show that the heat loss through the strut (approximately 12.6 Btu/hr per strut) will cause a local temperature depression at the SRM attach point of no more than -4.5°F, which dissipates to less than -0.5°F within 15.5 inches from the strut attach. Note that the aft field joint is approximately 19.5 inches from the attach point and is thus unaffected.
e. Summary of STS 51-L SRM Aft Joint Temperature Predications
A summary of the SRM aft joint predictions for the STS 51-L mission is presented in Table B.15 together with the ambient temperatures. The uncertainty was increased slightly from the predictions presented above to account for additional dispersions clue to possible convective cooling from the ET. The SRM thermal model was run with a constant 5°F depression in the ambient local temperature which resulted in SRM temperature 30°F colder at the time of launch and 2°F colder at the minimum temperature point. These dispersions were then added (root sum squared) to the uncertainty already established for the SRM predictions.
.
Launch Temps °F
.
Ground Level Ambient
Pad Level Ambient
Predicted SRM(aft field joint)
5. Comparison to Other Launches
The same type of thermal analysis performed for the STS 51-L mission was also performed for the next nine coldest launches as initially determined by the ambient temperature at the time of launch. The results of these analyses are presented in Table B.16 and Figure B.82. The next coldest SRM temperatures at launch was STS 51-C which had a minimum predicted temperature of 51°F at the time of launch. The STS 41-C mission is almost as cold with a predicted 52°F at the time of launch. A detailed report for each of the nine analyses is documented in the SRM Working Group Files.
C. Assessment of Joint Leakage From Near Ignition to 58 Seconds
1. Scenario Description
This scenario considers that the puff of smoke observed at 0.678 second was an indication of an SRM hot gas leak through the aft field joint which finally manifested itself as the observed hot jet at 50 seconds. The scenario assumes that between lift-off and 58 seconds the initial leak either: (1) continued unobserved at a low level without a joint burn through, or (2) plugged and sealed itself until a few seconds prior to 58 seconds when it again began leaking.
2. Investigation
The investigation included an analytical assessment of the flow, thermal, and structural behavior of a hot gas leak through the joint, supported by tests of small motors with simulated joint/seal defects. Both the analyses and tests assumed an initial leak path through the putty and the primary and secondary O-rings.
The analytical effort attempted to follow the progression of the leak through the initial 58 seconds of flight to either demonstrate or refute a continuous leak scenario if possible. Alternate analyses were conducted of the initial start transient to assess the mechanisms which may lead to a leak-stop-leak scenario.
The testing effort consisted of two series of hot fire tests. The first series was 5-inch motor tests which were conducted to obtain early data on the response of the O-ring joints to various defects and induced flow environments and were used to design the second test configuration: 70-pound charge motor tests which were designed to measure the thermal response of a full-scale cross section joint due to leakage of hot gas.
a. Analytical Approach
The analytical effort consisted of independent flow, thermal and structural analysis of the joint and hot gas leak, as well as limited coupled analysis of the flow and thermal problem. These....
Figure B.76. Ambient Temperature at Ground Level and on the Launch Pad for STS 51-L.
Figure B.77. Infrared Radiometer Error Due to Stabilization Effect.
Figure B.78. Ambient Temperature for 2.5 Days Prior to STS 51-L.
Figure B.79. Average Case and Mean Propellant Bulk Temperatures for STS 51-L.
Figure B.80. Aft Right Segment Temperatures for STS 51-L.
Figure B.81. Aft Left Segment Temperatures for STS 51-L.
Figure B.82. SRM Temperature for the Ten Coldest Launches.
Table B.16. Summary Chart for the 10 Coldest Launches.
[L173] ....analyses required an initial definition of a credible leak path as a starting point. Since several scenarios are possible leading to the initial leak, the various potential types of leak paths were classified according to geometry to assist in the analysis:
Joint leak type 1 places the blowhole, primary O-ring, and secondary O-ring are all in line at the same circumferential location. This would be typical of blow-by and would not include any circumferential flow. The character of this leak type is that only one blowing leak would be expected at 58 seconds.
Joint leak type 2 places one or more of the leak paths out of line such that circumferential flow exists either between the blowhole and the primary O-ring leak, or between the primary and secondary O-ring leaks, or both. This would be typical of O-ring damage or a possible tang scratch which would not be expected to line up with a blowhole. This type can be further subclassified according to the various possible offsets; however, the analysis is similar and suggests that the character at 58 seconds would be two or more blowing leaks.
The geometric configuration of these leak types is shown pictorially in Figure B.83, which also includes the subclassification of type 2.
Most of the analytical work to date has addressed the type I joint leak. These analyses include: (1) an erosion analysis of the NBR, putty, and O-rings, (2) a quasi-two-dimensional and three-dimensional steady-state flow analysis of the axial flow between the clevis and tang, (3) a two-dimensional and three-dimensional thermal analysis of the field joint, and (4) supporting structural analysis of the joint.
b. Erosion Analysis
The erosion rate of the V44 asbestos-filled NBR insulation was derived from the "char-motor" testing (Pendleton, S. B.; Evaluation of Elastomeric Insulation Materials at High Chamber Pressures, Morton Thiokol TWR-3896) which correlates material loss rate to convective heat transfer coefficient. The variation in erosion rate of the joint filler material (Randolph asbestos filled putty)with convective heat transfer coefficient was assumed to be equivalent to the V44 insulation material with a constant bias based on a single data point from subscale hot fire testing. Erosion of the Viton O-rings was based on the analytical model of Salita (Prediction of Pressurization and Erosion of the SRB Primary O-Rings During Motor Ignition, Part I and II; Morton Thiokol TWR-14952 and TWR-15186).
The O-ring erosion model has been used to predict O-ring erosion due to hot gas impingement and blowby (or leakage) of the O-ring. Since it is a transient cavity pressurization model, the erosion model has been used to determine the flow rate into and out of the O-ring cavities during the motor ignition transient due to leakage past the O-rings. The O-ring leakage areas were determined from the change in clearances between the clevis and tang during motor ignition (joint rotation) and measured O-ring resiliency at 25°F. These flow rates were input to the two-dimensional thermal model of the joint and O-rings to determine the time to heat the primary and secondary O-rings to their ablation temperature.
c. Flow Analysis
The steady-state models of flow through the joint include two three-dimensional Computational Fluid Dynamics (CFD) models used to determine the relative mass flux (flow spreading) due to the complex flow path through the joint, and a quasi-two-dimensional compressible flow model with friction. The quasi-2D model treats the flow path as a series of elements with the characteristics of width, height, and length. The putty blowhole and O-rings are allowed to erode with the erosion characteristics described above. Flow spreading is determined by a modified empirical free jet correlation where spreading is a function of element equivalent lengths. The calculated flow rate and relative mass flux is input to the thermal models to determine the joint temperature transient.
The CFD analysis assumes three-dimensional incompressible flow with friction and is based on the PHOENICS code of Cham, Inc. Two models have been developed to assess the flow through the clevis; one models only the gas path entrance into the clevis and O-ring area and has been used to assess flow turning effects upstream of the O-rings; the second model includes the gas path from the O-rings through the clevis and has been used to examine the three-dimensional effects as the gas flows around the pins and through the clevis. The combined spreading function (centerline flowrate divided by total flowrate) from these two models compare favorably with the empirical spreading function of the quasi-two-dimensional model described above.
d. Thermal Analysis
The thermal analysis of flow through the joint used two separate thermal math models of the joint, a detailed two-dimensional model, and a less detailed but more comprehensive three-dimensional model. Both models are based on the lumped parameter finite difference method and provide a transient analysis accounting for thermal mass, ablation of the internal insulation and putty, and melting of the clevis and tang.
The detailed two-dimensional model encompasses an axial slice of the joint between the pins and includes the clevis, tang, internal insulation, and putty. The flow network is one-dimensional with the internal heat transfer coefficients based on the gas temperature, density, velocity, and the current geometry of the flow path as a result of any insulation erosion of steel melting. Spreading of the flow circumferentially is simulated by varying the centerline flow rate through the joint.
The three-dimensional model uses a simplified nodal breakdown in the axial and radial directions to all inclusion of the third (circumferential) dimension. The three-dimensional model encompasses approximately a 90° arc of the SRM joint, and is built axisymmetric around the initial leak path, thus providing effective simulation of an 180° arc. The three-dimensional model includes a two-dimensional flow network providing a true simulation of circumferential flow and flow spreading through the joint.
Several methods were used to establish the leak flow rate through the joint for the two thermal models. These included: (1) the use of constant parametrics to determine the flow rates which would allow the joint to survive the leak flow heating for 60 seconds, (2) flow rates calculated by a simplified technique to account for the changing geometry used to assess the coupling between the flow rate, the enlarging flow gap due to melting, and the subsequent heating of the tang and clevis, and (3) flow rates developed by the flow analysis discussed above.
3. Findings and Conclusions
a. Flow Analysis
Several simplified (single element) flow analyses were conducted to determine the relative influence of the various flow paths. It was determined, due to the rapid erosion rate of the putty blow hole (NBR/putty and the O-rings) at choked flow conditions, that these elements would not control the flow rate through the joint after the start transient. The flow would be controlled by the metal clearance between the tang and clevis and the relative flow spreading from the entrance jet through the putty blow hole and would probably be influenced by the deposition of aluminum oxide on the metal surfaces. This required that the multi-element model described above be developed to obtain a reasonable estimate of flow. It was also determined that a substantial leakage rate could exist without being detectable on measured SRM internal pressure.
By comparison with the relative mass flux of the three-dimensional CFD models, it was determined that the simplified two-dimensional spreading function in the multi-element model was adequate to describe the flow and therefore, allows the flow model to be fully coupled with the two or three-dimensional thermal models.
b. Thermal Analysis
The results of the various parameteric analysis performed with the two-dimensional model are presented in Figure B.84. The [L174] constant flow parametrics indicate that the average flow rate should be no more than 0.006 lbm/sec per inch of circumference to survive for 58 seconds. A study which varied the putty blow hole size showed that an extremely small blow hole of 0.020 inch or smaller would similarly be required and a study which varied the tang to clevis clearance indicated that a very small clearance of less than 0.002 inch would be required to be consistent with a 58-second continuous leak. Both the blow hole size and the tang to clevis clearance quoted above are significantly less than what is expected for these dimensions. In addition to these parametric studies, both the two-dimensional and the three-dimensional models were run with the more exact flow and spreading analysis discussed in the previous section. The results of these analyses showed a time to failure of only 20 seconds for the three-dimensional case and when fully coupled with the flow analysis me time will be even less due to the rapid increase in flow when the clevis and tang begin to melt.
c. Testing
The 5-inch motor and 70-pound charge motor tests are reported in summary form in Appendix C. The 5-inch motor tests showed that the diametrical gap between metal joint surfaces is a significant parameter in establishing gas flow rates and time required to burn through the metal once a gas leak is initiated. The tests also demonstrated that putty could act as an effective seal and that O-rings with a clearance of up to 0.004 inch between the O-ring and metal surface (negative squeeze) have a high probability of sealing if pressurized uniformly over the entire circumference.
Diametrical gap also appears to be a significant parameter in the 70-pound charge motor tests. The test with a 0.0035-inch gap between the inside clevis leg and the tang showed no O-ring erosion or metal erosion over 84 seconds of burn time although a small leakage in the form of white smoke was visible. Two tests with a 0.040-inch gap resulted in a significant amount of black during the first 3 seconds and flame leakage between the Levis and tang. One of these tests experienced a burn through at 11.7 seconds, the other did not burn through in 24 seconds of bum time. However, the tang on the later test was eroded to within 0.07 inch of bum through.
Two additional tests at an intermediate gap of 0.020 inch also demonstrated a significant difference. The first test emitted very little smoke or flame but continued to leak at a small flow rate for 84 seconds with only minor O-ring damage. The second test of the same configuration produced obvious black smoke for 1.5 seconds, then appeared to virtually stop until 40 seconds when the pressure between the O-rings again indicated a small leak. This leak lasted for 10 seconds after which it stopped for the remainder of the 84-second test.
The time required to bum through is also significantly influenced by the angular location of the leak; i.e., leakage near an alignment pin hole burns through in a shorter time.
Black smoke was produced on tests with a relatively small O-ring leak path and large (0.020- to 0.040-inch) diametrical metal clearance for up to 3.5 seconds.
d. Conclusions
The conclusion from the analytical studies is that a leak would not have been expected to continue without a major bum through for 38 seconds. On the contrary, the models would predict a bum through in as little as 10 seconds. However, there are identified mechanisms which are not currently included in these models which could limit the flow through the joint, or the heating to the joint surfaces. These include joint clearance closure due to uneven thermal expansion, the redeposition of melted steel, the condensation and deposition of alumina from the SRM exhaust, and the deposition of insulation and/or putty decomposition products.
It is well known that the deposition of alumina may occur in this type situation, and recent small motor tests have demonstrated this in most cases. The analysis of this phenomenon however requires the use of empirical accommodation factors with an associated probability which render the analysis highly uncertain. This type analysis could well show blockage and thus a leak-stop-leak scenario, a limited continuous leak with constantly changing alumina depositions, or a burnthrough in a very short time. The difference would be the probability associated with the accommodation of the condensed alumina.
The results from the 70-pound charge motor tests of the full scale cross section joint indicate that, once an O-ring leak develops, it is very difficult to predict the thermal response of the field joint. Many factors, including the size of the diametrical gap, the amount of aluminum oxide plating, and extrusion of the putty can significantly influence the amount of hot gas flow and the direction of the gas path and whether the leak will continue or stop. The tests conducted to date have demonstrated conditions which will produce black smoke, those which result in a stop leak, and one mechanism by which an O-ring leak could start at ignition and continue without any observable burnthrough for a substantial period of time.
4. Scenario Conclusions
Both the analytical studies, and the testing described above have shown that in addition to a general erosion of the joint a burn through can occur in two locations depending on the leak flow rate and blockage due to alumina deposition. These are depicted in Figure B.85 and are called bypass burn throughs since the erosion through the metal parts eventually causes the flow to bypass the initial leak path through the O-rings. Both types occur where the flow path through the insulation gap makes a turn which causes the alumina particles in the flow to impinge on the turning surface. This type burn through, coupled with flow blockage from alumina deposition, forms one of three scenarios developed for the 0 to 59 seconds period. The other two scenarios relate more to the unpredictable nature of alumina deposition and blockage. The three scenarios identified to date are as follows:
Scenario 1. The initial leak at lift-off continues throughout the 0 to 58 second period, limited by alumina deposition and/or insulation and putty debris. During this time, by-pass erosion is occurring which finally results in a blowing hole at 50 seconds.
Scenario 2. The initial leak at lift-off continues throughout the 0 to 58 second period, limited by alumina deposition and/or insulation and putty debris. The alumina deposition is such that with the gradually closing tang to clevis gap (with chamber pressure) the blockage is maintained until near 58 seconds. At this point, either dynamics associated with wind shear or other loads, or the slight but noticeable opening of the tang to clevis gap as the chamber pressure begins to rise, causes the blockage to break down, resulting in a rapid bum through.
Scenario 3. The initial leak at lift-off does not continue past approximately 5 to 6 seconds (based on film coverage), being blocked by alumina deposition and/or insulation and putty debris. Initial analytical studies show that rapid alumina deposition could possibly block the clevis to tang gap. The flow stays blocked until near 58 seconds until the brittle alumina deposit breaks up, either due to dynamics associated with wind shear or other loads, or due to the slight but noticeable opening of the tang to clevis gap as the chamber pressure begins to rise, resulting in a rapid burn through.
At this time we cannot classify any of the three above described scenarios as significantly more probable than the other. Our conclusion, however, is that there is sufficient evidence to adequately support any of the three as a feasible explanation of the 0 to 58 seconds time period for STS 51-L.
D. Flow and Thermal Analyses
1. Organizational/Responsibilities
This section 4 the 51-L investigation report deals with the thermal and flow analyses which are not reported in other sections;....
Figure B.83. Joint Leak Classification.
Figure B.84. Summary of 2-Dimensional Thermal Analysis.
Figure B.85. By-Pass Burn Through.
[L177] ....e.g., sections B and C of this appendix and the Space Shuttle Systems Working Group Report.
Various organizations participated in this effort, including:
2. Method of Investigation
The methodology used for flow and thermal analyses was: (1) establish event phenomena necessary for analysis; (2) establish configuration specific geometry data base pertinent to the phenomena; (3) select individual available models capable of phenomena analysis; (4) develop additional mathematical models where needed; (5) establish requirements for necessary physical data; (6) when possible, integrate flow and thermal models: (7) perform flow and thermal analysis calibration runs for validity evaluation; (8) perform flow and thermal analysis of SRB case and joint behavior; (9) where necessary, iterate between thermal, flow, and structural geometry effects; and (10) evaluate results in terms of findings and conclusions based on test data comparison when possible.
Whenever possible, the fundamental physical data were extracted from SRB related simulations.
3. Terms and Acronyms
The terms and acronyms used in this appendix are identified in the main body of the report.
4. Narrative Description
The majority of the thermal and flow analyses were performed in support of prelaunch and launch SRB temperature predictions and the leaking joint flow and thermal assessments. These are covered in sections C and D, respectively. The remaining significant supporting analyses are covered in the following paragraphs. Each analysis is presented separately with individual findings and conclusions stated when appropriate.
4.1 Membrane and joint Heating Due to Impinging H2 jet from External Tank
(a) Model Description
The temperature distributions on the SRM membrane and clevis joint were computed using two separate math models. A standard plume model was used to calculate jet geometry from a 0.72-inch diameter hole in the LH2 tank ideally located (direct impingement) for SRB heating. This was a highly conservative approach based on film evidence of leak location. It was determined that direct plume impingement was limited to the first 21 seconds of flight due to interaction of the airstream and that the maximum heating location in the plume would sweep aft with time along the SRM. A schematical representation is given in Figure B.86. A maximum leak rate was chosen based on data and analysis of LH2 pressure during the 51-L mission. The predicted jet results were then used with a direct impingement heating model and a thermal conduction model to establish membrane and joint temperatures.
(b) Membrane and Clevis Temperatures
All analyses of the jet provided strong evidence that even with an ever present ignition source, complete combustion of the H2 was not possible. For conservatism in the thermal analysis, only the fully burned portion of the H2 jet was considered, and cooling due to circumferential heat conduction was eliminated. The analyses assumed an LH2 leak rate of 4.0 lbm/sec. The location of the initial jet impingement was varied between point A and point B, shown in Figure B.87, to assess the effect on membrane and joint temperatures. The maximum outer clevis temperature of 910°F resulted from the impingement at point B, and the maximum temperatures of the tang (node 161) adjacent to the O-ring of 360°F resulted from impingement at point A. A case was also analyzed which assumed direct impingement of a jet of liquid hydrogen on the joint. For this condition the minimum temperature of the tang node adjacent to the O-ring was -210°F. Tang node 161 bounds the O-ring temperature, but due to the highly transient response to the jet over the 21 seconds of total impingement, and to the simplified tang used in this area, the temperatures are conservative relative to the O-ring.
The assessment of the potential for a leak from the External Tank provided substantial evidence that a leak was improbable and conservative analyses indicated that a burning jet would not produce excessive O-ring temperatures. It was concluded that a more refined analysis was not necessary.
4.2 SRM Insulation and Inhibitor Flaws
(a) Analysis Description
The analysis consisted of: (1) definition of possible mechanisms by which internal convective heat transfer could cause abnormal heating to the joint or membrane; (2) definition of insulation or inhibitor flaws or debonds which would result in abnormal heating; and (3) evaluation of the identified heat transfer or failure mechanisms by flow and thermal analyses, based on flight and experimental data. All failures were evaluated on the assumptions that: (1) enough heating could be generated to either damage the O-ring to the point of leakage (1,000°F) or bum through the joint or membrane in 58 seconds; (2) the bum through would occur near the aft field joint; and (3) no effect on motor internal pressure would be observed (no deviation was observed prior to 60 seconds on 51-L).
The nominal convective and radiative heat transfer rates were obtained by considering the measured material loss rate in the SRM aft field joint slot along with the correlation of erosion rate to convective heat transfer coefficient from "char-motor" tests (Pendleton, S. B.; Evaluation of Elastomeric Insulation Materials at High Chamber Pressures, Morton Thiokol TWR-3896). The effects of nominal radiative and convective heating on the aft field joint O-rings and metal were evaluated using a two-dimensional thermal model of the joint. A simplified ballistics analysis was performed to estimate the effect of an inhibitor failure on motor internal pressure. A simplified flow analysis was performed to evaluate the possibility that an inhibitor failure or uneven burning of an inhibitor could create circumferential flow of hot gas in the O-ring groove, causing erosion of the O-ring in approximately 58 seconds. The SRM aft field joint configuration is depicted in Figure B.88.
(b) Findings
A failure due to increased convective heating on the joint was assessed and found to be highly improbable. A discussion of the flow disturbances depicted in Figures B.89 and B.90 has been included in Section IV.A.1 of this report (Failure Scenario 1).
Insulation flaws were examined to determine the probability of overheating the O-ring or joint with nominal convective and radiative heat transfer. A convective heating rate (based on "char-motor" data and an observed erosion rate in the SRM aft field joint slot of 5.4 mils per second) of 380 BtU/ft2-sec and a radiation heating rate of 480 BtU/ft2-sec were assumed on the metal surface in a two-dimensional thermal model of the joint. It was assumed that a failure would occur if the primary O-ring temperature reached 1,000°F or if a burn through occurred. Testing conducted at Morton Thiokol as part of this failure investigation indicated that O-ring failure occurs above 1,000°F. An example of the magnitude of a joint insulation flaw or debond has been included in Section IV.A.1 of this report (Failure Scenario 1) and indicates that an extremely large area of missing insulation would be required to precipitate a joint failure prior to 60 seconds.
A relatively small flaw the forward-facing inhibitor could cause ignition of the propellant behind the inhibitor which would bum back prematurely, erode through the case insulation, and bum through the case. The effects of a forward-facing inhibitor [L178] failure was evaluated by assuming a 1-inch-diameter hole at the base of the inhibitor as shown in Figure B.91. The burnback profile due to the inhibitor failure is shown in Figure B.92. Figure B.93 shows the effect of the described failure on internal motor pressure and indicates that a failure of this type would not be detectable on measured motor pressure. The time required for the burning propellant to erode through the insulation to the metal was estimated by assuming an erosion rate based on the maximum observed erosion rate near the aft field joint of 7 mils per second. The times required for the propellant to bum back and through the insulation are shown in Table B.16. The locations in Table B. 16 are at the point where the case insulation changes thickness. The last entry in Table B.16 assumes a + 3-sigma variation on V44 insulation erosion rate and could be assumed to represent either variation in the nominal erosion characteristics of the NBR or a higher heating rate than all previous records. The + 3-sigma case represents an erosion rate typical in the aft dome area where maximum heating rates occur. It is highly improbable that a hole in the forward-facing inhibitor could exist since it would require a failure of eight plies of material.
.
Location
.
Erosion at 2 inches from Inhibitor
Erosion at 5.5 inches
Erosion at 8.6 inches
Erosion at 12.6 inches
3-sigma erosion at 8.6 inches
Another result of the inhibitor failure described above would be to raise the pressure in the slot locally, due to the increased gas outflow through the slot. A persistent circumferential pressure difference could drive a steady flow of hot gas through a putty blow hole, past some length of primary O-ring, and back into the chamber through a second blow hole, causing seal failure by erosion. An analysis was performed to determine the magnitude of this circumferential pressure gradient, based on experimental pressures around solid spoilers and jet spoilers on flat plates. These data and the dynamic pressure in the SRM yield an estimate of the pressure difference between the gap and the central motor chamber (API). The effect of the additional burning in the slot is conservatively estimated as AP - 0.2 (API). It is proportional to the dynamic pressure in the central chamber, which decreases with time leading to the results shown in Table B.17. An estimate of the circumferential pressure required to force gas flow through the O-ring cavity, at a rate sufficient to erode the primary O-ring 0.120 inch in 60 seconds was estimated to be approximately 0.5 psid.
Time (sec)
Delta p (psia)
A failure at the base of the castable inhibitor could cause rapid ignition of the propellant above the split-flap to the stress relief bulb (approximately 14 inches from the joint slot). This failure would result in initially high convective heating rates on the insulation at the base of the castable inhibitor as well as high impingement heating rates on the forward-facing inhibitor across the slot. Figure B.94 depicts a failure of this type. An instantaneous propellant ignition to the flap bulb and 12-inch circumferential width was assumed at SRM ignition. It was assumed that the split-flap was not present and that the EPDM insulation beneath the flap had the same erosion characteristics as the V44 NBR insulation. The insulation erosion rate is shown in Table B.18. At approximately 12 seconds, due to propellant burn-back and insulation erosion, the erosion rate is reduced to a nominal material loss rate of approximately 7.5 mils/sec. The total insulation loss in 60 seconds is estimated to be 0.57 inch. The minimum insulation thickness is 1.8 inches. At the point of maximum heating, the insulation thickness is greater than 2 inches. The impingement erosion on the forward-facing inhibitor is estimated to be between 0.21 and 0.4 inch. Therefore, this failure scenario is not probable.
.
(c) Conclusions
A probable failure mechanism has not been identified which would cause a burn through prior to 58 seconds due to increased convective heating, insulation flaws, or inhibitor flaws and be consistent with flight observations.
4.3 Unseating of Secondary O-ring Due to Ice in the joint
(a) Model Description
The movement of freezing water into the secondary O-ring was evaluated by hypothesizing a water level within the joint consistent with previous observations and the 38 day environmental conditions experienced by 51-L on the launch pad. The assumed conditions were used in a standard thermal analyzer to predict metal and liquid temperatures as a function of time. The volumetric expansion of the freezing liquid was assessed with regard to the pressure exerted on the secondary O-ring. The effect of vehicle tilt was also considered by taking Inertial Measurement Unit (IMU) data and accounting for the tilt with regard to water location. Capillary effects were also considered but found to be significant only if coupled with other effects. Discussions of ice-in-joint scenarios are included in the main report.
(b) Analysis Results
The thermal math model representing nodes at the aft field Joint is shown in Figure B.95. The ambient environmental conditions for STS 51-L on the night of January 28, 1986, were input into the model. Liquid water was assumed to fill the outer clevis cavity and to fill to the top of the pins in the inner cavity. The freezing process at the coldest point on the SRB started approximately 10 hours before launch (1:30 a.m. EST) and took about 3 hours to totally freeze the joint, Figures B.96 and B.97. In order to account for the increased heat of fusion during freezing, it was necessary to model that process as if it started at 32°F and completed at 31°F. The first water to freeze was at the junction of the tang and the outer clevis, node 380. It took about 2.4 hours for that process to occur. This effectively formed an ice plug which sealed the joint, thereby forcing subsequent freezing water/ice down the outer cavity and up the inner cavity toward the O-ring. The next point to freeze was at the pin in the inner clevis, node 340, about 0.4 hours later. The final freezing in the clevis cavity was in the semicircular base, node 355, and that occurred about 0.6 hours after freezing of the initial ice plug. It is theoretically possible that some movement of water/ice mixture from the lower clevis cavity could occur in the circumferential direction toward warmer regions of the joint. However, the total time for freezing was three hours, compared to the time difference of 0.2 hour between total freezing at nodes 340 and 335. Therefore, the freezing process would occur with nearly imperceptible fluid motion. Furthermore, Figure B.98 shows the predicted start and complete freeze time as a function of SRB location. The joint regions to the northerly side freeze first; [L179] therefore, any displaced liquid would tend to collect on the southerly side away from the region of joint failure.
The volumetric expansion of water to ice is sufficient to unseat the secondary O-ring when the expansion path is blocked by a grease plug or if the cavity is filled with liquid water. The freezing process starts at the tang/clevis outer junction, thereby forming an ice plug. This causes the remaining water-ice to stroke down the outer cavity and then up the inner cavity toward the secondary O-ring. For zero tilt, the ice could be 0.17 to 0.30 inches from the secondary, O-ring (0.17 inch for minimum and 0.30 inch for nominal tang/clevis clearances).
A relatively small vehicle tilt of approximately 0.3° could position water locally at the secondary O-ring. Furthermore, because a column of liquid would grow toward the O-ring during the freezing process, vehicle tilt between 0.1° to 0.2° could result in ice at the O-ring. IMU alignment at the time of vehicle launch showed a vehicle tilt of 0.24° toward the east-southeast. This magnitude of tilt could give ice problems to the secondary O-ring; however, the tilt is approximately 180° opposite the area of prime interest on the right SRB. The history of vehicle tilt during the critical freezing period is not currently available but is being investigated. Because the available IMU data is of the same magnitude of calculated tilt limits, tilt should be considered as a contributor for possible secondary O-ring unseating.
4.4. Putty Flow Analysis
(a) Model Description
An analytical assessment of putty behavior as a result of motor pressurization was formulated to supplement putty testing and establish parametric trends. To the extent consistent with meaningful results, the model complexity was minimized in order to allow Clear interpretation of trends and adaptability to test results.
The initial model consists of two-dimensional, laminar, viscous flow between flat plates, coupled with isentropic compression of an isolated, trapped volume. The transient motor pressure rise rate is imposed as a boundary condition, and the trapped volume is incremented according to the expected pressure dependent joint expansion and rotation. Parameters varied include putty viscosity, gap thickness, and length of flow path. The putty was addressed as a homogeneous, Newtonian fluid: the viscosity as a function of temperature and moisture conditioning was taken from experimental measurements of complex viscosity performed by MSFC Materials and Processes Laboratory, Table B.19.
T (F)
.
5
14
23
43
61
77
(b) Trends Identified
Qualitatively, the trends identified were as anticipated; i.e., the time lag between putty/O-ring cavity pressure and motor pressure increases with increasing putty viscosity, and increasing flow path length, and the time lag decreases with increasing gap thickness. Quantitatively, the lag times are of the same order but somewhat larger than data derived from tests at MTI. For example, at 75°F, 100 percent humidity conditioning, 0.070-inch gap, and 5-inch length, the calculated lag time to 20 psi is 1.750 seconds (Figure B.99); the corresponding lag time derived from test data is 0.275 second. This may be the result of the simplicity of the analytical model or the accuracy of putty viscosity data. But it is more likely due to the deviation of the putty test phenomena from ideal putty slug flow.
(c) Activities Planned
Further investigation of the matching of analytical results to test data is necessary to confidently apply results to SRM motor phenomena. Parametric characterization of putty flow trends will be established and documented.
4.5 O-Ring Erosion Assessment
(a) Analysis Description
An O-ring erosion parametric analysis was based on the 0ring erosion model of Salita (Prediction of Pressurization and Erosion of the SRB Primary O-rings During Motor Ignition). The model assumes that hot gas flows into this primary O-ring cavity through a putty blow hole and impinges on the primary O-ring causing erosion. The blow hole is treated as an orifice by which hot gas flows and pressurizes the O-ring cavity. The cavity pressurization model is a "lumped parameter" model which determines the cavity pressure and temperature as a function of time based on flow into and out of the cavity, the change in cavity volume as a function of motor pressure, heat transfer to the metal surfaces. and mixing of the propellant exhaust with the air which initially fills the cavity. The impingement erosion of the O-ring is correlated to an impingement heat transfer coefficient by subscale tests and comparison to char-motor testing of a similar material.
The O-ring erosion model has also been used to estimate erosion due to blow by of the primary seal. An O-ring leakage area is assumed and the blow by erosion is estimated by applying a Dittus-Boelter pipe-flow, heat transfer coefficient to the erosion correlation. If the secondary O-ring is assumed to seal. the primary leakage flow pressurizes the secondary O-ring cavity and erodes the secondary O-ring by impingement erosion.
The time to heat the O-ring surface to the ablation temperature (1,000°F) and the O-ring bulk temperature transient were estimated by using the transient flow rates from this cavity pressurization model described above in a two-dimensional thermal model of the field joint. putty blow hole, and O-rings. This leakage area of the primary and secondary O-rings was based on structural analyses of the tang/clevis radial gap opening transient and resiliency tests of the O-rings at 25°F.
(b) Findings
(1) Impingement Erosion
The relationship of impingement erosion to putty blow hole width and gas heat loss to the metal is shown in Figure B.100. Blow hole observations from flight motor segment disassembly are also shown. The parameter analysis bounds the available erosion data from flight if worst case heat transfer is assumed; i.e., the gas is immediately cooled to the ambient metal temperature which slows cavity pressurization and increases the O-ring exposure to erosion. Although the analysis predicts that impingement erosion would always occur if a blow hole exists (unless the blow hole is very large), putty blow holes have been observed with no indication of O-ring erosion. The analysis would not predict that an O-ring could be eroded to failure unless the putty blow hole was much smaller than those observed on disassembly of flight motor. Subscale tests at Morton Thiokol have indicated that an O-ring can sustain approximately 130 mils of impingement erosion without leakage.
(2) Delayed Putty Blowthrough
The parametric analysis described in the previous paragraph assumed that the putty flow path existed at SRM ignition. The effort of delayed putty blowthrough on primary O-ring erosion due to hot gas impingement was evaluated by allowing motor pressure to reach steady-state before primary O-ring cavity pressurization began. The impingement erosion analysis could not be calibrated for delayed putty blowthrough because very small blow holes (less than or equal to 0.25 inch) had to be assumed to match flight erosion data. The model always estimated less erosion for a delayed blowthrough because the primary O-ring cavity pressurizes much more rapidly than during the motor ignition transient, decreasing the erosion exposure time. This [L180] analysis indicates that: (1) the observed flight O-ring erosion was due to putty flow paths which existed at SRM ignition (formed during assembly or during the. ignition transient); and (2) cases of observed blow holes with no O-ring erosion could be due to delayed putty blowthrough.
(3) O-Ring Heating Transient
The primary and secondary radial gap opening transients were assumed to be those shown in Figure B.101 where the initial condition is metal-to-metal (maximum O-ring squeeze). The O-rings were assumed to respond as indicated by resiliency data taken at a temperature of 23°F. The O-ring leakage area was assumed to be 0.773 inch in circumferential width with a radial height calculated by subtracting the O-ring movement (from resiliency data) from the tang/clevis gap opening at the primary and secondary O-ring locations.
The O-ring surface heating transient is shown in Figure B.102. The heating due to blow by is delayed approximately 0.130 second since the O-rings are able to track the tang/clevis gap opening. The time to raise the surface temperature of this O-rings to the 1,000°F ablation temperature is approximately 0.100 second after leakage begins. Figure B.103 shows the primary and secondary bulk temperature transients. The initial O-ring temperature of 28°F was based on 51-L prelaunch predictions.
The analysis indicates that O-ring erosion would not start until 0.250 second into the SRM ignition transient. Damage to the O-rings due to surface erosion would occur long before the bulk temperature of the O-rings could respond. Cold gas tests are therefore appropriate to evaluate O-ring leakage due to reduced :resiliency at low temperatures.
Figure B.86. SRB External Heating Due to Burning of Hydrogen from 4 lb/Sec ET Leakage.
Figure B.87. Burning H2 Plume Impingement Heating to Clevis.
Figure B.88. Aft Segment/Aft Center Segment Field Joint Configuration.
Figure B.89. Circumferential Vortex + Figure B.90. Acoustic Noise.
Figure B.91. Assumed Inhibitor Flaw.
Figure B.92. Aft Segment Joint Propellant Burnback with Assumed Inhibitor Burnthrough at SRM Ignition.
Figure B.93. Increase in Motor Pc Due to One-Inch-Diameter Inhibitor Flaw.
Figure B.94. Castable Inhibitor/Split-Flap Failure.
Figure B.95. Ice in Joint Thermal Math Model.
Figure B.96. Clevis 3D-Water in Clevis Joint, Ice in Joint, Case and Water Temperatures.
Figure B.97. Clevis 3D-Water in Clevis Joint, Insulated Case, Case and Water Temperatures.
Figure B.98. Joint Freeze History.
Figure B.99. Putty Analysis.
Figure B.100. Primary O-ring Erosion Due to Hot Gas Impingement.
Figure B.101. Tang to Clevis Radial Gap.
Figure B.102. O-Ring Surface Temperature Transient Prediction.
Figure B.103. O-Ring Bulk Temperature Transient Prediction.

外殻は厚さ13mmの鋼鉄でできている。

SRB is repeatedly used, and for example, the 2009 test launched Ares I-X rocket was built together with parts of the SRB used in the last 48 shuttle flights, including those used in the 1981 first flight (STS-1) NASA Ares I First Stage Motor to be Tested August 25, NASA, July 20, 2009.

NASA Ares I First Stage Motor to be Tested August 27
NASA and ATK unveil the completed Ares I first stage five-segment solid rocket booster today in Promontory, Utah. The completed solid rocket booster is now installed horizontally in a test stand that was modified from the space shuttle's four-segment configuration to fit the new five-segment Ares I booster. Instrumentation will be installed over the next month in preparation for the first major ground test of the NASA Constellation program August 27.
NASA's Ares I launch vehicle will launch the future explorers in the Orion spacecraft to the International Space Station, the moon and beyond. The upcoming test will provide valuable thrust, roll-control, acoustics and vibration data as engineers continue to design the Ares I rocket.
"Generating 3.6 million pounds of maximum thrust at liftoff, Ares I first stage provides the backbone of NASA's next-generation rocket," said Alex Priskos, first stage manager for the Ares Projects Office at NASA's Marshall Space Flight Center in Huntsville, Ala. "The planned two-minute test will be an awesome display of smoke and fire, a live testament reflecting the dedication, hard work and achievement of the people on this project. The entire Ares team is looking forward to the first development motor firing next month."
The Ares I first stage has been under development since 2005. Based on the design of the space shuttle's four-segment booster, the first stage differs from its predecessor in a few ways. These include the addition of a fifth segment, changes to the propellant grain, a larger nozzle opening and upgraded insulation and liner.
The propellant and cases remain the same as those used for more than three decades on the Space Shuttle Program. In fact, the cases used in this Ares I first stage ground test have collectively flown on 48 previous shuttle missions, including STS-1, the very first flight.
"Because we are using shuttle boosters in the design of the Ares I first stage, we are able to leverage decades of ground tests and shuttle flight processes," said Mike Kahn, executive vice president, ATK Space Systems. "Our streamlined processes, quality and infrastructure, combined with the knowledge and lessons learned that have been transferred to the Ares I first stage program, go a long way to ensure mission success for the human spaceflight program."
ATK is the prime contractor for the first stage five-segment solid rocket booster.
Media Contacts:
Jennifer Morcone, Marshall Space Flight Center, Huntsville, Ala.
Jennifer.J.Morcone@nasa.gov

Question: what is the exterior shell of the ares i rocket made of

SRB is repeatedly used, and for example, the 2009 test launched Ares I-X rocket was built together with parts of the SRB used in the last 48 shuttle flights, including those used in the 1981 first flight (STS-1) NASA Ares I First Stage Motor to be Tested August 25, NASA, July 20, 2009.

what is the exterior shell of the ares i rocket made of

Question: when was the ares i-x launched

Question: when was the ares i-x rocket launched

Question: how thick is the exterior shell of the ares i rocket

how thick is the exterior shell of the ares i rocket

Question: when was the first flight of the ares i rocket

when was the first flight of the ares i rocket

Question: what is the name of the rocket used in the last 48 shuttle flights

what is the name of the rocket used in the last 48 shuttle flights

SRB is repeatedly used , and for example , the 2009 test launched Ares I - X rocket was built together with parts of the SRB used in the last 48 shuttle flights , including those used in the 1981 first flight ( STS-1 ) NASA Ares I First Stage Motor to be Tested August 25 , NASA , July 20 , 2009 .

外殻は厚さ13mmの鋼鉄でできている。 SRBは何度も再使用されるもので、一例を挙げれば2009年に試験発射されたアレスI-Xロケットは、過去48回のシャトルの飛行で使用されたSRBの部品を寄せ集めて作られたものであり、その中には1981年の初飛行 (STS-1) で使われたものも含まれていたNASA Ares I First Stage Motor to be Tested August 25. NASA, July 20, 2009.。

外殻は厚さ13mmの鋼鉄でできている。

SRBは何度も再使用されるもので、一例を挙げれば2009年に試験発射されたアレスI-Xロケットは、過去48回のシャトルの飛行で使用されたSRBの部品を寄せ集めて作られたものであり、その中には1981年の初飛行 (STS-1) で使われたものも含まれていたNASA Ares I First Stage Motor to be Tested August 25. NASA, July 20, 2009.。

The exterior shell is made of 13mm thick steel. SRB is repeatedly used, and for example, the 2009 test launched Ares I-X rocket was built together with parts of the SRB used in the last 48 shuttle flights, including those used in the 1981 first flight (STS-1) NASA Ares I First Stage Motor to be Tested August 25, NASA, July 20, 2009.

SRB is repeatedly used, and for example, the 2009 test launched Ares I-X rocket was built together with parts of the SRB used in the last 48 shuttle flights, including those used in the 1981 first flight (STS-1) NASA Ares I First Stage Motor to be Tested August 25, NASA, July 20, 2009.

The shuttle uses five AP-101 computers from IBM, each of which is independent and long-lasting, and a 32-bit multi-computer that composes a built-in system.

Four of them operate on a specialized software called the Primary Avionics Software System (PASS), while the rest use a different software called the Backup Flight System (BFS).

“Implementing Space Shuttle Data Processing System Concepts in Programmable Logic Devices”
Roscoe C. Ferguson, Robert Tate, and Hiram C. Thompson
United Space Alliance
Abstract
Since 1981, the Space Transportation System (STS) or Space Shuttle has been the workhorse of the United States Man Space Program. Its primary and backup flight control system (consisting of five computers, software, and a network of twenty-four buses) was considered to be state of the art at it’s time of inception. It was one of the first successful implementations of a redundant digital fly by wire avionics system and has exhibited a near perfect performance record over its service life. While the heart of the Space Shuttle avionics system is implemented in out-dated and hard to maintain technology, most of the design concepts are still relevant for modern avionics systems. This same problem will exist for new systems designed today that must provide a long service life. Programmable Logic Devices (PLD) offer promise in the ability to preserve existing assets and the flexibility required to evolve system concepts over time. To investigate the promise of PLDs, two key Space Shuttle avionics concepts were prototyped in Field Programmable Gate Arrays (FPGA).
Each of the five Space Shuttle data processing units is an IBM AP101S General Purpose Computer (GPC). Each computer consists of a Central Processing Unit (CPU) and an Input/Output Processor (IOP). The CPU performs all “data crunching” and receives its input from and performs output to the IOP which is connected to each of the twenty-four data buses. To support redundancy management, the GPCs are synchronized using a hardware discrete interface (each GPC sends a three bit discrete code to the others and receives the corresponding code from four external GPCs) and software sync loop algorithms in the Flight Computer Operating System (FCOS).
The first concept prototyped was the Space Shuttle primary flight computer. The CPU and IOP of the AP101S were implemented in an Altera Stratix FPGA. The flight software from the STS-114 mission was successfully executed on this FPGA. This provides an example of the capability to preserve existing assets. The second concept prototyped was the Space Shuttle computer synchronization concept. The FCOS software sync algorithms were implemented in hardware to produce a prototype hardware synchronization chip than can support a four computer redundant set. This chip replaces the hardware discrete interface with a serial bus interface and uses “messages” instead of parallel discrete signals. This change allows for the built in exchange and voting of state data. Each chip provides an interface for a host processor. The host processor can provide state data to the synchronization chip for voting per frame (50 Hz) and receives interrupts from the chip to indicate the start of a minor frame along with information including major/minor cycle number, the results of data state voting (including isolation) and sync failure indication. The prototype synchronization chip was interfaced to a NIOS II RISC processor executing applications running on top of a MicroC/OS-II RTOS on an Altera Stratix FPGA. A configuration of two NIOS II processors were synchronized for days using the synchronization chips without failure.
2006 MAPLD International Conference Home Page

Question: how many ap-101 computers does the shuttle use

The shuttle uses five AP-101 computers from IBM, each of which is independent and long-lasting, and a 32-bit multi-computer that composes a built-in system.

Four of them operate on a specialized software called the Primary Avionics Software System (PASS), while the rest use a different software called the Backup Flight System (BFS).

how many ap-101 computers does the shuttle use

Question: what doesPASS stand for

ソフトウェア システム ( Primary Avionics Software System , PASS ) と いう 特製 の ソフトウェア で 稼働 し 、 残り の 1 台 は これ と は 別 の 、 バックアップ 飛行 システム ( Backup Flight System , BFS ) と いう ソフト を 使用 し

Question: what does bfs stand for

Question: who made the ap-101 computers used by the shuttle

who made the ap-101 computers used by the shuttle

Question: what does pass stand for

ソフトウェア システム ( Primary Avionics Software System , PASS ) と いう 特製 の ソフトウェア で 稼働 し 、 残り の 1 台 は これ と は 別 の 、 バックアップ 飛行 システム ( Backup Flight System , BFS ) と いう ソフト を 使用 し

Question: what is the name of the software used on the shuttle

) と いう 特製 の ソフトウェア で 稼働 し 、 残り の 1 台 は これ と は 別 の 、 バックアップ 飛行 システム ( Backup Flight System , BFS

what is the name of the software used on the shuttle

Question: how many computers does the shuttle have

how many computers does the shuttle have

The shuttle uses five AP-101 computers from IBM , each of which is independent and long - lasting , and a 32 - bit multi - computer that composes a built - in system .

Four of them operate on a specialized software called the Primary Avionics Software System ( PASS ) , while the rest use a different software called the Backup Flight System ( BFS ) .

シャトルは IBM製の5台のAP-101と呼ばれる、それぞれ独立して冗長性を持ち、組み込みシステムを構成する32ビット汎用コンピューターを使用している。 このうち4台は主飛行電子ソフトウェアシステム (Primary Avionics Software System, PASS) という特製のソフトウェアで稼働し、残りの1台はこれとは別の、バックアップ飛行システム (Backup Flight System, BFS) というソフトを使用している。

シャトルは IBM製の5台のAP-101と呼ばれる、それぞれ独立して冗長性を持ち、組み込みシステムを構成する32ビット汎用コンピューターを使用している。

このうち4台は主飛行電子ソフトウェアシステム (Primary Avionics Software System, PASS) という特製のソフトウェアで稼働し、残りの1台はこれとは別の、バックアップ飛行システム (Backup Flight System, BFS) というソフトを使用している。

The shuttle uses five AP-101 computers from IBM, each of which is independent and long-lasting, and a 32-bit multi-computer that composes a built-in system. Four of them operate on a specialized software called the Primary Avionics Software System (PASS), while the rest use a different software called the Backup Flight System (BFS).

The shuttle uses five AP-101 computers from IBM, each of which is independent and long-lasting, and a 32-bit multi-computer that composes a built-in system.

Four of them operate on a specialized software called the Primary Avionics Software System (PASS), while the rest use a different software called the Backup Flight System (BFS).

Since November 1983, the shuttle has been using the world’s earliest laptop computer called GRiD Compass.

The compass was cheap, which was sold for a maximum of $8,000, but at the time it performed uncomparable performance compared to its weight and size, and NASA was one of its important customers.

The grid compass was not related to the flight control system and was used to show the flight tracks of the shuttle twice.

Question: what is the name of the laptop computer the shuttle uses since 1983

Since November 1983, the shuttle has been using the world’s earliest laptop computer called GRiD Compass.

The compass was cheap, which was sold for a maximum of $8,000, but at the time it performed uncomparable performance compared to its weight and size, and NASA was one of its important customers.

The grid compass was not related to the flight control system and was used to show the flight tracks of the shuttle twice.

what is the name of the laptop computer the shuttle uses since 1983

Question: how much did the gid compass cost

Question: when did nasa start using the gid compass

when did nasa start using the gid compass

Question: who was one of the first customers of the gid compass

who was one of the first customers of the gid compass

Question: how many times did nasa use the grid compass

how many times did nasa use the grid compass

Question: why did nasa use the grid compass

Question: what was the grid compass not related to

what was the grid compass not related to

Since November 1983 , the shuttle has been using the world ’s earliest laptop computer called GRiD Compass .

The compass was cheap , which was sold for a maximum of $ 8,000 , but at the time it performed uncomparable performance compared to its weight and size , and NASA was one of its important customers .

The grid compass was not related to the flight control system and was used to show the flight tracks of the shuttle twice .

1983年11月から、シャトルにはグリッド・コンパス (GRiD Compass) と呼ばれる、世界で最も初期の頃に作られたラップトップ型コンピュータを使用していた。 コンパスはせいぜい8,000ドル程度で売られていた安価なものだが、その重量や大きさに比して不釣り合いなほどの性能を当時発揮し、NASAはその重要な顧客の一つだった。 なお、グリッド・コンパスは飛行制御系統には関係せず、シャトルの飛行軌跡を2周回分表示させるのに使用された。

1983年11月から、シャトルにはグリッド・コンパス (GRiD Compass) と呼ばれる、世界で最も初期の頃に作られたラップトップ型コンピュータを使用していた。

コンパスはせいぜい8,000ドル程度で売られていた安価なものだが、その重量や大きさに比して不釣り合いなほどの性能を当時発揮し、NASAはその重要な顧客の一つだった。

なお、グリッド・コンパスは飛行制御系統には関係せず、シャトルの飛行軌跡を2周回分表示させるのに使用された。

Since November 1983, the shuttle has been using the world’s earliest laptop computer called GRiD Compass. The compass was cheap, which was sold for a maximum of $8,000, but at the time it performed uncomparable performance compared to its weight and size, and NASA was one of its important customers. The grid compass was not related to the flight control system and was used to show the flight tracks of the shuttle twice.

Since November 1983, the shuttle has been using the world’s earliest laptop computer called GRiD Compass.

The compass was cheap, which was sold for a maximum of $8,000, but at the time it performed uncomparable performance compared to its weight and size, and NASA was one of its important customers.

The grid compass was not related to the flight control system and was used to show the flight tracks of the shuttle twice.

As the Shuttle is a spacecraft developed in the 1970s, there have been many improvements and modifications since that time to improve its performance and reliability in terms of safety.

The STS-101 pilot seat.

Question: when was the space shuttle developed

As the Shuttle is a spacecraft developed in the 1970s, there have been many improvements and modifications since that time to improve its performance and reliability in terms of safety.

The STS-101 pilot seat.

Question: what is the name of the pilot seat of the space shuttle

The STS-101 pilot seat.

what is the name of the pilot seat of the space shuttle

Question: what is the name of the seat in the shuttle

The STS-101 pilot seat.

what is the name of the seat in the shuttle

Question: when was the shuttle first developed

The STS-101 pilot seat.

Question: what has the shuttle improved in terms of

The STS-101 pilot seat.

what has the shuttle improved in terms of

Question: what spacecraft was developed in the 1970s

The STS-101 pilot seat.

what spacecraft was developed in the 1970s

Question: what is the name of the seat in a shuttle

The STS-101 pilot seat.

what is the name of the seat in a shuttle

Question: what has the shuttle improved in terms of safety

The STS-101 pilot seat.

what has the shuttle improved in terms of safety

As the Shuttle is a spacecraft developed in the 1970s , there have been many improvements and modifications since that time to improve its performance and reliability in terms of safety .

改良 シャトルは1970年代に開発された宇宙船であるため、その当時から安全面における性能や信頼性を向上させるべく多くの改良や改造が施されてきた。 STS-101の操縦席。

改良 シャトルは1970年代に開発された宇宙船であるため、その当時から安全面における性能や信頼性を向上させるべく多くの改良や改造が施されてきた。

As the Shuttle is a spacecraft developed in the 1970s, there have been many improvements and modifications since that time to improve its performance and reliability in terms of safety. The STS-101 pilot seat.

As the Shuttle is a spacecraft developed in the 1970s, there have been many improvements and modifications since that time to improve its performance and reliability in terms of safety.

1.752MN per base (178 tons, 104% when it occurs) Maximum load volume: 25.060kg Cargo room size: 4.6m x 18.0m Operating height: 190 to 960km (100 to 520 marine miles) Maximum speed: 7.743km/s (27.870km/h equal to 22.57km) Circuit range: 2,009km (1.085 marine miles) Specification: varies by flight.

At the beginning it was flighted with a minimum of two people, but in many subsequent flights it became common to consist of five, then seven (captain, pilot, a few flight operating technicians, and rarely an airline engineer).

In the two flights of STS-61-A and STS-71, eight passengers were on board.

Question: how many passengers were on the flight of sts-61-a and sts

1.752MN per base (178 tons, 104% when it occurs) Maximum load volume: 25.060kg Cargo room size: 4.6m x 18.0m Operating height: 190 to 960km (100 to 520 marine miles) Maximum speed: 7.743km/s (27.870km/h equal to 22.57km) Circuit range: 2,009km (1.085 marine miles) Specification: varies by flight.

At the beginning it was flighted with a minimum of two people, but in many subsequent flights it became common to consist of five, then seven (captain, pilot, a few flight operating technicians, and rarely an airline engineer).

In the two flights of STS-61-A and STS-71, eight passengers were on board.

how many passengers were on the flight of sts-61-a and sts

Question: how many passengers are on a typical flight

how many passengers are on a typical flight

Question: what is the maximum range of a circuit plane

what is the maximum range of a circuit plane

Question: how many passengers were on the first flight of sts-61

how many passengers were on the first flight of sts-61

Question: what is the maximum load volume of a cargo plane

what is the maximum load volume of a cargo plane

Question: what is the maximum load per base

Question: how many passengers are on a plane in the beginning

how many passengers are on a plane in the beginning

1.752MN per base ( 178 tons , 104 % when it occurs ) Maximum load volume : 25.060 kg Cargo room size : 4.6 m x 18.0 m Operating height : 190 to 960 km ( 100 to 520 marine miles ) Maximum speed : 7.743km / s ( 27.870 km/h equal to 22.57 km ) Circuit range : 2,009 km ( 1.085 marine miles ) Specification : varies by flight .

At the beginning it was flighted with a minimum of two people , but in many subsequent flights it became common to consist of five , then seven ( captain , pilot , a few flight operating technicians , and rarely an airline engineer ) .

In the two flights of STS-61 - A and STS-71 , eight passengers were on board .

1基あたり海面推力1.752MN（178トン、104%推力発生時） 最大搭載量：25,060kg 貨物室寸法：4.6m×18.0m 運用高度：190〜960km（100〜520海里） 最大速度：秒速7.743km(時速27,870km マッハ22.57相当) 軌道範囲：2,009km（1,085海里） 定員：飛行によって異なる。 初期の頃は最小人員の2名で飛行したが、後の多くの飛行では5名になり、その後7名（船長、パイロット、数人の搭乗運用技術者、まれに航空機関士(フライトエンジニア)）で構成するのが一般的になった。 STS-61-AとSTS-71の2回の飛行では8名が搭乗した。

1基あたり海面推力1.752MN（178トン、104%推力発生時） 最大搭載量：25,060kg 貨物室寸法：4.6m×18.0m 運用高度：190〜960km（100〜520海里） 最大速度：秒速7.743km(時速27,870km マッハ22.57相当) 軌道範囲：2,009km（1,085海里） 定員：飛行によって異なる。

初期の頃は最小人員の2名で飛行したが、後の多くの飛行では5名になり、その後7名（船長、パイロット、数人の搭乗運用技術者、まれに航空機関士(フライトエンジニア)）で構成するのが一般的になった。

1.752MN per base (178 tons, 104% when it occurs) Maximum load volume: 25.060kg Cargo room size: 4.6m x 18.0m Operating height: 190 to 960km (100 to 520 marine miles) Maximum speed: 7.743km/s (27.870km/h equal to 22.57km) Circuit range: 2,009km (1.085 marine miles) Specification: varies by flight. At the beginning it was flighted with a minimum of two people, but in many subsequent flights it became common to consist of five, then seven (captain, pilot, a few flight operating technicians, and rarely an airline engineer). In the two flights of STS-61-A and STS-71, eight passengers were on board.

1.752MN per base (178 tons, 104% when it occurs) Maximum load volume: 25.060kg Cargo room size: 4.6m x 18.0m Operating height: 190 to 960km (100 to 520 marine miles) Maximum speed: 7.743km/s (27.870km/h equal to 22.57km) Circuit range: 2,009km (1.085 marine miles) Specification: varies by flight.

At the beginning it was flighted with a minimum of two people, but in many subsequent flights it became common to consist of five, then seven (captain, pilot, a few flight operating technicians, and rarely an airline engineer).

In the two flights of STS-61-A and STS-71, eight passengers were on board.

There should not be any rainfall around the launch station or on the flight route.

Temperatures must be less than 2°C or 37°C.

There should be no clouds that hide the image of the aircraft until it climbs to 2,400 meters.

Question: what is the altitude of the plane at launch

2,400 m に 上昇 する まで の 間 に 機体 の 姿 を 隠し て しまう よう な 雲 が 存在 し

There should not be any rainfall around the launch station or on the flight route.

Temperatures must be less than 2°C or 37°C.

There should be no clouds that hide the image of the aircraft until it climbs to 2,400 meters.

what is the altitude of the plane at launch

Question: what should not be around the launch station or on the flight route

what should not be around the launch station or on the flight route

Question: what is the minimum temperature for the flight

what is the minimum temperature for the flight

Question: what is the minimum temperature needed for the flight

what is the minimum temperature needed for the flight

Question: what hide the image of the aircraft until it climbs to 2,400 meters

what hide the image of the aircraft until it climbs to 2,400 meters

Question: what must be less than 2 degrees celsius

what must be less than 2 degrees celsius

Question: what should hide the image of the aircraft until it reaches 2,400 meters

が 存在 し て は なら ない

what should hide the image of the aircraft until it reaches 2,400 meters

There should not be any rainfall around the launch station or on the flight route .

Temperatures must be less than 2 ° C or 37 ° C .

There should be no clouds that hide the image of the aircraft until it climbs to 2,400 meters .

発射台周辺や飛行経路に一切の降雨があってはならない。 気温は摂氏2℃以上37℃以下でなければならない。 高度2,400mに上昇するまでの間に機体の姿を隠してしまうような雲が存在してはならない。

発射台周辺や飛行経路に一切の降雨があってはならない。

気温は摂氏2℃以上37℃以下でなければならない。

高度2,400mに上昇するまでの間に機体の姿を隠してしまうような雲が存在してはならない。

There should not be any rainfall around the launch station or on the flight route. Temperatures must be less than 2°C or 37°C. There should be no clouds that hide the image of the aircraft until it climbs to 2,400 meters.

There should not be any rainfall around the launch station or on the flight route.

Temperatures must be less than 2°C or 37°C.

There should be no clouds that hide the image of the aircraft until it climbs to 2,400 meters.

Especially if there is a possibility of flooding, the shuttle will never be launched.

The aircraft may often suffer direct storms, but the aircraft will not be affected by the fact that the structure is a conductor or that the current is released into the air because it is not electrically connected.

Question: what is the structure of an aircraft made of

伝導 性 の アルミニウム

Especially if there is a possibility of flooding, the shuttle will never be launched.

The aircraft may often suffer direct storms, but the aircraft will not be affected by the fact that the structure is a conductor or that the current is released into the air because it is not electrically connected.

what is the structure of an aircraft made of

Question: the structure of an aircraft is made of what material

the structure of an aircraft is made of what material

Question: what type of storms can an aircraft suffer

what type of storms can an aircraft suffer

Question: what is the most likely event that will cause the shuttle to never be launched

what is the most likely event that will cause the shuttle to never be launched

Question: why is the shuttle not affected by storms

why is the shuttle not affected by storms

Question: why will the shuttle never be launched

落雷 が 起きる 可能 性 が ある 場合

Question: the explosion that is spread during the launch will play the role of what

electric line that connects the aircraft with the ground

機体 と 地面 を つなぐ 電線

electric line that connects the aircraft with the ground

the explosion that is spread during the launch will play the role of what

Question: what connects the aircraft with the ground

what connects the aircraft with the ground

Especially if there is a possibility of flooding , the shuttle will never be launched .

The aircraft may often suffer direct storms , but the aircraft will not be affected by the fact that the structure is a conductor or that the current is released into the air because it is not electrically connected .

On the contrary , the shuttle , the structure of the aircraft is made of conductive aluminum as the usual jet passenger aircraft , so the internal equipment will not be affected by the electric current , but the explosion that is spread during the launch will play the role of the electric line that connects the aircraft with the ground .

特に落雷が起きる可能性がある場合には、シャトルは絶対に発射されない。 航空機はしばしば雷の直撃を受けることがあるが、構造が伝導体であることや、電気的に接地されていないために電流が空気中に放電されることなどにより、機体が悪影響を受けることはない。 これに対してシャトルは、機体構造は通常のジェット旅客機と同じように伝導性のアルミニウムで作られているので内部機器が電流の影響を受けることはないが、発射時に噴射される噴煙が機体と地面をつなぐ電線の役目を果たしてしまう。

特に落雷が起きる可能性がある場合には、シャトルは絶対に発射されない。

航空機はしばしば雷の直撃を受けることがあるが、構造が伝導体であることや、電気的に接地されていないために電流が空気中に放電されることなどにより、機体が悪影響を受けることはない。

これに対してシャトルは、機体構造は通常のジェット旅客機と同じように伝導性のアルミニウムで作られているので内部機器が電流の影響を受けることはないが、発射時に噴射される噴煙が機体と地面をつなぐ電線の役目を果たしてしまう。

Especially if there is a possibility of flooding, the shuttle will never be launched. The aircraft may often suffer direct storms, but the aircraft will not be affected by the fact that the structure is a conductor or that the current is released into the air because it is not electrically connected. On the contrary, the shuttle, the structure of the aircraft is made of conductive aluminum as the usual jet passenger aircraft, so the internal equipment will not be affected by the electric current, but the explosion that is spread during the launch will play the role of the electric line that connects the aircraft with the ground.

Especially if there is a possibility of flooding, the shuttle will never be launched.

The aircraft may often suffer direct storms, but the aircraft will not be affected by the fact that the structure is a conductor or that the current is released into the air because it is not electrically connected.

For a long time, the shuttle was unable to fly around December 31 and January 1.

The software for shuttle, developed in the 1970s, has not been designed to be possible over the years, because if you force it during the flight you will have to re-set your computer, causing the possibility of unpredictable errors.

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Question: when was the software for the shuttle developed

For a long time, the shuttle was unable to fly around December 31 and January 1.

The software for shuttle, developed in the 1970s, has not been designed to be possible over the years, because if you force it during the flight you will have to re-set your computer, causing the possibility of unpredictable errors.

when was the software for the shuttle developed

Question: when was the shuttle unable to fly

Question: if you force the computer to re-set, what can happen to the software

if you force the computer to re-set, what can happen to the software

Question: what do you have to do if the shuttle can't fly

コンピューター を リセット

what do you have to do if the shuttle can't fly

Question: when does the shuttle stop running in japan

when does the shuttle stop running in japan

For a long time , the shuttle was unable to fly around December 31 and January 1 .

The software for shuttle , developed in the 1970s , has not been designed to be possible over the years , because if you force it during the flight you will have to re - set your computer , causing the possibility of unpredictable errors .

長い間、シャトルは12月31日と1月1日をまたがっては飛行できなかった。 1970年代に開発されたシャトル用のソフトウェアは年越しができるようには設計されておらず、もし飛行中にそれを強行するとコンピューターをリセットしなければならなくなり、予測できないようなエラーが発生する可能性が生じるからである。

長い間、シャトルは12月31日と1月1日をまたがっては飛行できなかった。

1970年代に開発されたシャトル用のソフトウェアは年越しができるようには設計されておらず、もし飛行中にそれを強行するとコンピューターをリセットしなければならなくなり、予測できないようなエラーが発生する可能性が生じるからである。

For a long time, the shuttle was unable to fly around December 31 and January 1. The software for shuttle, developed in the 1970s, has not been designed to be possible over the years, because if you force it during the flight you will have to re-set your computer, causing the possibility of unpredictable errors.

For a long time, the shuttle was unable to fly around December 31 and January 1.

The software for shuttle, developed in the 1970s, has not been designed to be possible over the years, because if you force it during the flight you will have to re-set your computer, causing the possibility of unpredictable errors.

16 seconds before the launch (T minus 16), the noise deterrence device is operating and the aircraft is not damaged by sharp sound and the SRB's 1,100m3 of water is released to the moving launch station and firefighting predisposition board (frame trailers). "Sound Suppression Water System" Revised 2000-08-28. Retrieved 2006-07-09.

Question: how much water is released to the moving launch station

how much water is released to the moving launch station

Question: what is the name of the water system used in helicopters

what is the name of the water system used in helicopters

Question: how long before launch does the noise deterrence device stop operating

how long before launch does the noise deterrence device stop operating

Question: what are the firefighting predisposition board called

what are the firefighting predisposition board called

Question: when was the sound suppression water system revised

when was the sound suppression water system revised

Question: what is operating 16 seconds before launch

騒音 抑制 装置 が 作動 し 、 猛烈 な 音響 で 機体 が 損傷 を 負わ ない よう に する ため に 移動 式 発射 台 や SRB の 火炎 偏向 板 （ フレーム トレンチ ） に 1,100 m³ の 水 が 放出 さ れ はじめる National Aeronautics and Space Administration . "" Sound Suppression Water System "" Revised 2000 - 08 - 28 . Retrieved

what is operating 16 seconds before launch

Question: what prevented the aircraft from being damaged by the noise deterrence device

what prevented the aircraft from being damaged by the noise deterrence device

16 seconds before the launch ( T minus 16 ) , the noise deterrence device is operating and the aircraft is not damaged by sharp sound and the SRB 's 1,100m3 of water is released to the moving launch station and firefighting predisposition board ( frame trailers ) . " Sound Suppression Water System " Revised 2000 - 08 - 28 . Retrieved 2006 - 07 - 09 .

発射16秒前（Tマイナス16）、騒音抑制装置が作動し、猛烈な音響で機体が損傷を負わないようにするために移動式発射台やSRBの火炎偏向板（フレームトレンチ）に1,100m³の水が放出されはじめるNational Aeronautics and Space Administration. "Sound Suppression Water System" Revised 2000-08-28. Retrieved 2006-07-09.。

発射16秒前（Tマイナス16）、騒音抑制装置が作動し、猛烈な音響で機体が損傷を負わないようにするために移動式発射台やSRBの火炎偏向板（フレームトレンチ）に1,100m³の水が放出されはじめるNational Aeronautics and Space Administration. "Sound Suppression Water System" Revised 2000-08-28. Retrieved 2006-07-09.。

6.6 seconds before the launch (T minus 6.6), the SSME lighting begins.

The firefighting command is sent through the orbit ship's GPC (multiple computer) by placing an interval of 120 milliseconds in order for the third engine (right), the second engine (left), and the first engine (center).

The countdown clock is one of the most-watched timepieces in the world. On this page, you'll learn how the countdown operates, and what milestones to watch for during our live launch coverage.
Pauses in the countdown, or "holds," are built into the countdown to allow the launch team to target a precise launch window, and to provide a cushion of time for certain tasks and procedures without impacting the overall schedule. For the space shuttle countdown, built-in holds vary in length and always occur at the following times: T-27 hours, T-19 hours, T-11 hours, T-6 hours, T-3 hours, T-20 minutes, and T-9 minutes.
Here are some of the key events that take place at each milestone after the countdown begins. Note: Event times and lengths are approximate and subject to change.
T-43 hours and counting
The Shuttle Test Director performs the traditional call to stations and the countdown clock is activated.
T-27 hours and holding
- Begin final vehicle and facility close-outs for launch
- Check out backup flight systems
- Review flight software stored in mass memory units and display systems
- Load backup flight system software into the orbiter's general purpose computers
- Remove middeck and flight deck platforms
- Activate and test navigational systems
- Complete preparation to load power reactant storage and distribution system
- Complete flight deck preliminary inspections
This is the first built-in hold and typically lasts four hours.
T-27 hours and counting
- Clear launch pad of all non-essential personnel
- Begin loading cryogenic propellants into orbiter's power reactant storage and distribution (PRSD) system
T-19 hours and holding
- Begin operations to load cryogenic reactants into the orbiter's fuel cell storage tanks
This built-in hold typically lasts four hours, but may be extended if PRSD offload is required.
T-19 hours and counting
- Demate the orbiter's midbody umbilical unit
- Clean and vacuum crew module
- External tank nosecone purge
T-11 hours and holding
- Begin final preparations of the orbiter's three main engines for main propellant tanking and flight
- Fill launch pad sound suppression system water tank
- Resume orbiter and ground support equipment close-outs
- Close out the tail service masts on the mobile launcher platform
This built-in hold varies between 13 to 14 hours.
T-11 hours and counting
- Weather and engineering briefings
- Pad debris inspection and closeout
- Flight crew equipment late stow
- Move rotating service structure to "park" position
- Activate the orbiter's inertial measurement units and communications systems
- Perform ascent switch list
T-6 hours and holding
- Activate the orbiter's fuel cells
- Clear the blast danger area of all nonessential personnel
- Switch the orbiter's purge air to gaseous nitrogen
This built-in hold typically lasts two hours, or one hour for a 24- or 48-hour scrub.
T-6 hours and counting
- Mission Management Team and launch director receive weather update
- Launch team verifies no violations of launch commit criteria before loading the external tank with propellants
- Clear pad of all personnel
- Chill-down of propellant transfer lines
- Begin loading the external tank with about 500,000 gallons of cryogenic propellants
T-3 hours and holding
- Finish filling the external tank with its flight load of liquid hydrogen and liquid oxygen propellants
This built-in hold typically lasts two-and-a-half hours.
T-3 hours and counting
- External tank loading enters stable replenish
- Perform inertial measurement unit preflight calibration
- Align Merritt Island Launch Area (MILA) tracking antennas
- Final Inspection Team proceeds to the launch pad to conduct a detailed analysis of the vehicle as the team walks up and down the entire launch tower
- Closeout Crew proceeds to the launch pad to configure the crew module for countdown and launch and assist the astronauts with entry into the orbiter
- Televised weather briefing
- Flight crew weather briefing
- Astronaut Support Person enters crew module and begins comm checks
T-20 minutes and holding
- Crew departs for the launch pad and, upon arriving at the pad, begins entry into the orbiter via the White Room
- Complete close-out preparations in the launch pad's White Room
- Check cockpit switch configurations
- Astronauts perform air-to-ground voice checks with Launch Control (Kennedy Space Center) and Mission Control (Johnson Space Center)
- Close the orbiter's crew hatch and check for leaks
- Complete White Room close-out
- Close-out crew retreats to fallback area
This built-in hold typically lasts 10 minutes.
T-20 minutes and counting
- NASA Test Director conducts final launch team briefings
- Complete inertial measurement unit preflight alignments
T-9 minutes and holding
- Transition the orbiter's onboard computers to launch configuration
- Start fuel cell thermal conditioning
- Close orbiter cabin vent valves
- Transition backup flight system to launch configuration
This is the final built-in hold, and varies in length depending on the mission.
T-9 minutes and counting
- Final launch window determination
- Activate flight recorders
- Final "go/no-go" launch polls conducted by NASA Test Director, Mission Management Team and launch director
T-0
- Start automatic ground launch sequencer
- Retract orbiter access arm (T-7 minutes, 30 seconds)
- Start auxiliary power units (T-5 minutes, 0 seconds)
- Arm solid rocket booster range safety safe and arm devices (T-5 minutes, 0 seconds)
- Start orbiter aerosurface profile test, followed by main engine gimbal profile test (T-3 minutes, 55 seconds)
- Retract gaseous oxygen vent arm, or "beanie cap" (T-2 minutes, 55 seconds)
- Crew members close and lock their visors (T-2 minutes, 0 seconds)
- Orbiter transfers from ground to internal power (T-50 seconds)
- Ground launch sequencer is go for auto sequence start (T-31 seconds)
- Activate launch pad sound suppression system (T-16 seconds)
- Activate main engine hydrogen burnoff system (T-10 seconds)
- Main engine start (T-6.6 seconds)
- Solid rocket booster ignition and liftoff!
NASA's John F. Kennedy Space Center

Question: how long is the firefighting command for the third engine

120 ミリ 秒 の 間隔 を

6.6 seconds before the launch (T minus 6.6), the SSME lighting begins.

The firefighting command is sent through the orbit ship's GPC (multiple computer) by placing an interval of 120 milliseconds in order for the third engine (right), the second engine (left), and the first engine (center).

how long is the firefighting command for the third engine

Question: how long before launch does the sme lighting start

how long before launch does the sme lighting start

Question: how many milliseconds does it take for the third engine to fire

how many milliseconds does it take for the third engine to fire

Question: what happens 6.6 seconds before launch

Question: what does gpc stand for

Question: where is the first engine of the japanese space shuttle

where is the first engine of the japanese space shuttle

Question: what does the sme do 6 seconds before launch

what does the sme do 6 seconds before launch

6.6 seconds before the launch ( T minus 6.6 ) , the SSME lighting begins .

The firefighting command is sent through the orbit ship 's GPC ( multiple computer ) by placing an interval of 120 milliseconds in order for the third engine ( right ) , the second engine ( left ) , and the first engine ( center ) .

発射6.6秒前（Tマイナス6.6）、SSMEの点火が始まる。 点火指令は軌道船のGPC（汎用コンピューター）を経由して、3番エンジン（右側）、2番エンジン（左側）、1番エンジン（中央）の順に120ミリ秒の間隔を置いて送られる。

発射6.6秒前（Tマイナス6.6）、SSMEの点火が始まる。

点火指令は軌道船のGPC（汎用コンピューター）を経由して、3番エンジン（右側）、2番エンジン（左側）、1番エンジン（中央）の順に120ミリ秒の間隔を置いて送られる。

6.6 seconds before the launch (T minus 6.6), the SSME lighting begins. The firefighting command is sent through the orbit ship's GPC (multiple computer) by placing an interval of 120 milliseconds in order for the third engine (right), the second engine (left), and the first engine (center).

6.6 seconds before the launch (T minus 6.6), the SSME lighting begins.

The firefighting command is sent through the orbit ship's GPC (multiple computer) by placing an interval of 120 milliseconds in order for the third engine (right), the second engine (left), and the first engine (center).

GPC reaches the push of the SSME to 90%, while at the same time fixing the direction of the nodes to the specified location by the National Aeronautics and Space Administration. "NASA - Countdown 101".

When the engine is lighted, the water of the noise deterrent device evaporates and becomes a large amount of water steam and is spread to the south.

The three-base SSME drive must then reach 100% within three seconds, and if it does not realize, the GPC will urgently stop the engine.

Question: what is the push of the sme drive in nasa countdown 101

GPC reaches the push of the SSME to 90%, while at the same time fixing the direction of the nodes to the specified location by the National Aeronautics and Space Administration. "NASA - Countdown 101".

When the engine is lighted, the water of the noise deterrent device evaporates and becomes a large amount of water steam and is spread to the south.

The three-base SSME drive must then reach 100% within three seconds, and if it does not realize, the GPC will urgently stop the engine.

what is the push of the sme drive in nasa countdown 101

Question: who fixes the direction of the nodes to the specified location

National Aeronautics and Space Administration

National Aeronautics and Space Administration . "" NASA - Countdown 101 "" . Retrieved

National Aeronautics and Space Administration

who fixes the direction of the nodes to the specified location

Question: how high must the three-base sme drive reach within 3 seconds

how high must the three-base sme drive reach within 3 seconds

Question: what is the name of the nasa countdown 101

NASA - Countdown 101 "" . Retrieved 2006 - 07 - 10 . 。

what is the name of the nasa countdown 101

Question: what direction is the water of the noise deterrent device spread to

what direction is the water of the noise deterrent device spread to

Question: what does the noise deterrent device evaporate and become

what does the noise deterrent device evaporate and become

Question: how long does it take for the three-base sme drive to reach 100%

how long does it take for the three-base sme drive to reach 100%

Question: how fast must the three-base sme drive reach in order to stop the engine

秒 以内 に 100 %

how fast must the three-base sme drive reach in order to stop the engine

GPC reaches the push of the SSME to 90 % , while at the same time fixing the direction of the nodes to the specified location by the National Aeronautics and Space Administration . " NASA - Countdown 101 " .

When the engine is lighted , the water of the noise deterrent device evaporates and becomes a large amount of water steam and is spread to the south .

The three - base SSME drive must then reach 100 % within three seconds , and if it does not realize , the GPC will urgently stop the engine .

GPCはSSMEの推力を90%にまで到達させると同時に、ノズルの向きを所定の位置に固定するNational Aeronautics and Space Administration. "NASA - Countdown 101". Retrieved 2006-07-10.。 エンジンに点火されると、騒音抑制装置の水が蒸発して大量の水蒸気となり、南側に向かって噴出される。 3基のSSMEの推力はそれから3秒以内に100%に達しなければならず、もしそれが実現しなかった場合はGPCがエンジンを緊急停止させる。

GPCはSSMEの推力を90%にまで到達させると同時に、ノズルの向きを所定の位置に固定するNational Aeronautics and Space Administration. "NASA - Countdown 101". Retrieved 2006-07-10.。

エンジンに点火されると、騒音抑制装置の水が蒸発して大量の水蒸気となり、南側に向かって噴出される。

3基のSSMEの推力はそれから3秒以内に100%に達しなければならず、もしそれが実現しなかった場合はGPCがエンジンを緊急停止させる。

GPC reaches the push of the SSME to 90%, while at the same time fixing the direction of the nodes to the specified location by the National Aeronautics and Space Administration. "NASA - Countdown 101". When the engine is lighted, the water of the noise deterrent device evaporates and becomes a large amount of water steam and is spread to the south. The three-base SSME drive must then reach 100% within three seconds, and if it does not realize, the GPC will urgently stop the engine.

GPC reaches the push of the SSME to 90%, while at the same time fixing the direction of the nodes to the specified location by the National Aeronautics and Space Administration. "NASA - Countdown 101".

When the engine is lighted, the water of the noise deterrent device evaporates and becomes a large amount of water steam and is spread to the south.

The three-base SSME drive must then reach 100% within three seconds, and if it does not realize, the GPC will urgently stop the engine.

There are many alternative landing facilities, but they were not used except the Edwards Air Force Base and the White Suns Air Force Base.

For the Edwards Air Force Base, the policy has been adopted in recent years as far as possible due to the large cost of the shuttle's transit, and the STS-131 was also determined to be used for a time when the Japanese last crew landed, but eventually landed in the Kennedy Space Center.

Question: where did the japanese last crew land

There are many alternative landing facilities, but they were not used except the Edwards Air Force Base and the White Suns Air Force Base.

For the Edwards Air Force Base, the policy has been adopted in recent years as far as possible due to the large cost of the shuttle's transit, and the STS-131 was also determined to be used for a time when the Japanese last crew landed, but eventually landed in the Kennedy Space Center.

Question: which air force base was not used by the japanese

空軍 基地 と ホワイト ・ サンズ 空軍 基地

which air force base was not used by the japanese

Question: where did the space shuttle land in the united states

エドワーズ 空軍 基地 と ホワイト ・ サンズ 空軍

where did the space shuttle land in the united states

Question: which two air force bases were not used

Edwards Air Force Base and the White Suns Air Force Base

エドワーズ 空軍 基地 と ホワイト ・ サンズ 空軍 基地

Edwards Air Force Base and the White Suns Air Force Base

Question: what shuttle landed in the kennedy space center

what shuttle landed in the kennedy space center

Question: what other air force base was not used

空軍 基地 と ホワイト ・ サンズ 空軍 基地

There are many alternative landing facilities , but they were not used except the Edwards Air Force Base and the White Suns Air Force Base .

For the Edwards Air Force Base , the policy has been adopted in recent years as far as possible due to the large cost of the shuttle 's transit , and the STS-131 was also determined to be used for a time when the Japanese last crew landed , but eventually landed in the Kennedy Space Center .

代替着陸施設は多数あるが、エドワーズ空軍基地とホワイト・サンズ空軍基地以外は使用されることはなかった。 エドワーズ空軍基地についても、シャトルの大陸横断に掛かる多額のコストのため近年はできるだけ利用しない方針が採られており、日本人最後の乗務となったSTS-131の着陸時にも一時は使用が決定していたが、最終的にはケネディ宇宙センターへの着陸となった。

代替着陸施設は多数あるが、エドワーズ空軍基地とホワイト・サンズ空軍基地以外は使用されることはなかった。

エドワーズ空軍基地についても、シャトルの大陸横断に掛かる多額のコストのため近年はできるだけ利用しない方針が採られており、日本人最後の乗務となったSTS-131の着陸時にも一時は使用が決定していたが、最終的にはケネディ宇宙センターへの着陸となった。

There are many alternative landing facilities, but they were not used except the Edwards Air Force Base and the White Suns Air Force Base. For the Edwards Air Force Base, the policy has been adopted in recent years as far as possible due to the large cost of the shuttle's transit, and the STS-131 was also determined to be used for a time when the Japanese last crew landed, but eventually landed in the Kennedy Space Center.

There are many alternative landing facilities, but they were not used except the Edwards Air Force Base and the White Suns Air Force Base.

For the Edwards Air Force Base, the policy has been adopted in recent years as far as possible due to the large cost of the shuttle's transit, and the STS-131 was also determined to be used for a time when the Japanese last crew landed, but eventually landed in the Kennedy Space Center.

Flight records The main flight records of the shuttle are as follows:

Landing in the Lake Rogers.

Question: when was the enterprise space shuttle first used

Flight records The main flight records of the shuttle are as follows:

Landing in the Lake Rogers.

when was the enterprise space shuttle first used

Question: when did the enterprise space shuttle take off

1977 年 2 月 18 日 エンタープライズ シャトル 輸送 機 に 搭載 さ れ て の 初 飛行 1977 年 8 月 12

Landing in the Lake Rogers.

when did the enterprise space shuttle take off

Question: where did the enterprise space shuttle land in 1977

Landing in the Lake Rogers.

where did the enterprise space shuttle land in 1977

Question: in which us state was the enterprise space shuttle first tested in 1977

Landing in the Lake Rogers.

in which us state was the enterprise space shuttle first tested in 1977

Question: where is the dryden flight research center

カリフォルニア 州 エドワーズ 空軍 基地 の ドライデン 飛行 研究 センター に 向け て 初 飛行 する エンタープライズ 号 ( OV - 101 ) + スペース シャトル 主 飛行 記録 日時 軌道 船 主な できごと / 注記 1977 年 2 月 18 日 エンタープライズ シャトル 輸送 機 に 搭載 さ れ て の 初 飛行 1977 年 8 月 12 日 エンタープライズ

Landing in the Lake Rogers.

where is the dryden flight research center

Flight records The main flight records of the shuttle are as follows :

In 1977 , as part of the Shuttle Plan , the first Enterprise ( OV-101 ) + Space Shuttle Flight Record toward the Dryden Flight Research Center at the Edwards Air Force Base in California in the entry landing test , the first Flight to the Enterprise Shuttle Carrier on February 18 , 1977 , the first Flight to the Enterprise Shuttle Carrier on August 12 , 1977 , the first Single Slayer Flight .

飛行記録 主なシャトルの飛行記録は以下のとおりである。 1977年、シャトル計画の一環である進入着陸試験で、カリフォルニア州エドワーズ空軍基地のドライデン飛行研究センターに向けて初飛行するエンタープライズ号 (OV-101) + スペースシャトル主飛行記録 日時 軌道船 主なできごと/注記 1977年2月18日 エンタープライズ シャトル輸送機に搭載されての初飛行 1977年8月12日 エンタープライズ 初の単独滑空飛行。 ロジャース乾湖に着陸。

飛行記録 主なシャトルの飛行記録は以下のとおりである。

1977年、シャトル計画の一環である進入着陸試験で、カリフォルニア州エドワーズ空軍基地のドライデン飛行研究センターに向けて初飛行するエンタープライズ号 (OV-101) + スペースシャトル主飛行記録 日時 軌道船 主なできごと/注記 1977年2月18日 エンタープライズ シャトル輸送機に搭載されての初飛行 1977年8月12日 エンタープライズ 初の単独滑空飛行。

Flight records The main flight records of the shuttle are as follows: In 1977, as part of the Shuttle Plan, the first Enterprise (OV-101) + Space Shuttle Flight Record toward the Dryden Flight Research Center at the Edwards Air Force Base in California in the entry landing test, the first Flight to the Enterprise Shuttle Carrier on February 18, 1977, the first Flight to the Enterprise Shuttle Carrier on August 12, 1977, the first Single Slayer Flight. Landing in the Lake Rogers.

Flight records The main flight records of the shuttle are as follows:

All seven pilots were killed.

Since the late 1960s, Pads A and B at Kennedy Space Center's Launch Complex 39 have served as backdrops for America's most significant space flight endeavors.
Since the late 1960s, Pads A and B at Kennedy Space Center's Launch Complex 39 have served as backdrops for America's most significant space flight endeavors.
Learn more about shuttle operations at Kennedy Space Center.
Detailed information about past space shuttle missions.
Learn more about each of the space shuttles.
Photos document the shuttles' transition from spaceflight to public display.
This interactive retrospective covers the legacy of the Space Shuttle Program.
Get a 360 view of Kennedy Space Center.› Google Street View of Kennedy →
Participate in a variety of interactive experiences to learn about the accomplishments of the space shuttle program and how it affects your daily life.› Experience site →
Kennedy Space Center: America's space gateway.› View This Video

Question: how many pilots were killed on the challenger

All seven pilots were killed.

how many pilots were killed on the challenger

Question: sts-31 was the name of which space shuttle

All seven pilots were killed.

sts-31 was the name of which space shuttle

Question: what was the longest space stay record for a space shuttle

17 日間 と 15 時間

All seven pilots were killed.

what was the longest space stay record for a space shuttle

Question: when did the challenger space shuttle crash

All seven pilots were killed.

when did the challenger space shuttle crash

Question: what was the name of the space shuttle that crashed into the atlantis in 2003

宇宙 望遠 鏡 発射 ( STS - 31 ) 1992 年 5 月 7 日 エンデバー エンデバー 初 飛行 ( STS - 49 ) 1996 年 11 月 19 日 コロンビア 17 日間 と 15 時間 に わたる シャトル の 最長 宇宙 滞在 記録 ( STS - 80 ) 2000 年 10 月 11 日 ディスカバリー シャトル 100 回 目 の 飛行 ( STS - 92 ) 2003 年 2 月 1日 コロンビア 大気 圏 再

All seven pilots were killed.

what was the name of the space shuttle that crashed into the atlantis in 2003

Question: when did the discovery space shuttle make its 100th flight

All seven pilots were killed.

when did the discovery space shuttle make its 100th flight

Question: how long did the space shuttle columbia stay in space

17 日間 と 15 時間

All seven pilots were killed.

how long did the space shuttle columbia stay in space

29 September 1988 Discovery Challenger Accident First Repeat Flight ( STS-26 ) 4 May 1989 Atlantis Shuttle First Space Shuttle ( Mazelan , STS-30 ) 24 April 1990 Discovery Hubble Space Telescope ( STS-31 ) 7 May 1992 Endover Endover First Flight ( STS-49 ) 19 November 1996 Columbia 17 days and 15 hours Shuttle 's Longest Space Stay Record ( STS-80 ) 11 October 2000 Discovery Shuttle 's 100th Flight ( STS-92 ) 1 February 2003 Coronavirus Re - Atlantic Space Shuttle ( STS-107 )

7名の飛行士全員が死亡。 1988年9月29日 ディスカバリー チャレンジャー号事故後の初の再開飛行 (STS-26) 1989年5月4日 アトランティス シャトルを使用しての初の探査機発射（マゼラン、STS-30） 1990年4月24日 ディスカバリー ハッブル宇宙望遠鏡発射 (STS-31) 1992年5月7日 エンデバー エンデバー初飛行 (STS-49) 1996年11月19日 コロンビア 17日間と15時間にわたるシャトルの最長宇宙滞在記録 (STS-80) 2000年10月11日 ディスカバリー シャトル100回目の飛行 (STS-92) 2003年2月1日 コロンビア 大気圏再突入時に空中分解 (STS-107)。

7名の飛行士全員が死亡。

1988年9月29日 ディスカバリー チャレンジャー号事故後の初の再開飛行 (STS-26) 1989年5月4日 アトランティス シャトルを使用しての初の探査機発射（マゼラン、STS-30） 1990年4月24日 ディスカバリー ハッブル宇宙望遠鏡発射 (STS-31) 1992年5月7日 エンデバー エンデバー初飛行 (STS-49) 1996年11月19日 コロンビア 17日間と15時間にわたるシャトルの最長宇宙滞在記録 (STS-80) 2000年10月11日 ディスカバリー シャトル100回目の飛行 (STS-92) 2003年2月1日 コロンビア 大気圏再突入時に空中分解 (STS-107)。

All seven pilots were killed. 29 September 1988 Discovery Challenger Accident First Repeat Flight (STS-26) 4 May 1989 Atlantis Shuttle First Space Shuttle (Mazelan, STS-30) 24 April 1990 Discovery Hubble Space Telescope (STS-31) 7 May 1992 Endover Endover First Flight (STS-49) 19 November 1996 Columbia 17 days and 15 hours Shuttle's Longest Space Stay Record (STS-80) 11 October 2000 Discovery Shuttle's 100th Flight (STS-92) 1 February 2003 Coronavirus Re-Atlantic Space Shuttle (STS-107)

Landing in the Lake Rogers.

飛行記録 主なシャトルの飛行記録は以下のとおりである。

All seven pilots were killed.

7名の飛行士全員が死亡。

Landing in the Lake Rogers.

飛行記録 主なシャトルの飛行記録は以下のとおりである。

The first flight to remove the back protective cover.

Landing in the Lake Rogers.

October 26, 1977 Enterprise Enterprise's last skiing test.

Question: when was enterprise enterprise's last skiing test

The first flight to remove the back protective cover.

Landing in the Lake Rogers.

October 26, 1977 Enterprise Enterprise's last skiing test.

when was enterprise enterprise's last skiing test

Question: where did the first flight of the enterprise enterprise land

Landing in the Lake Rogers.

where did the first flight of the enterprise enterprise land

Question: when was the last time the enterprise enterprise skied

Landing in the Lake Rogers.

when was the last time the enterprise enterprise skied

Question: what was the first flight to remove

尾部 保護 カバー を 取り 除い

Landing in the Lake Rogers.

Question: where did the first flight of enterprise enterprise land

Landing in the Lake Rogers.

where did the first flight of enterprise enterprise land

Question: what was the name of the first flight to remove the back protective cover of a ski

エンタープライズ エンタープライズ

Landing in the Lake Rogers.

what was the name of the first flight to remove the back protective cover of a ski

Question: what was enterprise enterprise's last flight

Landing in the Lake Rogers.

what was enterprise enterprise's last flight

The first flight to remove the back protective cover .

October 26 , 1977 Enterprise Enterprise 's last skiing test .

尾部保護カバーを取り除いての初飛行。 ロジャース乾湖に着陸。 1977年10月26日 エンタープライズ エンタープライズ最後の滑空試験。

尾部保護カバーを取り除いての初飛行。
ロジャース乾湖に着陸。
1977年10月26日 エンタープライズ エンタープライズ最後の滑空試験。

The first flight to remove the back protective cover. Landing in the Lake Rogers. October 26, 1977 Enterprise Enterprise's last skiing test.

The first flight to remove the back protective cover.

October 26, 1977 Enterprise Enterprise's last skiing test.

Landing in the Lake Rogers.

飛行記録 主なシャトルの飛行記録は以下のとおりである。

All seven pilots were killed.

7名の飛行士全員が死亡。

Landing in the Lake Rogers.

飛行記録 主なシャトルの飛行記録は以下のとおりである。

On February 1, 2003, the Space Shuttle Columbia was destroyed in the air when the main wing front side was destroyed by the fortified carbon-carbon-cutting heat material.

スペースシャトルコロンビア

Space Pioneers
Dr. Robert H. Goddard
Dr. Robert H. Goddard was an American physicist and inventor who designed and built the first liquid fueled rocket. Goddard filed for over two hundred patents. Including a multi-staged rocket along with a patent for liquid fueled rockets.
He was the first to theorize and design rockets for space travel, orbital research and intercontinental ballistic missiles. He has been called the man who started the space age. Much of his research he kept private due to early ridicule by the media.
Related Article – 5 Best Low Time Pilot Jobs With 250 Hours
Goddard didn’t find financial support from the government between the World Wars. He was funded by the Guggenheim family from 1930 to 1934. During World War II he worked for the Navy in Annapolis, Maryland.
Goddard’s first successful launch of a liquid fueled rocket was on March 16, 1926. He went on to launch 35 rockets during his career from 1926 to 1941.
Goddard developed many innovations during his rocket experiments. Germany was inspired by the news of his work and incorporated his ideas into their V-2 rocket program during World War II.
Goddard was born October 5, 1882 and died on August 10, 1945. NASA named the Goddard Space Flight Center in Maryland in his honor in 1959. There is a memorial and statue of Goddard at the location he launched the first liquid fueled rocket in Auburn, Massachusetts.
Mercury 3 Freedom 7 Mission
On May 5, 1961 Alan Shepard piloted Freedom 7 to become the first American in space. The mission known as Mercury-Redstone 3 whose goal was to launch an astronaut into orbit and return him safely. Freedom 7 was built by McDonnell Aircraft Corporation in St. Louis, Missouri.
The Soviet Union had launched the first man in space, Yuri Gagarin, three weeks earlier. Beating the Americans at getting the first man in space.
The Mercury-Redstone 3 rocket carried Freedom 7 to orbit from Cape Canaveral, Florida. An estimated 45 million Americans watched the launch at 9:34 Eastern Time. The launch was successful and Alan tested the spacecrafts control systems and retrorockets. The mission was short only covering a distance of 300 miles and attaining an altitude of 116 miles. The duration of the space flight was 15 minutes, 22 seconds.
Related Article – Instrument Proficiency Check (IPC): 4 Things You Need To Know
The capsule re-entered the atmosphere, deployed a parachute and landed near the bahamas. The pilot and capsule were recovered by helicopter and returned them to the aircraft carrier USS Lake Champlain. The recovery took just 11 minutes from splashdown to arriving on the carrier deck.
Freedom 7 is on display in Boston at the John F. Kennedy Library.
Gemeni Space Program
NASA’s Gemini Space Program spanned between the Mercury and Apollo projects, between 1961 and 1966. Gemini missions had two Astronaut crews. Gemini’s goal was to develop and perfect techniques needed for future Apollo missions to the Moon. These tasks included mission durations matching traveling to the moon and returning to Earth, space walks, rendezvous and docking with a second craft.
Gemini missions launched from Cape Canaveral, Florida. The rocket that carried Gemini craft into orbit was a modified Gemini-Titan II intercontinental ballistic missile. NASA used the recently completed Mission Control Center in Houston for the Gemini Space Program. NASA’s official pronunciation of the project as Jeh-mih-knee.
The Gemini capsules were built by McDonnell Aircraft in St. Louis, Missouri. The capsules were 18 feet long and 10 feet wide. There were twelve operational capsules built and used in space flight and 17 training capsules constructed that were not used in space flight. These are all on display in various museums and space centers around the country.
The first two Gemini missions were launched without crews for testing. These missions occurred in 1964 and 1965. There were ten crewed Gemini missions from 1965 to 1966. During Gemini 4 in 1965 Ed White made the first space walk. Gemini 5 lasted 8 days in orbit to test the effect on the crew during the time needed to get to the moon and back. Gemini 8 completed the first space docking with another spacecraft, the Agena target vehicle.
The Gemini Space program developed many advances that were used during the Apollo program to get to the Moon.
Apollo 12 Mission
Apollo 12 was launched on November 14, 1969 from Cape Canaveral, Florida. It was crewed by Commander Charles Conrad, Jr., Command Module Pilot Richard Gordon, and Lunar Module Pilot Alan Bean. Apollo 12 was the second mission to land on the Moon just 4 months after Apollo 11.
The Apollo 12 lunar module Intrepid made a successful descent and landing to the Moon on November 19, 1969. Conrad and Gordon landed on the Moon while Bean manned the lunar module. The target landing site was near the location of the Surveyor 3 probe. Surveyor 3 had touched down on the moon in 1967. The targeted landing was a success. The Astronauts were within walking distance of the probe. Conrad and Bean removed parts of the probe to be brought back to earth for scientific analysis.
Apollo 12 brought an improved color video camera to the Moon. An improvement over the monochrome video camera used in Apollo 11. When Bean was setting up the camera he accidentally pointed it directly at the Sun. This destroyed the internal imaging components of the camera ending television coverage of the mission.
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The astronauts also collected Moon rocks and setup equipment for measuring seismic activity, magnetic fields and solar winds on the Moon. After returning from the lunar surface and docking with the command module the ascent stage was jettisoned and crashed into the Moon. This was done to test the seismic sensors the astronauts had placed on the moon. The sensors detected the impact and vibrations continued for over an hour.
On their return trip from the Moon to Earth, the crew were able to view and video the Earth eclipsing the sun. They splashed down in the Pacific Ocean on November 24, 1969 concluding a successful mission.
Space Shuttle History
The US Space Shuttle concept began with the release of a joint study by the Air Force and NASA in 1966. NASA reviewed 29 designs for a reusable shuttle system. In January 1972 the shuttle program was approved and the final design was decided in March. Construction of the first shuttle began in 1974. This shuttle named Enterprise was for flight testing and landing and was not built to be launched into space.
The shuttle system consists of a reusable orbiter, the shuttle, an expendable large external fuel tank, and reusable solid fuel rocket boosters. The rocket boosters were recovered after each launch and refurbished and reused for multiple missions. The large external fuel tanks broke up on re-entry with debris landing in the ocean and were not recovered.
The shuttle system is officially known as the Space Transportation System (STS). Five operational shuttles were built. They were the Columbia, Challenger, Discovery, Atlantis and Endeavour. They flew 135 missions with the first in 1981 and the last in 2011. They were used to launch satellites, space probes, the Hubble Space Telescope and take supplies and Astronauts to the International Space Station.
The Shuttle Challenger was lost during launch on January 28, 1986 when seal failed in an O-ring in one of the solid rocket boosters. It allowed hot gas to escape burn through the large external fuel tank leading to a catastrophic failure killing all seven astronauts on board.
The Shuttle Columbia was destroyed during re-entry on February 1, 2003. During launch a piece of insulating foam shed from the external fuel tank struck the leading edge of the wing damaging it. During re-entry this allowed hot gases to enter the wing causing the shuttle to disintegrate on reentry over Texas. All seven astronauts on board were killed.
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Challenger Space Shuttle Accident
The Space Shuttle Challenger disintegrated just after launch on January 28, 1986 from Cape Canaveral, Florida. The launch was viewed by most school aged children in the US since the shuttle was to carry the first teacher in space. Christa McAuliffe was selected to be the first teacher in space from a contest with over 11,000 applicants. She was to be the first teacher in space perform experiments and teach two lessons from space.
The accident was caused when an O-ring failed in one of the joints in the right solid rocket booster. Hot gas escaped from the failed seal and burned through attachment hardware causing a structural failure and the disintegration of the orbiter and the death of all seven crew members.
The forecasts for launch day predicted temperatures below freezing close to 30 degrees fahrenheit with a previous overnight low temp down to 18 degrees fahrenheit . The shuttle was not certified to operate at temperatures below 40 degrees fahrenheit. As early as 1985 engineers at Thiokol who designed and built the solid rocket boosters raised concerns with using the boosters during cold weather and potential O-ring failure.
Nasa held a call with NASA and Thiokol the night before the launch. Thiokol engineers pushed for delaying the launch as they had no test information for the O-rings at that low temperature. During a second call with NASA and Thiokol there were discussions that the secondary O-rings could be depended on if the primary ones failed and it was decided to continue with the ill fated launch.
After the NASA call the night before launch Bob Ebeling told his wife about the O-rings and that the shuttle would blow up the next day during launch.
Columbia Space Shuttle Accident
The shuttle Columbia broke up on re-entry on February 1, 2003. The 7 crew members died during the accident. The crew included Commander Rick Husband, Pilot William McCool, Payload Commander Michael Anderson, Payload Specialist IIan Ramon, Mission Specialist Kalpana Chawla, Mission Specialist David Brown and Mission Specialist Laurel Blair Salton Clark.
During launch a piece of foam approximately 2’ x 2’ fell from the large external fuel tank striking the wing of the Columbia. This was determined by analyzing video footage the day after the launch. Previous shuttles had an issue with foam shedding and it wasn’t considered to be an issue that would affect the mission or re-entry.
This time the foam strike made a 6” to 10” hole in the wing. This would allow hot gases to enter the wing during re-entry. The re-entry was treated as a normal re-entry by NASA they weren’t concerned about the wing or any other issues.
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During re-entry amateur observers on the ground in California noticed debris being shed by the Shuttle and that it looked different than other re-entries they had watched. Shortly there after the Columbia became brighter and a glowing streak followed it. Over texas the Columbia disintegrated. The debris trail was picked up by the regional weather radar as a red and yellow band extending from texas into Lousiana. The crew is believed to have perished in less than 60 seconds.
Exploring Other Bodies
Scientists focus of study was initially on planets and moons. As they completed their exploration of these bodies with space probes they began to expand their focus to smaller space bodies known as planetesimals. These are smaller space bodies such as asteroids and comets.World scientists are devoting more study to these objects as they offer a glimpse into the formation of the solar system.
NASA’s Galileo space probe was launched in 1989. It’s primary mission was to study Jupiter and its moons. A secondary mission was to study two asteroids. The Galileo probe was the first spacecraft to visit an asteroid. It actually visited two asteroids during its mission Gaspra and Ida.
Halley’s comet gave scientists a great opportunity for study. The Euorpean Space Agency launched the Giotto probe to study the comet. It was the first probe to make a close observation of a comet. The probe came within 600 kilometers of the center of Halley’s comet.
Images collected by Giotto showed Halley’s comet to be a dark peanut shaped object approximately 15 kilometers long. Mass spectrometer analysis suggested it had a carbon rich surface.
The world space organizations are planning more visits to comets and meteors. There is also interest in mining asteroids. There are estimates that asteroids can each contain trillions of dollars of valuable metals.

Question: when was the space shuttle columbia destroyed in the air

On February 1, 2003, the Space Shuttle Columbia was destroyed in the air when the main wing front side was destroyed by the fortified carbon-carbon-cutting heat material.

when was the space shuttle columbia destroyed in the air

Question: where did the space shuttle columbia crash

where did the space shuttle columbia crash

Question: what damaged the main wing of the space shuttle columbia

カーボン ・ カーボン 断熱 材

what damaged the main wing of the space shuttle columbia

Question: what was the name of the space shuttle destroyed in the air by a carbon-cutting

what was the name of the space shuttle destroyed in the air by a carbon-cutting

Question: how many high-resolution photographs did nasa ask the ministry of defense to take

how many high-resolution photographs did nasa ask the ministry of defense to take

Question: what did nasa's thermal protection system technical director ask for permission to investigate

熱 保護 システム の 技術 主任 は コロンビア に 搭乗 し て いる 飛行 士 たち に 耐熱 タイル

what did nasa's thermal protection system technical director ask for permission to investigate

Question: when was the space shuttle columbia destroyed

when was the space shuttle columbia destroyed

Question: the space shuttle columbia was destroyed by a carbon-cutting heat-resistant tyre

the space shuttle columbia was destroyed by a carbon-cutting heat-resistant tyre

On February 1 , 2003 , the Space Shuttle Columbia was destroyed in the air when the main wing front side was destroyed by the fortified carbon - carbon - cutting heat material .

The technicians in the ground control room asked the Ministry of Defense to take three high - resolution photographs in order to understand the extent of damage more clearly , and NASA 's thermal protection system technical director asked the pilots on board in Colombia for permission for off - boat activities to investigate the damage of heat - resistant tyres .

2003年2月1日、スペースシャトルコロンビア号が発射の際に主翼前縁の強化カーボン・カーボン断熱材が損傷したことにより、大気圏再突入時に空中分解した。 地上管制室の技術者たちは損傷の広がりをより明確に把握できるよう、国防総省に対して三回にわたって高解像度の写真を撮影するよう要求し、NASAの熱保護システムの技術主任はコロンビアに搭乗している飛行士たちに耐熱タイルのダメージを調査させるべく船外活動の許可を求めた。

2003年2月1日、スペースシャトルコロンビア号が発射の際に主翼前縁の強化カーボン・カーボン断熱材が損傷したことにより、大気圏再突入時に空中分解した。

地上管制室の技術者たちは損傷の広がりをより明確に把握できるよう、国防総省に対して三回にわたって高解像度の写真を撮影するよう要求し、NASAの熱保護システムの技術主任はコロンビアに搭乗している飛行士たちに耐熱タイルのダメージを調査させるべく船外活動の許可を求めた。

On February 1, 2003, the Space Shuttle Columbia was destroyed in the air when the main wing front side was destroyed by the fortified carbon-carbon-cutting heat material. The technicians in the ground control room asked the Ministry of Defense to take three high-resolution photographs in order to understand the extent of damage more clearly, and NASA's thermal protection system technical director asked the pilots on board in Colombia for permission for off-boat activities to investigate the damage of heat-resistant tyres.

On February 1, 2003, the Space Shuttle Columbia was destroyed in the air when the main wing front side was destroyed by the fortified carbon-carbon-cutting heat material.

In order to fill the gap in the space development plan by the Shuttle withdrawal, a spacecraft is currently being developed that will not only carry pilots and equipment to the ISS, but will be able to reach the moon and Mars away from Earth.

The design concept, which was initially called “Crew Exploration Vehicle”, was then developed into the Orion Space Shuttle and the Constellation Plan.

Question: what was the design concept of the orion space shuttle originally called

In order to fill the gap in the space development plan by the Shuttle withdrawal, a spacecraft is currently being developed that will not only carry pilots and equipment to the ISS, but will be able to reach the moon and Mars away from Earth.

The design concept, which was initially called “Crew Exploration Vehicle”, was then developed into the Orion Space Shuttle and the Constellation Plan.

what was the design concept of the orion space shuttle originally called

Question: a spacecraft is being developed that will be able to reach where away from earth

a spacecraft is being developed that will be able to reach where away from earth

Question: a spacecraft is being developed to reach the moon and what other planet away from earth

a spacecraft is being developed to reach the moon and what other planet away from earth

Question: what is the name of the space shuttle

Question: what is needed to fill the gap in the space development plan

what is needed to fill the gap in the space development plan

Question: what is the goal of the Orion space shuttle

月 や 火星 まで 到達

what is the goal of the Orion space shuttle

In order to fill the gap in the space development plan by the Shuttle withdrawal , a spacecraft is currently being developed that will not only carry pilots and equipment to the ISS , but will be able to reach the moon and Mars away from Earth .

The design concept , which was initially called “ Crew Exploration Vehicle ” , was then developed into the Orion Space Shuttle and the Constellation Plan .

シャトル退役による宇宙開発計画の間隙を埋めるべく、飛行士や搭載物をISSに運ぶだけでなく、地球を離れて月や火星まで到達できるような宇宙船が現在開発中である。 当初「有人開発船（Crew Exploration Vehicle）」と呼ばれていた計画概念は、その後オリオン宇宙船やコンステレーション計画へと発展した。

シャトル退役による宇宙開発計画の間隙を埋めるべく、飛行士や搭載物をISSに運ぶだけでなく、地球を離れて月や火星まで到達できるような宇宙船が現在開発中である。

当初「有人開発船（Crew Exploration Vehicle）」と呼ばれていた計画概念は、その後オリオン宇宙船やコンステレーション計画へと発展した。

In order to fill the gap in the space development plan by the Shuttle withdrawal, a spacecraft is currently being developed that will not only carry pilots and equipment to the ISS, but will be able to reach the moon and Mars away from Earth. The design concept, which was initially called “Crew Exploration Vehicle”, was then developed into the Orion Space Shuttle and the Constellation Plan.

In order to fill the gap in the space development plan by the Shuttle withdrawal, a spacecraft is currently being developed that will not only carry pilots and equipment to the ISS, but will be able to reach the moon and Mars away from Earth.

The design concept, which was initially called “Crew Exploration Vehicle”, was then developed into the Orion Space Shuttle and the Constellation Plan.

However, in 2010, the Obama administration cut off the budget of the constellation plan, and in the future, it proposed to assign the operations of satellite launch to low orbits to private companies.

Until the next-generation spacecraft appears, there will be no means other than relying on the Russian Soyuz spacecraft or on the U.S. private commercial spacecraft in development to return to the ISS.

Question: when did the obama administration cut off the budget of the constellation plan

However, in 2010, the Obama administration cut off the budget of the constellation plan, and in the future, it proposed to assign the operations of satellite launch to low orbits to private companies.

Until the next-generation spacecraft appears, there will be no means other than relying on the Russian Soyuz spacecraft or on the U.S. private commercial spacecraft in development to return to the ISS.

when did the obama administration cut off the budget of the constellation plan

Question: what was the design concept for the orion space shuttle originally called

what was the design concept for the orion space shuttle originally called

Question: who is slated to launch satellites to low orbits

who is slated to launch satellites to low orbits

Question: what was the design concept for the crew exploration vehicle called

what was the design concept for the crew exploration vehicle called

Question: what was the design concept developed into

Orion Space Shuttle and the Constellation Plan

オリオン 宇宙 船 や コンステレーション 計画

Orion Space Shuttle and the Constellation Plan

what was the design concept developed into

Question: who cut off the budget of the constellation plan

who cut off the budget of the constellation plan

Question: until the next generation of spacecraft appears, what is the only way to return to the

ロシア 連邦 の ソユーズ 宇宙 船

until the next generation of spacecraft appears, what is the only way to return to the

However , in 2010 , the Obama administration cut off the budget of the constellation plan , and in the future , it proposed to assign the operations of satellite launch to low orbits to private companies .

Until the next - generation spacecraft appears , there will be no means other than relying on the Russian Soyuz spacecraft or on the U.S. private commercial spacecraft in development to return to the ISS .

当初「有人開発船（Crew Exploration Vehicle）」と呼ばれていた計画概念は、その後オリオン宇宙船やコンステレーション計画へと発展した。 しかし2010年にオバマ政権はコンステレーション計画の予算を打ち切り、今後は低軌道への衛星発射の事業は民間企業に委託することを提案した。 次世代の宇宙船が登場するまでは、飛行士がISSに到達しまた帰還するためにはロシア連邦のソユーズ宇宙船か、または開発中のアメリカの民間商用宇宙船に頼る以外に手段がなくなる。

しかし2010年にオバマ政権はコンステレーション計画の予算を打ち切り、今後は低軌道への衛星発射の事業は民間企業に委託することを提案した。

次世代の宇宙船が登場するまでは、飛行士がISSに到達しまた帰還するためにはロシア連邦のソユーズ宇宙船か、または開発中のアメリカの民間商用宇宙船に頼る以外に手段がなくなる。

The design concept, which was initially called “Crew Exploration Vehicle”, was then developed into the Orion Space Shuttle and the Constellation Plan. However, in 2010, the Obama administration cut off the budget of the constellation plan, and in the future, it proposed to assign the operations of satellite launch to low orbits to private companies. Until the next-generation spacecraft appears, there will be no means other than relying on the Russian Soyuz spacecraft or on the U.S. private commercial spacecraft in development to return to the ISS.

However, in 2010, the Obama administration cut off the budget of the constellation plan, and in the future, it proposed to assign the operations of satellite launch to low orbits to private companies.

Until the next-generation spacecraft appears, there will be no means other than relying on the Russian Soyuz spacecraft or on the U.S. private commercial spacecraft in development to return to the ISS.

On December 23, 2008, NASA announced that it has entered into a contract with SpaceX and Obital Sciences on the “NASA Awards Space Station Commercial Resupply Services Contracts” (NASA Awards Space Station Commercial Resupply Services Contracts).

Space X launched the Dragon spacecraft with Falcon 9 rocket in 2012 and Obital Sciences launched the Signas spacecraft with an Antarctic rocket in 2013 and performed an unmanned supply mission to the ISS instead of the Space Shuttle.

CONTRACT RELEASE
:
C08-069
NASA Awards Space Station Commercial Resupply Services Contracts
WASHINGTON -- NASA has awarded two contracts -- one to Orbital Sciences Corp. of Dulles, Va., and one to Space Exploration Technologies (SpaceX) of Hawthorne, Calif. -- for commercial cargo resupply services to the International Space Station. At the time of award, NASA has ordered eight flights valued at about $1.9 billion from Orbital and 12 flights valued at about $1.6 billion from SpaceX.
These fixed-price indefinite delivery, indefinite quantity contracts will begin Jan. 1, 2009, and are effective through Dec. 31, 2016. The contracts each call for the delivery of a minimum of 20 metric tons of upmass cargo to the space station. The contracts also call for delivery of non-standard services in support of the cargo resupply, including analysis and special tasks as the government determines are necessary.
NASA has set production milestones and reviews on the contracts to monitor progress toward providing services. The maximum potential value of each contract is about $3.1 billion. Based on known requirements, the value of both contracts combined is projected at $3.5 billion.
These agreements will fulfill NASA's need to procure cargo delivery services to the space station using a U.S. commercial carrier after the retirement of the space shuttle.
For more information about the space station, visit:
http://www.nasa.gov/station
- end -
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Question: when did nasa announce a contract with SpaceX and obital sciences

23 日 、 NASA は ISS へ の 物資 補給 を 民間 に 委ねる 商業 軌道 輸送 サービス ( COTS ) に 関する 契約 を 、 スペース X 社 および オービタル ・ サイエンシズ 社 と 取り 交わし た こと を 発表 し た "" NASA Awards Space Station Commercial Resupply Services Contracts "" . NASA , December 23 , 2008

On December 23, 2008, NASA announced that it has entered into a contract with SpaceX and Obital Sciences on the “NASA Awards Space Station Commercial Resupply Services Contracts” (NASA Awards Space Station Commercial Resupply Services Contracts).

Space X launched the Dragon spacecraft with Falcon 9 rocket in 2012 and Obital Sciences launched the Signas spacecraft with an Antarctic rocket in 2013 and performed an unmanned supply mission to the ISS instead of the Space Shuttle.

when did nasa announce a contract with SpaceX and obital sciences

Question: which two companies did nasa sign a contract with in 2008

X 社 および オービタル ・ サイエンシズ

which two companies did nasa sign a contract with in 2008

Question: when did obital sciences launch signas

Question: when did space x launch the dragon spacecraft

when did space x launch the dragon spacecraft

Question: what type of rocket was used to launch the dragon spacecraft

ファルコン 9 ロケット で ドラゴン 宇宙 船 を 、 オービタル ・ サイエンシズ は 2013 年 に アンタレス ロケット で シグナス

what type of rocket was used to launch the dragon spacecraft

Question: what is the name of the space station spacecraft launched by space x in 2012

what is the name of the space station spacecraft launched by space x in 2012

Question: in what year did space station commercial resupply services contracts begin

in what year did space station commercial resupply services contracts begin

Question: which company launched the dragon spacecraft in 2012

which company launched the dragon spacecraft in 2012

On December 23 , 2008 , NASA announced that it has entered into a contract with SpaceX and Obital Sciences on the “ NASA Awards Space Station Commercial Resupply Services Contracts ” ( NASA Awards Space Station Commercial Resupply Services Contracts ) .

Space X launched the Dragon spacecraft with Falcon 9 rocket in 2012 and Obital Sciences launched the Signas spacecraft with an Antarctic rocket in 2013 and performed an unmanned supply mission to the ISS instead of the Space Shuttle .

民間商用宇宙船への交代 2008年12月23日、NASAはISSへの物資補給を民間に委ねる商業軌道輸送サービス (COTS) に関する契約を、スペースX社およびオービタル・サイエンシズ社と取り交わしたことを発表した"NASA Awards Space Station Commercial Resupply Services Contracts". NASA, December 23, 2008.。 スペースXは2012年にファルコン9ロケットでドラゴン宇宙船を、オービタル・サイエンシズは2013年にアンタレスロケットでシグナス宇宙船を打ち上げ、スペースシャトルに代わってISSへの無人補給ミッションを果たした。

民間商用宇宙船への交代 2008年12月23日、NASAはISSへの物資補給を民間に委ねる商業軌道輸送サービス (COTS) に関する契約を、スペースX社およびオービタル・サイエンシズ社と取り交わしたことを発表した"NASA Awards Space Station Commercial Resupply Services Contracts". NASA, December 23, 2008.。

スペースXは2012年にファルコン9ロケットでドラゴン宇宙船を、オービタル・サイエンシズは2013年にアンタレスロケットでシグナス宇宙船を打ち上げ、スペースシャトルに代わってISSへの無人補給ミッションを果たした。

On December 23, 2008, NASA announced that it has entered into a contract with SpaceX and Obital Sciences on the “NASA Awards Space Station Commercial Resupply Services Contracts” (NASA Awards Space Station Commercial Resupply Services Contracts). Space X launched the Dragon spacecraft with Falcon 9 rocket in 2012 and Obital Sciences launched the Signas spacecraft with an Antarctic rocket in 2013 and performed an unmanned supply mission to the ISS instead of the Space Shuttle.

On December 23, 2008, NASA announced that it has entered into a contract with SpaceX and Obital Sciences on the “NASA Awards Space Station Commercial Resupply Services Contracts” (NASA Awards Space Station Commercial Resupply Services Contracts).

Space X launched the Dragon spacecraft with Falcon 9 rocket in 2012 and Obital Sciences launched the Signas spacecraft with an Antarctic rocket in 2013 and performed an unmanned supply mission to the ISS instead of the Space Shuttle.

NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing.

STA was also used in the Kennedy Space Center in Florida.

Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing
How do you help astronauts feel like they've landed the Space Shuttle hundreds of times without actually using a Shuttle?
That is one task given to research pilots at Johnson Space Center in Houston. They take the sleek Shuttle Training Aircraft (STA) and use it to train astronauts to land the massive-by-comparison Space Shuttle.
Image to right: The Shuttle Training Aircraft is used to train astronauts to land the Space Shuttle. Credit: NASA
"Astronauts only get one try to land the Shuttle; it lands as a glider," said Research Pilot Triple Nickel. "We use the very slick, sports car-like feel of the STA to simulate the 'falling brick' of the Space Shuttle."
According to astronauts who have flown both, the STA gives a remarkably good simulation of a landing Shuttle.
"The STA is great! It flies very closely to the actual orbiter," said STS-114 Commander Eileen Collins. "The handling qualities are close enough that we have no problems transitioning from training to the real thing."
The STA is a "highly modified Gulfstream-2 aircraft," Nickel said. The plane’s cockpit has been adapted to closely resemble a Shuttle's flight deck; the left-hand seat features the same hand controls that a Shuttle commander uses during a mission. Several passenger seats in the back of the aircraft have been replaced by a bank of Shuttle-simulating computers.
"When a pilot flies the aircraft, he's actually flying a computer model of the Shuttle," Nickel said. NASA has four of these customized aircraft, most of which are kept at the NASA Forward Operating Location in El Paso, Texas. The fleet is rotated through Ellington Field for maintenance as needed.
Most of the STA training takes place at the White Sands Space Harbor, a part of the White Sands Test Facility in New Mexico. The location is ideal because it not only has an actual Shuttle landing strip -- an STA training requirement -- but it provides more remote, undisturbed airspace than its counterparts at Kennedy Space Center and at Edwards Air Force Base in California.
Astronauts are shuttled to El Paso on another NASA training aircraft, the T-38. Each STA training session includes 10 approaches and landings. After a session, the training crew refuels the aircraft, picks up a different astronaut and begins again.
Image to left: The Shuttle Training Aircraft flies over the desert accompanied by a T-38 aircraft. Credit: NASA
"STA training is all about producing a familiar path through the sky for the astronauts," Nickel said. "No matter the height of the astronaut, we put their eyes at a consistent height so when they fly the Shuttle, the view looks familiar. We even mask off the windows to match the view from a Shuttle."
The STA flies at the same speeds, dive angle and approach trajectory as the Shuttle. But simulating that "falling brick" feeling is a delicate maneuver.
"To accomplish that, we lower the main gear of the STA and put the engines in reverse thrust," he said. "You know when a commercial plane lands and you're thrown forward after the wheels touch down? We do that at 30,000 feet."
Prior to a Shuttle mission, a commander has to complete 1,000 STA landings. The STS-114 commander and pilot, Eileen Collins and Jim Kelly, will continue STA training on a weekly basis until launch. Such thorough practice leads to confident astronauts and successful Shuttle landings -- not to mention great feedback for the STA teams.
"Shuttle astronauts come back and tell us it felt like they had done it a thousand times before," Nickel said.
"It is some of the best training we have," Collins said.

Question: where is the nasa shuttle used in florida

ケネディ 宇宙 センター

NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing.

STA was also used in the Kennedy Space Center in Florida.

where is the nasa shuttle used in florida

Question: what was used in the kennedy space center in florida

what was used in the kennedy space center in florida

Question: the kennedy space center is in which us state

the kennedy space center is in which us state

Question: what is the nasa term for a shuttle that preps astronauts for landing

what is the nasa term for a shuttle that preps astronauts for landing

Question: which space agency operates the'shuttle'

which space agency operates the'shuttle'

Question: where was the shuttle STA used in florida

ケネディ 宇宙 センター

where was the shuttle STA used in florida

Question: what does STA stand for

Shuttle Training Aircraft Preps Astronauts for Landing

Shuttle Training Aircraft Preps Astronauts for Landing

Question: where was the space shuttle STA used

フロリダ 州 の ケネディ 宇宙 センター

NASA - Test Drive : Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive : Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive : Shuttle Training Aircraft Preps Astronauts for Landing .

STA was also used in the Kennedy Space Center in Florida .

運用の歴史 4機のSTAが通常はテキサス州エルパソで飛行訓練を行い、ヒューストンで整備を行ったNASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing。 STAは同様にフロリダ州のケネディ宇宙センターでも使用された。

運用の歴史 4機のSTAが通常はテキサス州エルパソで飛行訓練を行い、ヒューストンで整備を行ったNASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing。

STAは同様にフロリダ州のケネディ宇宙センターでも使用された。

NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing. STA was also used in the Kennedy Space Center in Florida.

NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing - NASA - Test Drive: Shuttle Training Aircraft Preps Astronauts for Landing.

STA was also used in the Kennedy Space Center in Florida.

Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle

It was planned to be used in a combination of external fuel tanks (ET) and solid fuel assistant rocket (SRB) and cargo modules with the main engine.

Shuttle-C
Between 1984 and 1995 the government investigated numerous concepts for the development of a replacement to the shuttle. Each of these concepts focused on the vehicle only and did not address the entire flight and ground systems. Major emphasis was placed on the reduction of flight element production costs and retained the support the design approach, starting with the rocket engines, then the vehicle, and then the ground facilities and support equipment required. It was recognized that operations experience was needed to evolve to a more operable approach for the future.
However, due to priorities at the launch operations center, personnel with hands-on experience were not actively involved in these advanced study efforts. In many cases the concept was targeted to replace shuttle; however, there was no definitive plan to bridge the gap from shuttle until the new program was certified as operational. Additionally, the lack of definition of a mission model in many cases led to the premature termination of each effort.
Shuttle-C was envisioned to fill the gap between shuttle and ALS5. Using an unmanned cargo carrier in place of the orbiter, Shuttle-C was compatible with existing shuttle infrastructure and offered a heavy-lift capability (up to 150K lbs. to low earth orbit) within a four year window. In the spring of 1989 a Shuttle-C Users Conference was sponsored at Huntsville, Alabama to enlighten the user community of system capabilities and determine a preliminary mission model. The then Space Station Freedom was targeted at that time as a potential user in the space station assembly sequence and the logistics resupply missions.
Shuttle-C, a partially reusable cargo vehicle, would capitalize on the existing Space Transportation System infrastructure at KSC and elsewhere to keep costs down. With its heavy lift capability -- anywhere from 100,000 to 170,000 pounds (45,360 to 77,112 kilograms) -- Shuttle-C could reduce by 50 percent the number of launches and length of assembly time for space station elements. It also could be used to carry into orbit scientific spacecraft. Among the candidate missions under study were two of the Great Observatories, AXAF and SIRTF. A cargo Shuttle could serve as a test bed for new Shuttle elements like the ASRM.
Several Shuttle-derived vehicle (SDV) designs have been envisioned for near-term NASA and DOD heavy lift launch vehicle (HLLV) missions into the 21st century. Studies performed at the Marshall Center have included SRB-X, side mount and in-line boostered configurations, each an evolutive concept incorporating STS features to deliver 144,000-195,000 lb payloads into LEO or 10,000 lb into GEO. A three-stage multitank design sporting a cluster of eight 1.757 Mlb thrust engines on the first stage is emerging as a favored HLLV. The second stage would be lofted by four 481,000 lb thrust SSME derivative engines and the third stage would have two of the derivatives. All stages would be drogue-parachuted to water touchdown for reuse.
The technology requiring the greatest advances to realize the design is a reusable, long-life liquid oxygen/hydrocarbon fueled booster. The 5-6 yr development span would require associated selections of launch sites and the construction of launch facilities which would not interfere with STS operations.
In October 1987, NASA selected three contractors to perform the first of a two-phase systems definition study for Shuttle-C. The efforts focused on vehicle configuration details, including the cargo element's length and diameter, the number of liquid-fueled main engines, and an operations concept evaluation that included ground and flight support systems. A major purpose of the study was to determine whether Shuttle-C would be cost effective in supporting the space station. Using Shuttle-C could free the Space Shuttle for STS-unique missions, such as solar system exploration, astronomy, life sciences, space station crew rotation, and logistics and materials processing experiments. Shuttle-C also would be used to launch planetary missions and serve as a test bed for new Shuttle boosters.
The results of the Shuttle-C efforts were to be coordinated with other ongoing advanced launch systems studies to enable a joint steering group, composed of DOD and NASA senior managers. The purpose of the steering group was to formulate a national heavy-lift vehicle strategy that best accommodated both near-term requirements and longer term objectives for reducing space transportation operational costs.
Due to sharing the development costs, the space station community was not ready to commit to using the Shuttle-C. The message from other potential users was "build it and we will come"; however, no one was willing to step up and commit to Shuttle-C. With no clear definition of annual utilization the program was ultimately terminated.
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シャトルC シャトルCの夜間打ち上げの想像図 シャトルCはアメリカ航空宇宙局が提案したスペースシャトルの構成要素を流用した無人貨物打ち上げロケットである。 外部燃料タンク (ET) と固体燃料補助ロケット (SRB) とメイン・エンジンを備えた貨物用モジュールを組み合わせて使用される予定だった。

シャトルC シャトルCの夜間打ち上げの想像図 シャトルCはアメリカ航空宇宙局が提案したスペースシャトルの構成要素を流用した無人貨物打ち上げロケットである。

外部燃料タンク (ET) と固体燃料補助ロケット (SRB) とメイン・エンジンを備えた貨物用モジュールを組み合わせて使用される予定だった。

Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle It was planned to be used in a combination of external fuel tanks (ET) and solid fuel assistant rocket (SRB) and cargo modules with the main engine.

Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle C Shuttle

It was planned to be used in a combination of external fuel tanks (ET) and solid fuel assistant rocket (SRB) and cargo modules with the main engine.

The concept of Shuttle C was theoretically expected to reduce the development costs of heavy-match launch missiles by the reuse technology developed in the Shuttle plan.

The proposal was made several times, both of which were to use the existing shuttle structures and the immediate main engines and navigation computers with a limit of use.

[164] With the 1982 completion of the orbital flight test series, NASA began a planned acceleration of the Space Shuttle launch schedule. One early plan contemplated an eventual rate of a mission a week, but realism forced several downward revisions. In 1985, NASA published a projection calling for an annual rate of 24 flights by 1990. Long before the Challenger accident, however, it was becoming obvious that even the modified goal of two flights a month was overambitious.
In establishing the schedule, NASA had not provided adequate resources for its attainment. As a result, the capabilities of the system were strained by the modest nine-mission rate of 1985, and the evidence suggests that NASA would not have been able to accomplish the 15 flights scheduled for 1986. These are the major conclusions of a Commission examination of the pressures and problems attendant upon the accelerated launch schedule.
On the same day that the initial orbital tests concluded-July 4, 1982-President Reagan announced a national policy to set the direction of the U. S. space program during the following decade. As part of that policy, the President stated that:
"The United States Space Transportation System (STS) is the primary space launch system for both national security and civil government missions."
Additionally, he said:
"The first priority of the STS program is to make the system fully operational and cost-effective in providing routine access to space."
From the inception of the Shuttle, NASA had been advertising a vehicle that would make space operations "routine and economical." The greater the annual number of flights, the greater the degree of routinization and economy, so heavy emphasis was placed on the schedule. However, the attempt to build up to 24 missions a year brought a number of difficulties, among them the compression of training schedules, the lack of spare parts, and the focusing of resources on nearterm problems.
One effect of NASA's accelerated flight rate and the agency's determination to meet it was the dilution of the human and material resources that could be applied to any particular flight.
The part of the system responsible for turning the mission requirements and objectives into flight software, flight trajectory information and crew training materials was struggling to keep up with the flight rate in late 1985, and forecasts showed it would be unable to meet its milestones for 1986. It was falling behind because its resources were strained to the limit, strained by the flight rate itself and by the constant changes it was forced to respond to within that accelerating schedule. Compounding the problem was the fact that NASA had difficulty evolving from its singleflight focus to a system that could efficiently support the projected flight rate. It was slow in developing a hardware maintenance plan for its reusable fleet and slow in developing the capabilities that would allow it to handle the higher volume of work and training associated with the increased flight frequency.
[165] Pressures developed because of the need to meet customer commitments, which translated into a requirement to launch a certain number of flights per year and to launch them on time. Such considerations may occasionally have obscured engineering concerns. Managers may have forgotten-partly because of past success, partly because of their own well-nurtured image of the program-that the Shuttle was still in a research and development phase. In his testimony before a U.S. Senate Appropriations subcommittee on May 5, 1982, following the third flight of the Space Shuttle, James Beggs, then the NASA Administrator, expressed NASA's commitment:
"The highest priority we have set for NASA is to complete development of the Shuttle and turn it into an operational system. Safety and reliability of flight and the control of operational costs are primary objectives as we move forward with the Shuttle program."1
Sixteen months later, arguing in support of the Space Station, Mr. Beggs said, "We can start anytime.... There's no compelling reason [why] it has to be 1985 rather than '86 or '87. The point that we have made is that the Shuttle is now operational." 2 The prevalent attitude in the program appeared to be that the Shuttle should be ready to emerge from the developmental stage, and managers were determined to prove it "operational."
Various aspects of the mission design and development process were directly affected by that determination. The sections that follow will discuss the pressures exerted on the system by the flight rate, the reluctance to relax the optimistic schedule, and the attempt to assume an operational status.
Planning of a Mission
The planning and preparation for a Space Shuttle flight require close coordination among those making the flight manifest, those designing the flight and the customers contracting NASA's services. The goals are to establish the manifest; define the objectives, constraints and capabilities of the mission; and translate those into hardware, software and flight procedures.
There are major program decision points in the development of every Shuttle flight. At each of these points, sometimes called freeze points, decisions are made that form the basis for further engineering and product development. The disciplines affected by these freeze points include integration hardware, engineering, crew timeline, flight design and crew training.
The first major freeze point is at launch minus 15 months. At that time the flight is officially defined: the launch date, Orbiter and major payloads are all specified, and initial design and engineering are begun based on this information.
The second major freeze point is at launch minus 7.7 months, the cargo integration review. During this review, the integration hardware design, Orbiter vehicle configuration, flight design and software requirements are agreed to and specified. Further design and engineering can then proceed.
Another major freeze point is the flight planning and stowage review at launch minus five months. At that time, the crew activity timeline and the crew compartment configuration, which includes middeck payloads and payload specialist assignments, are established. Final design, engineering and training are based on these products.
Development of Flight Products
The "production process" begins by collecting all mission objectives, requirements and constraints specified by the payload and Space Shuttle communities at the milestones described above. That information is interpreted and assimilated as various groups generate products required for a Space Shuttle flight: trajectory data, consumables requirements, Orbiter flight software, Mission Control Center software and the crew activity plan, to name just a few.
Some of these activities can be done in parallel, but many are serial. Once a particular process has started, if a substantial change is made to the flight, not only does that process have to be started again, but the process that preceded it and supplied its data may also need to be repeated. If one group fails to meet its due date, the group that is next in the chain will start late. The delay then cascades through the system.
Were the elements of the system meeting their schedules? Although each group believed it had an adequate amount of time allotted to perform its function, the system as a whole was falling....
....behind. An assessment of the system's overall performance is best made by studying the process at the end of the production chain: crew training. Analysis of training schedules for previous flights and projected training schedules for flights in the spring and summer of 1986 reveals a clear trend: less and less time was going to be available for crew members to accomplish their required training. (See the Shuttle mission simulator training chart.)
The production system was disrupted by several factors including increased flight rate, lack of efficient production processing and manifest changes.
Changes in the Manifest
Each process in the production cycle is based on information agreed upon at one of the freeze points. If that information is later changed, the process may have to be repeated. The change could be a change in manifest or a change to the Orbiter hardware or software. The hardware and software changes in 1985 usually were mandatory changes; perhaps some of the manifest changes were not.
The changes in the manifest were caused by factors that fall into four general categories: hardware problems, customer requests, operational [167] constraints and external factors. The significant changes made in 1985 are shown in the accompanying table. The following examples illustrate that a single proposed change can have extensive impact, not because the change itself is particularly difficult to accommodate (though it may be), but because each change necessitates four or five other changes. The cumulative effect can be substantial. (See the Impact of Manifest Changes chart.)
When a change occurs, the program must choose a response and accept the consequences of that response. The options are usually either to maximize the benefit to the customer or to minimize the adverse impact on Space Shuttle operations. If the first option is selected, the consequences will include short-term and/or longterm effects.
Hardware problems can cause extensive changes in the payload manifest. The 51-E mission was on the launch pad, only days from launch, with a Tracking and Data Relay Satellite and Telesat satellite in the cargo bay, when a hardware problem in the tracking satellite was discovered. That flight was canceled and the payload reassigned. The cancellation resulted in major changes to several succeeding flights. Mission 51 -D, scheduled to fly two months later, was changed to add the Telesat and delete the retrieval of the Long Duration Exposure Facility. The retrieval mission was then added to mission 61-I, replacing another satellite. A new mission (61-M) was scheduled for July, 1986, to accommodate the Tracking and Data Relay Satellite and the displaced satellite, and all flights scheduled later in 1986 slipped to make room for 61-M.
Customers occasionally have notified NASA Headquarters of a desire to change their scheduled launch date because of development problems, financial difficulties or changing market conditions. NASA generally accedes to these requests and has never imposed the penalties available. An example is the request made to delay the flight of the Westar satellite from mission 61-C (December, 1985) to a flight in March, 1986. Westar was added to flight 61-E, and the Getaway Special bridge assembly was removed to make room for it; the HS-376 satellite slot was deleted from 51-L and added to 61-C; the Spartan-Halley satellite was deleted from 61-D and added to 51 -L. Thus, four flights experienced major payload changes as a result of one customer's request.
1985 Changes in the Manifest
Hardware Problems
Tracking and Data Relay Satellite (canceled 51 -E, added 61 -M).
Synchronous Communication Satellite (added to 61-C).
Synchronous Communication Satellite (removed from 61-C).
OV- 102 late delivery from Palmdale (changed to 51 -G, 51 -I, and 61 -A).
Customer Requests
HS-376 (removed from 51-I). G-Star (removed from 61-C). Satellite Television Corporation-Direct Broadcast Satellite (removed from 61-E).
Westar (removed from 61-C).
Satellite Television Corporation-Direct Broadcast Satellite (removed from 61-H).
Electrophoresis Operations in Space (removed from 61-B).
Electrophoresis Operations in Space (removed from 61-H).
Hubble Space Telescope (swap with Earth Observation Mission).
Operational Constraints
No launch window for Skynet/Indian Satellite Combination (61-H).
Unacceptable structural loads for Tracking and Data Relay Satellite/Indian Satellite (61-H).
Landing weight above allowable limits for each of the following missions: 61-A, 61-E, 71-A, 61 -K.
External Factors
Late addition of SenatorJake Garn (A-Utah) (51 -D)
Late addition of Representative Bill Nelson (D-Florida) (61-C).
Late addition of Physical Vapor Transport Organic Solid experiment (51-I).
Operational constraints (for example, a constraint on the total cargo weight) are imposed to insure that the combination of payloads does not exceed the Orbiter's capabilities. An example involving the Earth Observation Mission Spacelab flight is presented in the NASA Mission Planning and Operations Team Report in Appendix J. That case illustrates that changes resulting from a single instance of a weight constraint violation can cascade through the entire schedule.
External factors have been the cause of a number of changes in the manifest as well. The changes discussed above involve major payloads, but changes to other payloads or to payload specialists can create problems as well. One small change does not come alone; it generates several others. A payload specialist was added to mission 61-C only two months before its scheduled lift off. Because there were already seven crew members assigned to the flight, one had to be removed. The Hughes payload specialist was moved from 61-C to 51-L just three months before 51-L was scheduled to launch. His experiments were also added to 51-L. Two middeck experiments were deleted from 51-L as a result, and the deleted experiments would have reappeared on later flights.
Again, a "single" late change affected at least two flights very late in the planning and preparation cycles.
The effects of such changes in terms of budget, cost and manpower can be significant. In some cases, the allocation of additional resources allows the change to be accommodated with little or no impact to the overall schedule. In those cases, steps that need to be re-done can still be accomplished before their deadlines. The amount of additional resources required depends, of course, on the magnitude of the change and when the change occurs: early changes, those before the cargo integration review, have only a minimal impact; changes at launch minus five months (two months after the cargo integration review) can carry a major impact, increasing the required resources by approximately 30 percent. In the missions from 41-C to 51-L, only 60 percent of the major changes occurred before the cargo integration review. More than 20 percent occurred after launch minus five months and caused disruptive budget and manpower impacts.3
Engineering flight products are generated under a contract that allows for increased expenditures to meet occasional high workloads.
[170] Even with this built-in flexibility, however, the requested changes occasionally saturate facilities and personnel capabilities. The strain on resources can be tremendous. For short periods of two to three months in mid-1985 and early 1986, facilities and personnel were being required to perform at roughly twice the budgeted flight rate.
If a change occurs late enough, it will have an impact on the serial processes. In these cases, additional resources will not alleviate the problem, and the effect of the change is absorbed by all downstream processes, and ultimately by the last element in the chain. In the case of the flight design and software reconfiguration process, that last element is crew training. In January, 1986, the forecasts indicated that crews on flights after 51-I. would have significantly less time than desired to train for their flights.4 (See the Simulation Training chart.)
According to Astronaut Henry Hartsfield:
"Had we not had the accident, we were going to be up against a wall; STS 61-H . . . would have had to average 31 hours in the simulator to accomplish their required training, and STS 61-K would have to average 33 hours. That is ridiculous. For the first time, somebody was going to have to stand up and say we have got to slip the launch because we are not going to have the crew trained." 5
"Operational" Capabilities
For a long time during Shuttle development, the program focused on a single flight, the first Space Shuttle mission. When the program became "operational," flights came more frequently, and the same resources that had been applied to one flight had to be applied to several flights concurrently. Accomplishing the more pressing immediate requirements diverted attention from what was happening to the system as a whole. That appears to be one of the many telling differences between a "research and development" program and an "operational program." Some of the differences are philosophical, some are attitudinal and some are practical.
Elements within the Shuttle program tried to adapt their philosophy, their attitude and their requirements to the "operational era." But that era came suddenly, and in some cases, there had not been enough preparation for what "operational" might entail. For example, routine and regular post-flight maintenance and inspections are critical in an operational program; spare parts arc critical to flight readiness in an operational fleet; and the software tools and training facilities developed during a test program may not be suitable for the high volume of work required in an operational environment. In many respects, the system was not prepared to meet an "operational" schedule.
As the Space Shuttle system matured, with numerous changes and compromises, a comprehensive set of requirements was developed to ensure the success of a mission. What evolved was a system in which the preflight processing, flight planning, flight control and flight training were accomplished with extreme care applied to every detail. This process checked and rechecked everything, and though it was both labor- and time-intensive, it was appropriate and necessary for a system still in the developmental phase. This process, however, was not capable of meeting the flight rate goals.
After the first series of flights, the system developed plans to accomplish what was required to support the flight rate. The challenge was to streamline the processes through automation, standardization, and centralized management, and to convert from the developmental phase to the mature system without a compromise in quality. It required that experts carefully analyze their areas to determine what could be standardized and automated, then take the time to do it.
But the increasing flight rate had priority- quality products had to be ready on time. Further, schedules and budgets for developing the needed facility improvements were not adequate. Only the time and resources left after supporting the flight schedule could be directed toward efforts to streamline and standardize. In 1985, NASA was attempting to develop the capabilities of a production system. But it was forced to do that while responding-with the same personnel-to a higher flight rate.
At the same time the flight rate was increasing, a variety of factors reduced the number of skilled personnel available to deal with it. These included retirements, hiring freezes, transfers to other programs like the Space Station and transitioning to a single contractor for operations support.
[171] The flight rate did not appear to be based on assessment of available resources and capabilities and was not reduced to accommodate the capacity of the work force. For example, on January 1, 1986, a new contract took effect at Johnson that consolidated the entire contractor work force under a single company. This transition was another disturbance at a time when the work force needed to be performing at full capacity to meet the 1986 flight rate. In some important areas, a significant fraction of workers elected not to change contractors. This reduced the work force and its capabilities, and necessitated intensive training programs to qualify the new personnel. According to projections, the work force would not have been back to full capacity until the summer of 1986. This drain on a critical part of the system came just as NASA was beginning the most challenging phase of its flight schedule.6
Similarly, at Kennedy the capabilities of the Shuttle processing and facilities support work force became increasingly strained as the Orbiter turnaround time decreased to accommodate the accelerated launch schedule. This factor has resulted in overtime percentages of almost 28 percent in some directorates. Numerous contract employees have worked 72 hours per week or longer and frequent 12-hour shifts. The potential implications of such overtime for safety were made apparent during the attempted launch of mission 61-C on January 6, 1986, when fatigue and shiftwork were cited as major contributing factors to a serious incident involving a liquid oxygen depletion that occurred less than five minutes before scheduled lift off. The issue of workload at Kennedy is discussed in more detail in Appendix G.
Another example of a system designed during the developmental phase and struggling to keep up with operational requirements is the Shuttle Mission Simulator. There are currently two simulators. They support the bulk of a crew's training for ascent, orbit and entry phases of a Shuttle mission. Studies indicate two simulators can support no more than 12- 15 flights per year. The flight rate at the time of the accident was about to saturate the system's capability to provide trained astronauts for those flights. Furthermore, the two existing simulator s are out-of-date and require constant attention to keep them operating at capacity to meet even the rate of 12-15 flights per year. Although there are plans to improve capability, funds for those improvements are minimal and spread out over a 10-year period. This is another clear demonstration that the system was trying to develop its capabilities to meet an operational schedule but was not given the time, opportunity or resources to do it.7
Responding to Challenges and Changes
Another obstacle in the path toward accommodation of a higher flight rate is NASA's legendary "can-do" attitude. The attitude that enabled the agency to put men on the moon and to build the Space Shuttle will not allow it to pass up an exciting challenge-even though accepting the challenge may drain resources from the more mundane (but necessary) aspects of the program.
A recent example is NASA's decision to perform a spectacular retrieval of two communications satellites whose upper stage motors had failed to raise them to the proper geosynchronous orbit. NASA itself then proposed to the insurance companies who owned the failed satellites that the agency design a mission to rendezvous with them in turn and that an astronaut in a jet backpack fly over to escort the satellites into the Shuttle's payload bay for a return to Earth.
The mission generated considerable excitement within NASA and required a substantial effort to develop the necessary techniques, hardware and procedures. The mission was conceived, created, designed and accomplished within 10 months. The result, mission 51-A (November, 1984), was a resounding success, as both failed satellites were successfully returned to Earth. The retrieval mission vividly demonstrated the service that astronauts and the Space Shuttle can perform .
Ten months after the first retrieval mission, NASA launched a mission to repair another communications satellite that had failed in low-Earth orbit. Again, the mission was developed and executed on relatively short notice and was resoundingly successful for both NASA and the satellite insurance industry.
The satellite retrieval missions were not isolated occurrences. Extraordinary efforts on NASA's part in developing and accomplishing missions will, and should, continue, but such efforts will be a substantial additional drain on resources. NASA cannot both accept the relatively spur-of [172] the-moment missions that its "can-do" attitude tends to generate and also maintain the planning and scheduling discipline required to operate as a "space truck" on a routine and cost-effective basis. As the flight rate increases, the cost in resources and the accompanying impact on future operations must be considered when infrequent but extraordinary efforts are undertaken. The system is still not sufficiently developed as a "production line" process in terms of planning or implementation procedures. It cannot routinely or even periodically accept major disruptions without considerable cost. NASA's attitude historically has reflected the position that "We can do anything," and while that may essentially be true, NASA's optimism must be tempered by the realization that it cannot do everything.
NASA has always taken a positive approach to problem solving and has not evolved to the point where its officials are willing to say they no longer have the resources to respond to proposed changes. Harold Draughon, manager of the Mission Integration Office at Johnson, reinforced this point by describing what would have to happen in 1986 to achieve the flight rate:
"The next time the guy came in and said 'I want to get off this flight and want to move down two' [the system would have had to say,] We can't do that,' and that would have been the decision." 8
Even in the event of a hardware problem, after the problem is fixed there is still a choice about how to respond. Flight 41-D had a main engine shutdown on the launch pad. It had a commercial payload on it, and the NASA Customer Services division wanted to put that commercial payload on the next flight (replacing some NASA payloads) to satisfy more customers. Draughon described the effect of that decision to the Commission: "We did that. We did not have to. And the system went out and put that in work, but it paid a price. The next three or four flights all slipped as a result." 9
NASA was being too bold in shuffling manifests. The total resources available to the Shuttle program for- allocation were fixed. As time went on, the agency had to focus those resources more and more on the near term-worrying about today's problem and not focusing on tomorrow's.
NASA also did not have a way to forecast the effect of a change of a manifest. As already indicated, a change to one flight ripples through the manifest and typically necessitates changes to many other flights, each requiring resources (budget, manpower, facilities) to implement. Some changes are more expensive than others, but all have an impact, and those impacts must be understood.
In fact, Leonard Nicholson, manager of Space Transportation System Integration and Operations at Johnson, in arguing for the development of a forecasting tool, illustrated the fact that the resources were spread thin: "The press of business would have hindered us getting that kind of tool in place, just the fact that all of us were busy . . . . "10
The effect of shuffling major payloads can be significant. In addition, as stated earlier, even apparently "easy" changes put demands on the resources of the system Any middeck or secondary payload has, by itself, a minimal impact compared with major payloads. But when several changes are made, and made late, they put significant stress on the flight preparation process by diverting resources from higher priority problems.
Volume III of JSC 07700, Revision B, specifies that all middeck experiments must be scheduled, and payload specialists assigned, 11 weeks before launch. 11 That rule has not been enforced-in fact, it is more honored in the breach than in the observance. A review of missions 41-G through 61-C revealed that of the 16 payload specialists added to those flights, seven were added after launch minus five months.
Even "secondary" payloads take a lot of time and attention when they are added to a flight late. Harold Draughon:
"I spend more than half of my time working on things that are not very important because they get put in so late. Rather than working on PAM's [Payload Assist Modules] and IUS's [Inertial Upper Stages], I am working on chicken eggs." 12
Those directing the changes in the manifest were not yet sensitive to the problem. Each change nibbles away at the operational resources, and the changes were occurring frequently, even routinely. Much of the capacity of the system was being used up responding to late changes in lower priority experiments. That flexibility toward secondary experiments tied up the resources that would have been better spent building capability to meet the projected flight rate.
[173] Tommy Holloway, chief of the Johnson Flight Director Office, emphasized that, given finite resources, one must decide: "It's flight rate versus [manifest] flexibility.'' l3
The portion of the system forced to respond to the late changes in the manifest tried to bring its concerns to Headquarters. As Mr. Nicholson explained,
"We have done enough complaining about it that I cannot believe there is not a growing awareness, but the political aspects of the decision are so overwhelming that our concerns do not carry much weight.... The general argument we gave about distracting the attention of the team late in the process of implementing the flight is a qualitative argument .... And in the face of that, political advantages of implementing those late changes outweighed our general objections. "14
It is important to determine how many flights can be accommodated, and accommodated safely. NASA must establish a realistic level of expectation, then approach it carefully. Mission schedules should be based on a realistic assessment of what NASA can do safely and well, not on what is possible with maximum effort. The ground rules must be established firmly, and then enforced.
The attitude is important, and the word operational can mislead. "Operational" should not imply any less commitment to quality or safety, nor a dilution of resources. The attitude should be, "We are going to fly high risk flights this year; every one is going to be a challenge, and every one is going to involve some risk, so we had better be careful in our approach to each."15
Effect of Flight Rate on Spare Parts
As the flight rate increases, the demand on resources and the demand for spare parts increases. Since 1981, NASA has had logistics plans for Shuttle flight rates of 12 and 24 flights a year. It was originally forecast (in mid-1983) that the supply of spares required to support 12 flights annually could be accomplished in the spring of 1986. Actual inventory of spare parts had run close to plan until the second quarter of fiscal year 1985. At that time, inventory requirements for spares began to increase faster than deliveries. A year later, when inventory stockage should have been complete, only 32,000 of the required 50,000 items (65 percent) had been delivered. 16
The spare parts plan to support 24 flights per year had called for completing inventory stockage by June, 1987. By mid-1985, that schedule was in jeopardy.
The logistics plan could not be fully implemented because of budget reductions. In October, 1985, the logistics funding requirement for the Orbiter program, as determined by Level III management at Johnson, was $285.3 million. That funding was reduced by $83.3 million-a cut that necessitated major deferrals of spare parts purchases. Purchasing deferrals come at great cost. For example, a reduction due to deferral of $11.2 million in fiscal year 1986 would cost $11.2 million in fiscal year 1987, plus an additional $21.6 million in fiscal year 1988. This three-to-one ratio of future cost to current savings is not uncommon. Indeed, the ratio in many instances is as high as seven to one. This practice cannot make sense by any standard of good financial management.
According to Johnson officials, reductions in spares expenditures provided savings required to meet the revised budgets. As Program Manager Arnold Aldrich reported to the Commission:
"There had been fund contentions in the program for a number of years, at least starting in the mid-seventies and running through into the early to mid-eighties . . . intentional decisions were made to defer the heavy build-up of spare parts procurements in the program so that the funds could be devoted to other more pressing activities. . . . It was a regular occurrence for several annual budget cycles. And once the flight rate really began to rise and it was really clear that spare parts were going to be a problem, significant attention was placed on that problem by all levels of NASA and efforts had been made to catch up. But . . . our parts availability is well behind the flight need . . ." 17
Those actions resulted in a critical shortage of serviceable spare components. To provide parts required to support the flight rate, NASA had to resort to cannibalization. Extensive cannibalization of spares, i.e., the removal of components [174] from one Orbiter for installation in another, became an essential modus operandi in order to maintain flight schedules. Forty-five out of approximately 300 required parts were cannibalized for Challenger before mission 51-L. These parts spanned the spectrum from common bolts to a thrust control actuator for the orbital maneuvering system to a fuel cell. This practice is costly and disruptive, and it introduces opportunities for component damage.
This concern was summarized in testimony before the Commission by Paul Weitz, deputy chief of the Astronaut Office at Johnson:
"It increases the exposure of both Orbiters to intrusion by people. Every time you get people inside and around the Orbiter you stand a chance of' inadvertent damage of whatever type, whether you leave a tool behind or whether you, without knowing it, step on a wire bundle or a tube or something along those lines." 18
Cannibalization is a potential threat to flight safety, as parts are removed from one Orbiter, installed in another Orbiter, and eventually replaced. Each handling introduces another opportunity for imperfections in installation and for damage to the parts and spacecraft.
Cannibalization also drains resources, as one Kennedy official explained to the Commission on March 5, 1986:
"It creates a large expenditure in manpower at KSC. A job that you would have normally used what we will call one unit of' effort to do the job now requires two units of effort because you've got two ships [Orbiters] to do the task with." 19
Prior to the Challenger accident, the shortage of' spare parts had no serious impact on flight schedules, but cannibalization is possible only so long as Orbiters from which to borrow are available. In the spring of 1986, there would have been no Orbiters to use as "spare parts bins." Columbia was to fly in March, Discovery was to be sent to Vandenberg, and Atlantis and Challenger were to fly in May. In a Commission interview, Kennedy director of Shuttle Engineering Horace Lamberth predicted the program would have been unable to continue:
"I think we would have been brought to our knees this spring [1986] by this problem [spare parts] if we had kept trying to fly " 20
NASA's processes for spares provisioning (determining the appropriate spares inventory levels), procurement and inventory control are complicated and could be streamlined and simplified.
As of spring 1986, the Space Shuttle logistics program was approximately one year behind. Further, the replenishment of all spares (even parts that are not currently available in the system) has been stopped. Unless logistics support is improved, the ability to maintain even a three-Orbiter fleet is in jeopardy.
Spare parts provisioning is yet another illustration that the Shuttle program was not prepared for an operational schedule. The policy was shortsighted and led to cannibalization in order to meet the increasing flight rate.
The Importance of Flight Experience
In a developmental program it is important to make use of flight experience, both to understand the system's actual performance and to uncover problems that might not have been discovered in testing. Because Shuttle flights were coming in fairly rapid succession, it was becoming difficult to analyze all the data from one flight before the next was scheduled to launch. In fact, the Flight Readiness Review for 51-L was held while mission 61-C was still in orbit. Obviously, it was impossible to even present, much less analyze and understand, anomalies from that flight.
The point can be emphasized by citing two problems that occurred during mission 61-C but were discovered too late to be considered at the 51-L Flight Readiness Review:
1. The Space Shuttle brakes and tires have long been a source of concern. In particular, after the 51-D Orbiter blew a tire at Kennedy in April, 1985, there was considerable effort (within budgetary constraints) to understand and resolve the problems, and Kennedy landings were suspended until certain improvements were made. (See section "Landing: Another Critical Phase!" page 186.) Mission 51 -L was to be the first flight to land [175] in Florida since 51 -D had experienced brake problems. STS 61-C landed at Edwards Air Force Base in California on January 19, 1986, four days after the 51-L Flight Readiness Review. The 61-C brakes were removed following landing and shipped to the vendor for further inspection and analysis. That inspection revealed major brake damage. The subsystem manager at Johnson in charge of the brakes did not receive the information until January 27, 1986, one day before 51-L was launched, and did not learn the extent of the problem until January 30, 1986.
2. The inspection of the 61 -C Solid Rocket Booster segments was completed on January 19, 1986, four days after the 51-L Level I Flight Readiness Review. The post-recovery inspection of the 61 -C Solid Rocket Booster segments revealed that there was O-ring erosion in one of the left booster field joints and additional O-ring anomalies on both booster nozzles. Although the information was available for Marshall's 51 -L Level III review at launch minus one day, it was clearly not available in time for consideration in the formal launch preparation process.21 These examples underscore the need to establish a list of mandatory post-flight inspections that must precede any subsequent launch.
Effect on Payload Safety
The payload safety process exists to ensure that each Space Shuttle payload is safe to fly and that on a given mission the total integrated cargo does not create a hazard. NASA policy is to minimize its involvement in the payload design process. The payload developer is responsible for producing a safe design, and the developer must verify compliance with NASA safety requirements. The Payload Safety Panel at Johnson conducts a phased series of safety reviews for each payload. At those reviews, the payload developer presents material to enable the panel to assess the payload's compliance with safety requirements.
Problems may be identified late, however, often as a result of late changes in the payload design and late inputs from the payload developer. Obviously, the later a hazard is identified, the more difficult it will be to correct, but the payload safety process has worked well in identifying and resolving safety hazards.
Unfortunately, pressures to maintain the flight schedule may influence decisions on payload safety provisions and hazard acceptance. This influence was evident in circumstances surrounding the development of two high priority scientific payloads and their associated booster, the Centaur.
Centaur is a Space Shuttle-compatible booster that can be used to carry heavy satellites from the Orbiter's cargo bay to deep space. It was scheduled to fly on two Shuttle missions in May, 1986, sending the NASA Galileo spacecraft to Jupiter and the European Space Agency Ulysses spacecraft first to Jupiter and then out of the planets' orbital plane over the poles of the Sun. The pressure to meet the schedule was substantial because missing launch in May or early June meant a year's wait before planetary alignment would again be satisfactory.
Unfortunately, a. number of safety and schedule issues clouded Centaur's use. In particular, Centaur's highly volatile cryogenic propellants created several problems. If a return-to-launch-site abort ever becomes necessary, the propellants will definitely have to be dumped overboard. Continuing safety concerns about the means and feasibility of dumping added pressure to the launch preparation schedule as the program struggled to meet the launch dates.
Of four required payload safety reviews, Centaur had completed three at the time of the Challenger accident, but unresolved issues remained from the last two. In November, 1985, the Payload Safety Panel raised several important safety concerns. The final safety review, though scheduled for late January, 1986, appeared to be slipping to February, only three months before the scheduled launches.
Several safety waivers had been granted, and several others were pending. Late design changes to accommodate possible system failure would probably have required reconsideration of some of the approved waivers. The military version of the Centaur booster, which was not scheduled to fly for some time, was to be modified to provide added safety, but because of the rush to get the 1986 missions launched, these improvements were not approved for the first two Centaur boosters. After the 51-L accident, NASA allotted more than $75 million to incorporate the [176] operational and safety improvements to these two vehicles.22 We will never know whether the payload safety program would have allowed the Centaur missions to fly in 1986. Had they flown, however, they would have done so without the level of protection deemed essential after the accident.
Outside Pressure to Launch
After the accident, rumors appeared in the press to the effect that persons who made the decision to launch mission 51-L might have been subjected to outside pressure to launch. Such rumors concerning unnamed persons, emanating from anonymous sources about events that may never have happened, are difficult to disprove and dispel. Nonetheless, during the Commission's hearings all persons who played key roles in that decision were questioned. Each one attested, under oath, that there had been no outside intervention or pressure of any kind leading up to the launch.
There was a large number of other persons who were involved to a lesser extent in that decision, and they were questioned. All of those persons provided the Commission with sworn statements that they knew of no outside pressure or intervention. 23
The Commission and its staff also questioned a large number of other witnesses during the course of the investigation. No evidence was reported to the Commission which indicated that any attempt was ever made by anyone to apply pressure on those making the decision to launch the Challenger.
Although there was total lack of evidence that any outside pressure was ever exerted on those who made the decision to launch 51-L, a few speculative reports persisted.
One rumor was that plans had been made to have a live communication hookup with the 51-L crew during the State of the Union Message. Commission investigators interviewed all of the persons who would have been involved in a hookup if one had been planned, and all stated unequivocally that there was no such plan. Furthermore, to give the crew time to become oriented, NASA does not schedule a communication for at least 48 hours after the launch and no such communication was scheduled in the case of flight 51-L.
The flight activity officer who was responsible for developing the crew activity plan testified that three live telecasts were planned for the Challenger, but they related in no way to the State of the Union Message: 24
The Commission concluded that the decision to launch the Challenger was made solely by the appropriate NASA officials without any outside intervention or pressure.
Findings
1. The capabilities of the system were stretched to the limit to support the flight rate in winter 1985/1986. Projections into the spring and summer of 1986 showed a clear trend; the system, as it existed, would have been unable to deliver crew training software for scheduled flights by the designated dates. The result would have been an unacceptable compression of the time available for the crews to accomplish their required training.
2. Spare parts are in critically short supply. The Shuttle program made a conscious decision to postpone spare parts procurements in favor of budget items of perceived higher priority. Lack of spare parts would likely have limited flight operations in 1986.
3. Stated manifesting policies are not enforced. Numerous late manifest changes (after the cargo integration review) have been made to both major payloads and minor payloads throughout the Shuttle program.
[177] 4. The scheduled flight rate did not accurately reflect the capabilities and resources.
5. Training simulators may be the limiting factor on the flight rate: the two current simulators cannot train crews for more than 12-15 flights per year.
6. When flights come in rapid succession, current requirements do not ensure that critical anomalies occurring during one flight are identified and addressed appropriately before the next flight.
References

Question: what was the name of the shuttle that was supposed to reduce the development costs of heavy

The concept of Shuttle C was theoretically expected to reduce the development costs of heavy-match launch missiles by the reuse technology developed in the Shuttle plan.

The proposal was made several times, both of which were to use the existing shuttle structures and the immediate main engines and navigation computers with a limit of use.

what was the name of the shuttle that was supposed to reduce the development costs of heavy

Question: what did the concept of shuttle c reduce development costs of

重量 物 打ち 上げ ロケット

what did the concept of shuttle c reduce development costs of

Question: how many times was the proposal for shuttle c made

how many times was the proposal for shuttle c made

Question: where was the reuse technology developed

where was the reuse technology developed

Question: what could reduce the development costs of heavy match launch missiles

what could reduce the development costs of heavy match launch missiles

Question: the concept of shuttle c was developed from what

the concept of shuttle c was developed from what

Question: besides the main engines, what other component of shuttle c was used

besides the main engines, what other component of shuttle c was used

Question: what was the concept of shuttle c supposed to reduce

development costs of heavy-match launch missiles

重量 物 打ち 上げ ロケット の 開発 費

development costs of heavy - match launch missiles

what was the concept of shuttle c supposed to reduce

The concept of Shuttle C was theoretically expected to reduce the development costs of heavy - match launch missiles by the reuse technology developed in the Shuttle plan .

The proposal was made several times , both of which were to use the existing shuttle structures and the immediate main engines and navigation computers with a limit of use .

シャトルCの概念は理論的にはシャトル計画で開発された再利用技術によって重量物打ち上げロケットの開発費を減らす事が期待された。 提案は複数回行われ、いずれも既存のシャトルの構造体や使用回数限度の迫ったメイン・エンジンや航法コンピュータを流用するというものだった。

シャトルCの概念は理論的にはシャトル計画で開発された再利用技術によって重量物打ち上げロケットの開発費を減らす事が期待された。

提案は複数回行われ、いずれも既存のシャトルの構造体や使用回数限度の迫ったメイン・エンジンや航法コンピュータを流用するというものだった。

The concept of Shuttle C was theoretically expected to reduce the development costs of heavy-match launch missiles by the reuse technology developed in the Shuttle plan. The proposal was made several times, both of which were to use the existing shuttle structures and the immediate main engines and navigation computers with a limit of use.

The concept of Shuttle C was theoretically expected to reduce the development costs of heavy-match launch missiles by the reuse technology developed in the Shuttle plan.

The proposal was made several times, both of which were to use the existing shuttle structures and the immediate main engines and navigation computers with a limit of use.

DIRECT DIRECT was proposed as an alternative to Ares I and Ares V proposed by NASA’s Vision for Space Exploration.

The original Shuttle derivative launch was called "Jupiter" and was planned to be launched in 2011 with a more ambitious "Project 2" composed of heavy-matter launch rocket Levitarian, orbit-circuit support station Olympia, Garon heavy-load spacecraft, spacecraft Algo and Helios and crew cargo ship Artemis.

Question: what was the name of nasa's vision for space exploration project

DIRECT DIRECT was proposed as an alternative to Ares I and Ares V proposed by NASA’s Vision for Space Exploration.

The original Shuttle derivative launch was called "Jupiter" and was planned to be launched in 2011 with a more ambitious "Project 2" composed of heavy-matter launch rocket Levitarian, orbit-circuit support station Olympia, Garon heavy-load spacecraft, spacecraft Algo and Helios and crew cargo ship Artemis.

what was the name of nasa's vision for space exploration project

Question: what was the original shuttle derivative launch called

what was the original shuttle derivative launch called

Question: when was the first shuttle scheduled to be launched

when was the first shuttle scheduled to be launched

Question: what is the name of the crew cargo ship of the shuttle project 2

what is the name of the crew cargo ship of the shuttle project 2

Question: DIRECT DIRECT was proposed as an alternative to what two Ares missions

アレス I と アレス V

DIRECT DIRECT was proposed as an alternative to what two Ares missions

Question: what is the name of the new shuttle launched in 2011

what is the name of the new shuttle launched in 2011

Question: what is the name of the orbit-circuit support station in project 2

what is the name of the orbit-circuit support station in project 2

Question: director DIRECT was proposed as an alternative to ares i and ares v proposed by

ビジョン ・ フォー ・ スペース ・ エクスプロレーション

director DIRECT was proposed as an alternative to ares i and ares v proposed by

DIRECT DIRECT was proposed as an alternative to Ares I and Ares V proposed by NASA ’s Vision for Space Exploration .

The original Shuttle derivative launch was called " Jupiter " and was planned to be launched in 2011 with a more ambitious " Project 2 " composed of heavy - matter launch rocket Levitarian , orbit - circuit support station Olympia , Garon heavy - load spacecraft , spacecraft Algo and Helios and crew cargo ship Artemis .

DIRECT DIRECTはNASAのビジョン・フォー・スペース・エクスプロレーションで提案されたアレスIとアレスVの代替案として提案された。 元のシャトル派生打ち上げ機では"ジュピター"と称され、より野心的な"プロジェクト2"で重量物打ち上げロケットのレビタリアン、軌道周回支援ステーションオリンピア、ガロン重量貨物宇宙船、宇宙ステーションアルゴとヘリオスと乗員貨物船アルテミスから構成され2011年に打ち上げる計画だった。

DIRECT DIRECTはNASAのビジョン・フォー・スペース・エクスプロレーションで提案されたアレスIとアレスVの代替案として提案された。

元のシャトル派生打ち上げ機では"ジュピター"と称され、より野心的な"プロジェクト2"で重量物打ち上げロケットのレビタリアン、軌道周回支援ステーションオリンピア、ガロン重量貨物宇宙船、宇宙ステーションアルゴとヘリオスと乗員貨物船アルテミスから構成され2011年に打ち上げる計画だった。

DIRECT DIRECT was proposed as an alternative to Ares I and Ares V proposed by NASA’s Vision for Space Exploration. The original Shuttle derivative launch was called "Jupiter" and was planned to be launched in 2011 with a more ambitious "Project 2" composed of heavy-matter launch rocket Levitarian, orbit-circuit support station Olympia, Garon heavy-load spacecraft, spacecraft Algo and Helios and crew cargo ship Artemis.

DIRECT DIRECT was proposed as an alternative to Ares I and Ares V proposed by NASA’s Vision for Space Exploration.

The original Shuttle derivative launch was called "Jupiter" and was planned to be launched in 2011 with a more ambitious "Project 2" composed of heavy-matter launch rocket Levitarian, orbit-circuit support station Olympia, Garon heavy-load spacecraft, spacecraft Algo and Helios and crew cargo ship Artemis.

DIRECT has three derivatives and the latest version 3.0 was announced in May 2009.

It was revealed in the public hearing of the Human Space Planning Committee held in Washington DC on June 17, 2009.

Question: when was the public hearing of the human space planning committee held

"2009 年 6 月 17 日 に ワシントン DC で 開催 さ れ た 有人 宇宙 飛行 計画 委員 会 の 公聴

DIRECT has three derivatives and the latest version 3.0 was announced in May 2009.

It was revealed in the public hearing of the Human Space Planning Committee held in Washington DC on June 17, 2009.

when was the public hearing of the human space planning committee held

Question: how many derivatives does DIRECT have

Question: when was the latest version of DIRECT released

when was the latest version of DIRECT released

Question: where was the public hearing of the human space planning committee held

where was the public hearing of the human space planning committee held

Question: who held a public hearing on the new version of DIRECT

有人 宇宙 飛行 計画 委員

who held a public hearing on the new version of DIRECT

Question: where was director 3.0 revealed in 2009

有人 宇宙 飛行 計画 委員 会 の

Question: how many derivatives does director have

DIRECT に は 三 つ の 派生 機種 が

DIRECT has three derivatives and the latest version 3.0 was announced in May 2009 .

It was revealed in the public hearing of the Human Space Planning Committee held in Washington DC on June 17 , 2009 .

DIRECTには三つの派生機種があり2009年5月に最新の3.0版が発表された。 2009年6月17日にワシントンDCで開催された有人宇宙飛行計画委員会の公聴会で明らかになった。

DIRECTには三つの派生機種があり2009年5月に最新の3.0版が発表された。

2009年6月17日にワシントンDCで開催された有人宇宙飛行計画委員会の公聴会で明らかになった。

DIRECT has three derivatives and the latest version 3.0 was announced in May 2009. It was revealed in the public hearing of the Human Space Planning Committee held in Washington DC on June 17, 2009.

DIRECT has three derivatives and the latest version 3.0 was announced in May 2009.

It was revealed in the public hearing of the Human Space Planning Committee held in Washington DC on June 17, 2009.