# Materials Design for Hypersonics

Adam B. Peters<EMAIL_ADDRESS>Dajie Zhang  Samuel Chen  Catherine Ott
Corey Oses
Stefano Curtarolo  Ian McCue  Tresa Pollock  Suhas Eswarappa Prameela
<EMAIL_ADDRESS>Department of Materials Science and Engineering, Johns Hopkins
University, Baltimore, MD 21218, USA  Johns Hopkins Applied Physics
Laboratory, Laurel, MD 20723  Department of Materials Science and Engineering,
Northwestern University, IL, USA. 60208  Department of Mechanical Engineering
and Materials Science, Duke University, NC, USA. 27708  Center for Autonomous
Materials Design, Duke University, NC, USA. 27708  Hopkins Extreme Materials
Institute, Johns Hopkins University, Baltimore, MD, USA, 21218  Departments of
Materials Engineering, University of California, Santa Barbara, CA, 93106
Department of Materials Science and Engineering, MIT, Cambridge, MA, USA,
02139  Department of Aeronautics and Astronautics, MIT, Cambridge, MA, USA,
02139

###### Abstract

Hypersonic vehicles must withstand extreme conditions during flights that
exceed five times the speed of sound. This class of vehicles has the potential
to facilitate rapid access to space, bolster defense capabilities, and create
a new paradigm for transcontinental earth-to-earth travel. However, the
extreme aerothermal environments resulting from high Mach number trajectories
create significant challenges for vehicle materials and structures. As
hypersonic systems advance, there is a critical need to develop novel
materials that are resilient to a combination of thermal and mechanical loads,
aggressive oxidizing environments, and rapid heating rates. This work aims to
provide a succinct discussion of emerging design strategies for refractory
alloys, composites, and ceramics used for hypersonic vehicles. We will
highlight key design principles for critical vehicle areas such as primary
structures, thermal protection, and propulsion systems; the role of theory and
computation in elucidating structure-property-processing relationships; and
strategies for advancing laboratory scale materials to flight-ready components
such as aerostructures and thermal protection systems.

###### keywords:

Hypersonics $|$ Materials Design $|$ Extreme Environments $|$ Thermal
Protection Systems $|$ High Entropy Alloys $|$ Ultra High-Temperature Ceramics

††journal: Journal

###### Contents

  1. 1 Introduction - Background to Hypersonics
  2. 2 Hypersonic Vehicle Configurations and Design Requirements
    1. 2.1 Primary Aerostructure
    2. 2.2 Thermal Protection System
    3. 2.3 Air-Breathing Propulsion Systems
  3. 3 Materials-class-specific considerations and design criteria
    1. 3.1 Metallic Materials for Hypersonics
    2. 3.2 Carbon Composites for Hypersonics
    3. 3.3 Ultra-high Temperature and Refractory Ceramics for Hypersonics
      1. 3.3.1 Ceramics-Matrix Composites
  4. 4 Advances in Computational Tools for Materials Development
    1. 4.1 Advances in Theory and Computational Tools
  5. 5 Role of integrated experiments and flight readiness pathways
  6. 6 Conclusion

##  1 Introduction - Background to Hypersonics

In the last decade, there has been a resurgence in hypersonic vehicle
development driven by the desire to increase flight performance and
reusability. Hypersonics refers to flight and aerodynamic phenomena that occur
above Mach 5 (5 times the speed of sound). To frame hypersonic speeds, a non-
stop flight from Los Angeles to Tokyo aboard a commercial airliner (Mach 0.8)
takes roughly twelve hours, whereas onboard an emerging Mach 9 hypersonic
vehicle it takes one. Although the first hypersonic flight was achieved
${\sim}$70 years ago, there has been increasing interest from a broader
audience due to modern engineering advances that are poised to revolutionize
defensive capabilities, sub-orbital travel, and rapid access to space [1, 2,
3] (Figure 1). New vehicle systems with ever-increasing capabilities and Mach
numbers are being developed, including: boost-glide systems, reusable
aircraft, space-launch vehicles, and missile technologies [1]. However, these
remarkable leaps in Mach number and performance during atmospheric flight come
with an array of formidable challenges in the domain of materials multi-
property optimization, simulation, and design [4]. Vehicles are purpose-built
with bespoke materials to operate at vastly different Mach numbers (5-25+),
altitudes (spanning sea level to orbit), hypersonic flight times (ranging from
seconds to hours), and trajectories.

When vehicle speeds increase past supersonic conditions and into the
hypersonic regime, the physics of external aerodynamic flows become dominated
by aerothermal heating rather than aerodynamic forces (Figure 1a). Aerodynamic
compression and friction create high-enthalpy gas dynamics that impart
additional physical phenomena from the energy exchange of a superheated
atmosphere. This superheated atmosphere results in: high heat fluxes (3-7
orders of magnitude greater than the 1.4 kW/m2 from the sun); extreme thermal
gradients (changing from -170°C to 3000°C across distances of order 1 cm);
high stagnation pressures (${\sim}$105-107 Pascals); and destructive plasma
from gas ionization which accelerates materials oxidation [1, 5, 6]. As
operational Mach numbers increase, these formidable phenomena must be
accommodated by materials in the principal subsystems of a hypersonic vehicle:
aeroshell/primary structure, leading edges, control surfaces, acreage thermal
protection, propulsion, and guidance systems. Developing hypersonic materials
has become the focus of cutting-edge research and are presently rate-limiting
steps for the resilience of structures during operation in extreme
environments.

Recent work reports on materials development for the propulsion system [7, 8,
9] thermoelectric generators [10], radomes [11], structural materials [12],
and thermal protection systems [12, 13]. Materials for hypersonics can be
broadly classified into three types: refractory metals, composites and
ceramics. Each material offers distinct tradeoffs for a given sub-system and
environmental application. Common metals and alloys in hypersonics, such as
aluminum and nickel-base superalloys, are favorable for primary structural
components and moderate thermal loads, while refractory metals with higher
operating temperatures, are employed for structures that see more demanding
operating conditions (see Box 1). Refractory ceramics combine high-temperature
capability with moderate thermal conductivity, but lack monolithic thermal
shock resistance and tend to be used as a thermal barrier coating or thin
structural materials [14]. By contrast, fiber-reinforced composite materials,
such as carbon/carbon or ultra-high temperature ceramics matrix composites,
incorporate carbon or ceramic fibers in dense matrices to improve high
temperature strength-to-weight ratios beyond metals [5], [14, 15, 16].

Advancement of these materials from laboratory scale studies to flight is
hindered by standardization of materials processing; reproducibility of
materials data; and difficulties testing representative thermal, oxidative,
and mechanical flight conditions. High-fidelity models have historically been
used to design flight trajectories to bound materials selection criteria. New
materials design tools with integrated computation and predictive frameworks
are emerging that can aid the design of complex materials and expand vehicle
performance and reliability. We will explore how refractory metals,
composites, and ceramics are designed and selected for hypersonic applications
according to vehicle-specific design criteria.

Figure 1: (a) Computational fluid dynamics (CFD) simulation of the X-43
vehicle at a Mach 7 test condition with the engine operating. The solution
includes internal (air-breathing scramjet engine) and external flow fields,
including the interaction between the engine exhaust and vehicle aerodynamics.
The image illustrates surface heat transfer on the vehicle (red is the highest
heating) and flow field contours at the local Mach number. Structural
components and the associated materials used for the design of the X-43
hypersonic vehicle are indicated: (b) alumina borosilicate insulation tile
with an emissive coating used for acreage protection thermal protection [17];
(c) nose and leading-edge design integrating carbon composites and refractory
tungsten alloy SD 18; (d) sharp leading edge cross-section showing the carbon
composite with a refractory Ir coating [5]; (e) airframe of the vehicle
composed of steel/aluminum skin and Al/Ti bulkheads. (f-n) Timeline of
hypersonic vehicle development spanning hypersonic airplanes, space access,
re-entry, boost-glide vehicles, and cruise missile applications, where colors
indicate the hypersonic vehicles configuration: (f) the first to reach
hypersonic speeds, Project Bumper-WAC” Without Any Control” (1949), (g) the
reusable X-15 research aircraft (1959), (h) Apollo re-entry capsules
(1961-1972), (i) Space Shuttle (1972-2011), (j) NASA X-43 airplane (2001), (k)
the HVT-2 boost-glide vehicle (2010-2012), (i) Boeing X-51 scramjet
(2010-2013), (m) SpaceX Starship (slated for hypersonic re-entry in 2023), (n)
a notional future hypersonic airplane. (Image sources: NASA (a,b,e,f,h,i,j),
(c) – adapted from [18] , (d) – adapted from [5], U.S. Airforce (g,l), DARPA
(k), Creative Commons - Offical SpaceX (m), U.S. Govt. images not subject to
copywrite)

##  2 Hypersonic Vehicle Configurations and Design Requirements

Material requirements for hypersonic flight are sensitively coupled to the
vehicle design and flight envelope, which impose two-principal environmental
challenges: (1) thermal loads that are dependent on both geometry and location
on the vehicle; (2) strongly oxidizing conditions that drive changes in both
material properties (oxidation) and geometry (ablation). As a result,
aerostructures, wing leading edges, acreage thermal protection systems, and
propulsion systems necessitate vastly different materials to accommodate these
diverse thermo-chemo-mechanical loads. Depending on the flight conditions
(Mach and altitude), flight time at a given Mach number and altitude (known as
time on condition), and location on the vehicle, qualified materials may not
exist for the desired application [12].

Aerothermal heating arises as the hypersonic vehicle pierces through the
atmosphere. Fundamentally, the adiabatic dissipation of a vehicle’s kinetic
energy into the viscous gas environment is responsible for the extreme thermal
conditions of flight [19]. In the vehicle’s shock layer, (volume gas between
the body and the shock wave), stagnation temperature increases proportionally
to Mach to the third power and root of the atmospheric density, and can reach
values as high as 10,000°C [20]. Although much of the energy is swept away
with the surrounding gas flow around the vehicle, energy transfer by
convective or radiative heating generates high heat fluxes that necessitate
materials capable of resisting high temperatures [7].

Material requirements are further exacerbated due to the dissociation of O2
and N2 into free radicals at gas-phase temperatures above 3000°C (typically at
speeds greater than Mach 8). These conditions lead to highly reactive surface
chemical interactions, causing materials degradation, microstructural
evolution, phase formation, and property changes during flight [6, 12, 21].
Critical challenges for materials designers remain in both the leading-edge
surfaces from direct aerothermal exposure (nose, cowl lips, and control
surfaces, Figure 1a), and in the propulsion flow path where radiative cooling
is not viable [12]. In the following section, we will highlight how the
characteristics of the primary aerostructures, thermal protection systems, and
propulsion systems influence material design and selection.

###  2.1 Primary Aerostructure

Lightweight primary structures (e.g., aeroshells and airframes) may be formed
into either lifting bodies or ballistic structural elements, where the
leading-edge profile and flight trajectory govern the aerothermal load during
flight. Unlike the traditional atmospheric re-entry vehicle designs in Figure
1h – which employ blunt features to increase drag and push the shock region
away from the structure and transfer energy into the air – hypersonic vehicles
require slender primary structures and sharp control surfaces to reduce drag
and enable stable long-distance accuracy. However, the heating rate is
inversely proportional to the square root of the tip radius and must be
accommodated through various energy dissipation mechanisms.

In modern vehicles, aeroshells are designed using solid or sandwich
constructions with honeycomb, lattice, corrugated, or foam cored to minimize
weight while maintaining rigidity and enable advanced passive cooling
strategies [5, 15, 16, 18, 22, 23]. Robust carbon and ceramic composites
remain materials of choice for modern leading-edge structures [5, 15, 16, 18,
22, 23], and enable peak temperature reduction through passive cooling by
employing favorable composite weave patterns, or thermally conductive
materials to more effectively transport heat to the colder regions of the
aeroshell main body [16, 22]. Such designs are commonly referred to as “hot
structures” (Figure 1k, l) as compared to the insulated “cold structure”
design adopted by the Space Shuttles and many other types of reentry vehicles
or bodies that use thick outer surface thermal insulation (Figure 1i,m ).

###  2.2 Thermal Protection System

Thermal protection materials and system design has become an engineering field
of its own because materials with unique property combinations can enable new
flight capabilities. Thermal protection systems (TPS) are employed for thermal
regulation of leading edges, nose, and propulsion features that experience the
greatest heat flux, as well as acreage locations that protect the aeroshell’s
fuselage and control surfaces (i.e., rudders, elevons; Figure 1a). In modern
vehicles, the aerostructure may have an integrated TPS for optimal heat
transfer and dissipation. TPS materials are selected to best accommodate the
local aerothermodynamic criteria according to their combinations of high-
temperature strength, thermal conductivity, heat capacity, melting/oxidation
temperature, and emissivity. Broadly, there are three fundamental types of TPS
used to increase vehicle resilience to aerothermal heating [5, 12], [14, 15,
16].

Passive thermal protection systems are ideal for moderate transient heat flux
scenarios, and may be composed of: (i) insulated cold structures (e.g., Space
Shuttle tiles); (ii) heat sink surface structures that both absorb and radiate
energy (e.g., skin of the X-15); or (iii) or emissive “hot structures” that
lower the thermal load through both environmental radiation and conduction
into the vehicle (e.g., nose of the X-51). Examples of these appear in Figure
1g, i. Semi-passive systems are implemented for high heat fluxes that persist
for long durations, and encompass: (i) reusable heat pipes that transfer and
radiate thermal energy via evaporative cooling and capillary wicking (e.g.
liquid lithium or potassium [24]); or (ii) single-use ablatives materials that
absorbed energy via pyrolysis or charring of a reinforced polymer/resin (used
in the first re-entry capsules, Figure 1h).

Active thermal protection systems, using the forced flow of a liquid or vapor,
are employed for the most extreme heat fluxes and extended flight durations
[25, 26, 27, 28, 29]. These systems include: (i) convective cooling
architectures, which transfer heat into a working fluid (e.g., Shuttle main
engine), (ii) film cooling, whereby a fluid is injected over a large area into
the flow to form an insulating blanket (e.g., X-43 propulsion system), (iii)
or transpiration cooling where a fluid (e.g., H2O or He) is injected into hot
gas flow through porous structures. Examples of these appear in Figure 1a, i,
j.

Nose and wing leading edges that are subjected to intense heat loading may
employ heat pipes or actively cooled structures for thermal regulation [12,
30]. By contrast, acreage locations — large fuselage regions on the vehicle —
experience lower heat fluxes and have historically been passively cooled using
materials with low thermal conductivities and thermal expansion coefficients
(see Box 1). Although heat sinks and thermal insulation are attractive from a
risk management perspective, they suffer from excessive mass and low fracture
toughness. Ablative materials have been used to great effect for shuttle re-
entry but do not favor reusability. In contrast, hot structures, heat pipes,
and active thermal management systems (enabled by additive manufacturing) have
dominated research efforts for new materials and system designs [31, 32, 33,
34, 35, 36, 37, 38].

Each TPS will have a unique architecture and thermal profile, which results in
their own set material property requirements. An example of simulated
aerodynamic heating for a passive (hot structure), semi-active (heat pipe),
and active (transpiration cooling) cooled leading edge is shown in Figure 2a,
b at one flight condition. In this example, each leading-edge experiences the
same heat flux and stagnation temperature because these are dictated by the
component geometry and flight condition. However, the resulting temperature
profile depends on the TPS mechanism. The passive leading-edge exhibits the
highest peak temperature and thermal gradient because it is solely relying on
intrinsic material properties (conductivity, heat capacity and emissivity).
The semi-passive leading edge exhibits a small thermal gradient (but a similar
peak temperature to the passive structure) because heat pipes increase thermal
conductivity by 1-3 orders of magnitude [39]. Lastly, the actively cooled
leading edge has the lowest peak temperature because transpiration reduces the
incident heat flux [40, 41, 42, 43].

Figure 2: Steady-state finite element (FE) simulations of aerodynamic heating
of a leading edge, carried out for a range of structural materials and
hypothetical hypersonic flight conditions. (a) Illustration of the sharp
leading-edge geometry used in these simulations with the following dimensions:
2.5 mm tip radius, 3-degree wedge angle, 5 cm span, 10 cm cord length. (b)
Thermal profiles across a 2D TZM leading-edge, considering passive thermal
management, a semi-passive Li heat pipe operating at 1500 K, and active
cooling via transpiration where the incident heat flux is reduced by a factor
of 2. (c) Ashby map highlighting operational tradeoffs for metal alloy, UHTC,
refractory alloy, and carbon-base material classes. (d) Ashby-style plot of
the FE simulation results from passive leading edges, where: the y-axis is the
normalized mechanical stress resulting from a thermal expansion gradient, and
the x-axis is the normalized peak temperature at the tip where the heat flux
is highest. Only 4 materials are not viable from a temperature standpoint
(ignoring oxidation), whereas 8 are not viable due to the expansion stress
exceeding the yield strength of the material at that temperature. Constraints
via oxidation will decrease the overall maximum operating temperature; there
is limited availability on oxidation kinetics for these materials. (e) The
culmination of (d) for different flight conditions is shown as a hypothetical
hypersonic flight corridor, where each line represents the “survivability
limit” for a monolithic material with this specific (sharp) wedge geometry;
known flight conditions of the X-43A, X-15, and typical space re-entry are
indicated for reference.

###  2.3 Air-Breathing Propulsion Systems

Similar to environmentally facing thermal protection systems, existing
approaches to hypersonic propulsion systems can be significantly improved with
refractory materials capable of operating in stressing aerothermal oxidizing
and reducing environments [44]. Currently, the ram/scramjet engine is the
standard form of propulsion for air-breathing hypersonic vehicles. Unlike
rocket-propelled hypersonic vehicles (e.g., X-15 and the Space shuttle), the
oxidizer for the propellent is supplied by the surrounding air and mixed
within a combustor using onboard fuel [45]. More advanced combined cycle
multi-mode propulsion systems under development include: rocket-based combined
cycle (RBCC); turbine-based combined cycle (TBCC); and turbo rocket combined
cycle (TRCC) that are capable of transition propulsion modes (such as rocket
propulsion during the initial ascent phase and then transition to air-
breathing scramjet engines at hypersonic speeds Figure 1a).

Specific components in these propulsion systems, including inlet ducts,
nozzles, and combustors, experience extreme temperature and mechanical stress
without the ability to readily dissipate heat through radiative cooling.
Propulsion system materials include a combination of refractory alloys, CMC’s,
C/C, and metal matrix composites (MMC) [46]. Textile-based CMCs may be formed
into complex structures for internal coolant flow and mechanical stiffening,
but their surface temperatures are limited to ${\sim}$1600°C. For velocities
near Mach 6, passively cooled refractory materials can be used to operate at
temperatures near that of the propulsive flow, but active cooling is required
above Mach 6. Active cooling methods require materials that can accommodate
high temperatures, pressure, and temperature gradients between the cooled fuel
and combustion chamber [12].

Materials lifetimes in this environment are largely controlled by oxidation,
which is highly dependent on flow conditions when water vapor is present in
the propulsion flow – this oxidation mechanism is quite different from the
ionized flow in leading-edge applications. Oxidation is exacerbated by thermal
gradient-induced microcracking, and limitations in high-fidelity modeling make
materials properties insufficient for lifetime prediction [12, 18, 22]. These
futuristic propulsion systems could enable low-cost and reusable air-breathing
hypersonic vehicles for manned flights and civilian transportation, but
materials development is necessary. More efficient materials will require
accurate prediction of engine thermal balance, heat loads, shock conditions
and oxidative character of the burning atmosphere [44].

Box 1: Material Classes for Hypersonics. Materials properties such as tensile
strength ($\sigma_{TS}$), yield strength ($\sigma_{y}$), flexural strength
($\sigma_{FS})$, density ($\rho$), coefficient of thermal expansion (CTE),
thermal conductivity ($\kappa$), various use temperatures (T), and heat flux
(Q) are given for metallic alloys, composites, ceramics and synthetics. Some
materials data are not available due to limited availability. Some of the
variation in the material’s properties is highly variable based on weave,
composite structure, and processing.The numbers in boxes correspond to the
materials used on the vehicles depicted in Figure 1. Common material property
data given from [47, 48]

##  3 Materials-class-specific considerations and design criteria

Materials selection is typically applied after structural components have been
designed and trajectories have been determined [49]. Initial material
screening can be carried out using thermo-mechanical simulations. For a given
set of material properties, conditions (heat flux and stagnation temperature)
are applied across a component to calculate the resulting thermal profile,
which is then used as boundary conditions to calculate thermal stresses. This
screening is useful in determining whether the peak temperature exceeds a
material’s melting point and/or the thermal stress exceeds the material’s flow
stress at the given temperature.

An example of this screening is illustrated in Figure 2c-e for a sharp passive
leading edge. As a hypothetical evaluation, steady-state thermal simulations
were carried out on high-temperature structural materials for over 300 Mach
and altitude combinations (i.e., assuming the material was exposed to these
conditions indefinitely and allowed to equilibrate). For each simulation,
following the work of [50], we extracted the peak temperature and estimated
the tip stress from the thermal gradient in this region. The incident heat
flux changes substantially around the tip, Figure 2b, causing a steep thermal
gradient (as large as 1000 K across 4 mm), which generates stresses of order
100 MPa due to non-uniform thermal expansion. If these stresses are above the
materials’ flow stress at the peak temperature, the tip will deform and affect
the boundary layer, potentially causing a laminar-to-turbulent transition

These mesoscale models provide critical insights into guiding what materials
should be used in the various structures (mentioned in the previous section)
for a given set of flight conditions. For instance, Figure 2 highlights that
traditional alloys (e.g., Ti-base, Ni-base and steels) have limited use as
leading edges due to low melting points, high-thermal expansion coefficients
and moderate thermal conductivities (Figure 2d). Monolithic ceramics suffer
from high thermal stresses, but their strengths can be modified through
secondary phases (see subsection 3.3) or employed as coatings. Refractory
metals – owing to their high strength at temperature, thermal transport
properties, and low thermal expansion coefficients – are ideal, but oxidation
kinetics will constrain their maximum service temperature (see subsection
3.1).

###  3.1 Metallic Materials for Hypersonics

Metallic materials are ubiquitously used in hypersonic vehicles – as nose and
wing leading edges, control surfaces, and engine inlets – due to their
tendency to their damage tolerance and manufacturability. These components
need to withstand extremely high heat fluxes and thermal strains, which demand
materials with high melting points that maintain strength at high
temperatures.

Pure elements with high melting points (W, Re, Ta, Mo, Nb, V, Cr, Ti, Ni) form
the basis of fielded high-temperature alloys. For instance, Ti was employed in
hot aeroshell structures in the SR-71 [5, 15] the nose section of the X-43
contained a SD 180 tungsten heavy alloy [51], a Haynes Ni-base alloy was used
in the Mach 7 X-43 variant [52], and both MoRe and Ni-base alloys have been
tested for heat pipe structures [11]. Other refractory metals, such as the
TaWHf alloys T111 and T222 (see Box 1) exhibit favorable creep-resistant
materials properties and are ideal for the extended containment of heated
liquid alkali-metal working fluids (1000-1300°C) for heat pipe type leading
edge designs [53]. T111/Li and niobium-based C-103/Na designs have been
assessed to satisfy the requirements for Mach 8 and Mach 10 flights
respectively [24].

However, melting point alone is not a singularly meaningful parameter of
structural design. For comparison, carbon-carbon will not melt at 1 atm but
will sublime at 3727°C and oxidize to CO(g) at temperatures starting as low as
${\sim}$370°C. As in all flight applications, density, oxidation resistance
and the ability to tolerate thermomechanical loading are important
considerations (Figure 2c). Nickel alloys are capable of operating at high
fractions of their melting points due to both coherent precipitation
strengthening and self-passivation that persist to very high temperatures.
Meanwhile, W and Mo can maintain over 50% of their Young’s Modulus at 2000°C
but oxidize well below 1000°C, with more rapid oxidation at increasing
temperature due to the high vapor pressures of their trioxides [5, 13, 14].
The combined property advantages in Ni- and Co-base alloys are often not found
in state-of-the-art refractory alloys, inhibiting their true operational
potential.

One promising class of metallic materials are the newly developed multi-
principal element (“high entropy”) alloys [54, 55, 56, 57, 58]. In particular,
refractory multi-principal element (RMPE) alloys offer the opportunity to
maintain high-temperature properties while, for example, decreasing density
and oxidation kinetics. RMPE alloys, such as MoNbTaVW ($\sigma_{y}$ = 1246
MPa, $\rho$ = 12.4 g/cm3) [56] and NbMoTiVZr ($\sigma_{y}$ = 1785 MPa, $\rho$
= 7.1 g/cm3) [59], could provide significant benefits over legacy refractory
alloys such as Ta-10W, which were chosen for re-crystallization behavior
rather than strength ($\sigma_{y}$ = 471 MPa, $\rho$ = 16.8 g/cm3) [60].
Unfortunately, many thermophysical properties and operation-relevant
properties, such as recrystallization temperature, have yet to be measured for
many of these alloys. While multidisciplinary design approaches have
successfully been implemented for the aerothermal and mechanical design of
hypersonic vehicles [61], materials have yet to be factored into this dynamic
design optimization loop.

Due to their limited oxidation resistance, alloys in hypersonic environments
typically rely on a compatible coating. Coatings may be multilayered,
functioning as both thermal and environmental barriers, with oxide-forming
metallic layers and porous, low-conductivity ceramic overcoats [62]. Nickel
alloys are designed to form alumina as a protective oxide and have well-
developed aluminide coating systems due to their extensive use in aircraft
engines [62].However, coatings are much less developed for refractory alloys
and typically contain metal silicides, which have limited protection below 850
°C and slough off above 1700°C due to aeroshearing However, coatings are much
less developed for refractory alloys and typically contain metal silicides,
which have limited protection below 850°C and slough off above 1700°C due to
aeroshearing 1700°C due to aeroshearing [63, 64]. A wide range of potential
coating failure modes need to be considered during design, and while there has
been considerable progress on understanding the mechanics of coatings [57,
65], often, the material properties are missing.

Out of the numerous legacy refractory alloys developed by NASA, only a handful
are manufactured and used today (such as C103, TZM, W25Re). In the past, cost
and formability at ambient temperatures were the roadblocks in fielding
superior refractory alloys. For instance, Nb alloys with high fractions of W
and Hf (greater than 15 wt.%) suffer from ductile-to-brittle transitions
several hundred degrees above room temperature [66]. However, rapidly evolving
additive manufacturing capabilities can now produce complex component designs
– negating machining constraints [67, 68, 69, 70]. This provides new
opportunities for the design of advanced cooling methodologies such as active
TPS [67, 68, 69, 70]. These new manufacturing techniques can also be combined
with high-throughput characterization and machine learning algorithms to
rapidly discover and develop the next-general refractory alloys for
hypersonics [30, 71, 72].

###  3.2 Carbon Composites for Hypersonics

Carbon-carbon composites (C/C) are historically considered the de facto
materials for the fabrication of hypersonic aeroshells and leading edges owing
to their excellent performance characteristics, including: low density
(1.60-1.98 g/cm3), low coefficient of thermal expansion (-0.85 to 1.1 x
10${}^{-}6$/K), high modulus of elasticity (200 GPa), high thermal
conductivities (${\sim}$4-35 W/mK) and retained mechanical properties up to
${\sim}$2000°C in inert environments [5, 12, 16, 18, 22], [73, 18, 74]. Low-
density C/C is often preferred over metals for severe environment
aerostructure elements. For instance, the horizontal Haynes control surfaces
of the Mach 7 variant of X-43 were replaced with coated C/C for the Mach 10
variant [18].

C/C is manufactured using one of two processes to densify a 2-D “zero
condition” C/C preform: (1) infiltration and pyrolysis (PIP) in which a high
carbon yield resin (phenolic or pitch) is infiltrated into the fibrous pre-
form (${\sim}$40-60 vol% fiber) before undergoing high-temperature
graphitization of the resin matrix; or (2) chemical vapor infiltration (CVI),
where densification occurs by the infiltration and decomposition of
carbonaceous gases. Both manufacturing methods have repeated steps to
consolidate the matrix, increase density, and increase strength. For PIP, 4-6
pyrolysis cycles under hot isostatic pressure (HIP) are preferred to inhibit
the formation of closed pores during resin pyrolysis and achieve a high
density ($>$ 98%). On the other hand, surface carbon deposition via CVI
inhibits further infiltration over time so the surface deposit needs to be
removed and the process restarted. Advanced C/C 6 (ACC-6) remains the state-
of-the-art composite, having undergone six impregnation cycles to increase
density and achieve high yield strengths at elevated temperatures.

Porosity reduction is a focal point of processing research because it is
critical in limiting oxygen diffusion and aerothermal erosion during flight.
Uncoated C/C formed by PIP and CVI have demonstrated in-plane tensile
strengths on the order of 165 MPa (with strength increasing as a function of
density) [75]. Furthermore, the thermomechanical properties of C/C are highly
anisotropic and dependent on the processing method, residual porosity, and
fiber architecture (fiber-woven fabrics or fiber tow windings may be
orientated at an angle, such as 30°, 60°, or 90°). Given this complexity, the
need for high-fidelity modeling capabilities for anisotropic, and locally
varied, volume elements is critical, which makes materials properties
standards insufficient for design, performance, and life prediction [12].

Despite these promising properties, uncoated C/C erodes rapidly at elevated
temperatures. The oxidation of carbonaceous composites begins ${\sim}$370°C in
air, with dramatic oxidation occurring beyond 500°C [13]. Present hypersonic
materials design efforts aim to protect C/C from high-temperature oxidation,
ablation, and erosion from prolonged and repeated aerothermal exposure. With
increasingly more extreme hypersonic environments, two protection approaches
are being developed: (1) deposition of high-temperature protective coatings,
and (2) modification of the carbon-carbon matrix. Oxidation-resistant coating
materials may be applied to C/C surface (or fibers before infiltration) to
limit diffusion and modify emissivity for passive thermal regulation.

The initial development of anti-oxidation coatings for C/C was focused on
refractory compositions containing SiC, HfB2, and ZrB2 (HfB2/SiC and ZrB2/SiC
blends) [76]. Other additives such as tetraethyl orthosilicate have been
applied as silica-forming impregnates (i.e. on Shuttle Orbiter), which seal
microstructural defects to limit oxidization [77]. These systems can protect
C/C from oxidation up to 1500-1600°C, but thin oxide coatings become
ineffective at higher temperatures due to melting, evaporation/active
oxidation of SiO(g), foaming, and/or visco-elastic erosion of the HfO2 or ZrO2
oxide scales containing high vapor pressure borosilicate phases. To satisfy
the needs of long-term reusable hypersonic service environments, both matrix
modification and deposition of coatings onto fibers and substrates are
required, but challenges persist. Annular matrix cracking and thermal
expansion mismatches between the C/C and coatings often lead to rupture during
thermal cycling and oxidative erosion, both of which are issues for sustained
high Mach number flight [78].

In an attempt to improve mechanical resilience and resistance to thermal
shock, advancements in composite materials design have focused on multi-scale
reinforcement strategies. Low dimensional micro/nanoscale reinforcements may
include nanoparticles (0D), carbon nanotubes/fibers (CNTs/CNFs), whiskers (1D,
e.g Si3N4, TaC, ZrC) or graphene (2D). Such additives serve to improve
properties at the fiber-matrix interface via grain refinement, debonding,
deformation, pullout, bridging, crack, and deflection mechanisms propagation
[76]. Higher dimensions, i.e., 2.5-D, 3-D, and 4-D, can be achieved by adding
short fiber to the resin matrix or waving and winding fiber strands into hoop
or braid-like structures [76].

The synergistic effects of complex multi-scale coating structural modification
and fiber-matrix interface optimization need to be further developed to expand
thermomechanical and erosion resistance properties for increasingly severe
service environments. The incorporation of modern computational approaches,
such as those developed in the national Materials Genome Initiative [79], and
integrated design of the fiber-matrix interface with unique combinations of
additives, may serve to improve materials performance. However, capabilities
for high-fidelity modeling of complex architectures are still limited [12,
80]. There remains a large gap in the comprehensive performance of
oxidation/ablation-resistant modified C/Cs in standard databases of material
properties insufficient for design and scalability to relevant components such
as aeroshells and leading edge [12].

###  3.3 Ultra-high Temperature and Refractory Ceramics for Hypersonics

Ultra-high temperature (UHT) and refractory ceramics are a developing class of
materials for leading edges due to their stability at high-temperatures [14].
UHTCs, encompassing carbides, nitrides, and borides of early transition metals
(Zr, Hf, Ti, Nb, Ta), possess high melting points (${\sim}$3000°C), tunable
densities (4.5-12.5 g/cm3), high thermal conductivities (40-120 W/mK),
moderate coefficient of thermal expansion (6.3 to 8.6 x 10-6/K), and high IR
spectral emissivity for passive radiative cooling [80, 81]. The complexity of
multiphase materials and their ability to survive extreme aerodynamic
conditions is shown in Figure 4. Challenges with respect to oxidation thermal
shock can be mitigated through tailoring the architecture and lead to
materials that rival metals in leading edge applications.

Among the UHTCs, ZrB2-SiC and HfB2-SiC have received the most attention
because they uniquely combine high thermal conductivity, specific strength
($\sigma$ $>$ 460 MPa at T = 2500°C, $\rho$ = 5.5 g/cm3 [82]), and superior
oxidation resistance ${\sim}$1650°C [78, 80, 83]. More recently, transition
metal carbides have garnered interested as components for nozzle throats,
divert/attitude control thrusters, and nozzle liners, where higher thermal and
mechanical loads are encountered [14]. Replacing HfB2 and ZrB2 with HfC or ZrC
can increase service above 2000°C. Still, the oxidation of refractory carbide
and boride ceramics containing SiC remains a significant challenge for
extended applications beyond ${\sim}$1600°C. Above these temperatures, active
oxidation generates gaseous oxidative products (i.e SiO(g) versus SiO2(s)),
which no longer provide protection against oxygen [14, 80].

A certain level of porosity can improve resistance to thermal shock and
thermal expansion mismatch. These pores help compensate for oxidation-induced
volume expansion, leading to the formation of a fully dense and cohesive
surface oxide scale but processing conditions and microstructure must be
carefully controlled. The structure-processing property relationships for
UHTCs is not well understood and additional information is needed to isolated
fundamental factors that control thermomechanical behavior of emerging
compositions [80]. For example, ZrB2-MoSi2 ceramics processed at elevated
temperatures are noted to have improved oxidation resistance [84].

The best oxidation performance for monolithic ceramics are obtained using hot
pressure assisted ceramics powder processing techniques, including: hot
pressing, hot isostatic pressing, and spark pressure sintering methods; these
methods facilitate porosity reduction and sintering while limiting coarsening
mechanisms [80]. Small UHTC grain sizes resist grain boundary fracturing
during oxidation and restrict molecular oxygen transport, minimizing the
disruptive effects of high-temperature martensitic transformations (phase
changes) of HfO2 and ZrO2 (Figure 4). Other recent approaches for improved
aerothermal resilience after evaporation of the protective B2O3 layer form
from diboride oxidation include additions of W, Mo, and Nb [85] and graphene
nanoplatelets reinforcement. These additions are indicated to suppress crack
formation, bursting, and oxide layer by up to 60%, while improving high heat-
dissipative abilities, important to surface resiliency when exposed to plasma
[86].

The high densities of UHTC materials, low thermal shock resistance, and low
fracture toughness impose additional physical limitations for bulk ceramics
[14]. Modern air-breathing hypersonic vehicles are extremely weight sensitive.
The high materials density (${\sim}$3-6x density compared to C/C) and poor
thermal shock resistance (1/5 of ACC-6, and half that of RCC and CVI C/SiC
1100°C [14]) of monolithic ceramics become a limiting factor for structural
components and dense segmented leading-edge inserts (Figure 3a, b). AAs a
result, the preferred instantiation of UHTCs is for emissive, anti-oxidative
coatings on Cf composites or refractory alloys. UHTC coatings can be improved
by adopting graded or layered compositions, enhancing bond strength by
structural integration, enhancing toughness and crack bridging via nanoscale
and micron-scale carbide fibers, and including emissivity enhancing dopants.
Compositional complexity can lead to further property improvements, where for
example TaHf-C has the highest recorded melting temperature [87].

Figure 3: Hypersonic wing leading edge designs and associated materials
microstructures processed under using different conditions showing materials
oxidation (a) Schematic drawings of wing leading-edge conceptual design using
a monolithic ceramic segmented edge. (b) Photographs of monolithic ZrB2/20
vol% SiC leading edges shown before and after arcjet testing in the H2 arc-jet
facility (see Figure 4) with failed ceramics due to oxidation and thermal
shock. (c) Failed and successful coated-C/C X-43 leading edges following arc-
jet for simulated flight conditions of Mach 10, 32 km altitude using 1475
W/cm2, 130 seconds. (d) HfB2-SiC UHTC nose cone subjected to a total 80
minutes of arc jet exposure at heat fluxes of 200 W/cm2. The sample formed an
oxide layer and a SiC depletion zone, which leaves behind a porous oxide
surface (e). (f) depicts SEM cross sections of HfB2-SiC or HfB2-SiC-TaSi2
materials formed via hot pressing and/or field-assisted sintering and with or
without TaSi2 additives. The images indicate how grain structure, oxide layer
formation, and SiC depletion is dramatically impacted by processing conditions
and the inclusion of tertiary phases. (g) SEM cross-section of a UHTCMC
incorporating Cf and high aspect ratio SiC. (Images from NASA and adapted from
[5, 14, 88].)

Whether for monolithic ceramic bodies or barrier coatings, processing
variables significantly affect materials properties and impose difficulties
for obtaining standardized performance data. Despite a significant body of
research examining both practical densification mechanisms and compositional
variation, limited information is reported on the kinetics of sintering and
densification of bulk ceramics. Many studies have reported the
thermomechanical properties of single ceramic compositions (strength,
hardness, elastic constants, thermal conductivity, and fracture toughness),
but limited information on the structure-processing-property relationship
stemming from first principles are understood [80, 89]. To expand the
utilization of refractory ceramics and UHTCs for hypersonic and extreme
environments, discovery/synthesis of new materials is needed, yet the
probability of undiscovered compounds remains low [80].

Further advancement of materials design these systems will likely parallel the
development of structural metal alloys. Simultaneous improvement of oxidation
resistance, creep at elevated temperatures, and transformation toughening will
require additions of secondary, ternary additions, and high-entropy
compositions that occupy uncharted areas of phase diagrams [12, 80, 90, 91].
Emerging work on high entropy, multi-component (ZrHfTi)C solid-solution
materials is being conducted, where for example high Hf-content is beneficial
for forming an amorphous oxycarbide layer enhancing initial oxidation
resistance, while an equiatomic ratio of metallic atoms increased high-
temperature phase stability [92, 87]. Multi-scale computational modeling of
UHTCs can aid in the development of high entropy materials by integrating ab-
initio (fundamental chemistry and electronic properties), atomistic
(thermomechanical properties), and continuum frameworks (mechanical
properties, thermomechanical analysis of microstructure) [88], [93, 94, 95,
96].

####  3.3.1 Ceramics-Matrix Composites

The poor thermal shock resistance and high densities of bulk UHTCs and
refractory ceramics can be overcome by the incorporation of ceramic fibers
(${\sim}$35-60 vol%) to create ceramics matrix composites (CMCs) [97, 98, 99].
On the other hand, the oxidation resistance of C/C can be improved by
replacing a carbonaceous with a ceramic matrix that forms a self-healing
glassy oxide passivation layer. For the latter, SiC was determined to be a
suitable substitute for the carbon matrix due to its high oxidation
temperature, thermal shock stability, and creep resistance [22]. The most
well-established carbon fiber-reinforced CMCs for hot structures are carbon
fiber-reinforced silicon carbide composite (C/SiC), and carbon fiber-
reinforced carbon-silicon carbide (C/C-SiC) composites [22]. While the active
oxidation of C/C starts at ${\sim}$500°C in air and becomes more significant
above 600°C, C/SiC and C-C/SiC can be stable up to 1600°C due to the formation
of a thin self-passivating SiO2 scale. Alternatively, environmentally stable
oxide/oxide ceramic composites have been produced to combat oxidation
experienced by non-oxide materials. These “Ox/Ox” CMCs incorporate
polycrystalline alumina or aluminosilicate fibers (e.g Nextel 610 or Nextel
720) into alumina, aluminosilicate, or SiOC matrixes and were initially
suggested for hypersonic thermal barrier materials. However, actual service
applications are limited to temperatures ${\sim}$1000-1200°C due to
degradation in tensile strength, stiffness, and creep [100].

The fracture behavior of damage-tolerant CMCs is dominated by the stiff
reinforcing Cf (or SiCf), where fiber-matrix debonding is associated with
frictional effects and crack deflection within porous or multilayer
interfaces. Fiber/matrix bonding using a coating with adapted interphases
(e.g. CVD pyrolytic carbon, silicon, $\beta$-SiC, BN, alumina) serves to: (1)
increase fiber-matrix bonding for enhanced mechanical properties (tensile
strength ${\sim}$350 MPa for CVI C/SiC); (2) protect carbonaceous fibers from
oxidative degradation at ${\sim}$450°C during crack formation when exposed to
an oxidative atmosphere; and (3) mitigate the severity of CTE mismatch between
Cf and the SiC matrix in the absence of matrix damage [14, 22]. Due to the
anisotropic nature of thermal expansion in composite materials and the CTE
mismatch between the SiC matrix and the Cf, these materials may be more prone
to cracking compared to SiCf/SiC CMCs. In recent years, melt silicon carbide
fiber-reinforced carbon-silicon carbide composites (SiC/SiC) can reach use
temperatures up to 1600°C. Yet, high temperature “sweating” of unreacted
silicon leaves processing improvements to be desired.

Practically, similar manufacturing techniques for C/C fabrication can be used
to infiltrate a carbon matrix with SiC or other refractory ceramic
compositions (including UHTCs). CMC fabrication techniques include chemical
vapor infiltration/deposition, PIP, reactive melt infiltration, slurry
infiltration, in-situ reaction, hot pressing, and powder pre-infiltration.
Several techniques can be combined to achieve multi-component compositions and
gradient/sandwich structures. Still, the fabrication of complex ultra-high-
temperature ceramic matrix composites (UHTCMCs) compositions remains of
significant interest in replacing C/C or C/SiC CMC materials for improved
thermomechanical capabilities. The advancement of UHTCMCs that incorporate
HfC, ZrC, TaC, HfB2, and ZrB2 matrix compositions will serve to greatly
benefit the development of propulsion platforms for hypersonic flights [98].

Carbon fiber-reinforced UHTC-matrix composites as shown in Figure 3g,
especially those containing HfC and ZrC, can resist oxidation at temperatures
above 2000°C under hypersonic flight conditions. However, CTE mismatches
between UHTC matrix phases are more prone to the microcracking during
processing and aerothermal heating that reduces strength [101]. The use of SiC
fiber will improve the structural properties, and oxidation resistance
(compared to Cf), and decrease the density UHTCMCs. HfC composites reinforced
with ${\sim}$15% short linear chopped SiC fibers showed strengths up to 370
MPa at room temperature, decreasing to 290 MPa by 2200°C in Ar [102]. These
properties far exceed the strengths reported for C/C which has demonstrated
strengths of ${\sim}$200 MPa at room temperature to 2200°C. The addition of up
40-60 vol% SiCf is suggested to further improve mechanical properties [103].

##  4 Advances in Computational Tools for Materials Development

###  4.1 Advances in Theory and Computational Tools

Many experimental characterizations remain costly and/or inaccessible at the
extreme conditions experienced by hypersonic vehicles [104, 105, 106], making
modeling/simulation efforts difficult to validate. Nonetheless, the lack of
experimental data can be an opportunity for theory and computation to provide
some insight. A history of simulation codes modeling thermal protection system
materials is presented in [106] and dates to the 1960s and includes software
like the CMA code by the Aerotherm Corporation [107, 108] and the FIAT code by
the NASA Ames Research Center [108], that incorporates internal energy
balance, decomposition equations, general surface energy balance boundary
conditions, and a thermochemical ablation model [106, 109]. Other simulations
have been performed modeling ablation in carbon-phenolic (charring) composites
[110, 111, 112], electromagnetic shielding at microwave frequencies [113],
water mass flow rate in transpiration (active) cooling systems [114], flow
fields and thermal behavior of solid samples, temperature-dependent fracture
toughness of particulate-reinforced UHTCs [115], and mechanical-thermoelectric
performance for multi-functional thermal protection systems [10]. High-
temperature thermal and elastic properties of high entropy borides have also
been modeled using molecular dynamics [116]. All of these methods contribute
to framing an understanding of the relevant vehicle systems level and
materials design criteria.

The challenge in performing experimental characterizations at conditions
relevant for hypersonic flight translates to a scarcity of empirical data.
Computational data is similarly limited: the complex mass and heat transfer
behavior at play spans several scales, and first-principles modeling is
largely disjointed across these scales and far too expensive for high-
throughput workflows. Any relevant data available exists in unstructured and
programmatically inaccessible formats. This makes it difficult to leverage
emerging artificial intelligence methods for accelerated screening. Solutions
are on the horizon: synchrotron x-ray computed microtomography was able to
fully resolve microcrack damage as cracks grew under load at temperatures up
to 1750°C [105] and high-throughput first-principles frameworks are becoming
capable of accurately modeling finite-temperature properties [95, 96, 117].
Disorder plays an ever-increasing role as the environment becomes more extreme
[91, 118], further complicating characterization and modeling efforts. The
prediction and optimization materials with useful properties will require an
understanding of the interplay between ultra-high-temperature phenomena, which
will be assisted by high-fidelity structured data, artificial intelligence and
solid thermodynamic-kinetic analysis.

To further the development of new and existing hypersonic materials, a
computational approach alone is insufficient. Real-world data is needed to
validate truly predictive material modeling and design. Furthermore, material
phenomenology and behavior at extreme temperatures are difficult to predict by
first principles alone, especially as material systems become more complex.

##  5 Role of integrated experiments and flight readiness pathways

Within the material design framework, computational models and experiments
serve vital and complementary roles. The intended application of the material
being designed largely drives the environmental loads. For example, a TPS
material designed for the exterior of a hypersonic vehicle will experience a
vastly different heating and chemical environment than the inside of a
scramjet combustor. These testing and modeling methodologies must take into
account the coupled thermal, structural, and chemical nature of the material-
environment interactions [6]. In addition, the intended reusability of the
material dictates the relevant timescales that must be examined. For typical
ballistic or boost-glide trajectories, flight times span tens of seconds to
tens of minutes. A single-use application may be able to ignore slower
phenomena that occur (e.g. creep), while a material designed for a reusable
application must account for the overall life of the vehicle.

Figure 4: Multi-scale modeling and testing framework for materials design and
flight testing. Length scales for both modeling and testing approaches span
many orders of magnitude. Smaller scale models can inform and be validated by
successively larger scale tests (Images adapted from [6, 110], [119, 120, 121,
122, 123])

Computational models for materials are traditionally broken down by domain
(material, fluid) and scale. At the larger scales, continuum codes rely on
finite-element analysis (FEA) or similar numerical methods to solve the
macroscopic governing equations over the physical domain. In the fluid domain,
the focus is usually to predict the environment that the material experiences,
such as the heat flux, pressure, and shear force. The field of computational
fluid dynamics (CFD) has expanded to include thermal and chemical
nonequilibrium, turbulence effects, surface reactions, and multiphase flow
[124, 125]. On the material side, thermal and structural analyses have
traditionally been considered separately. Material response tools evaluate the
thermal and chemical response, including any ablation and pyrolysis within the
material [126]. Thermo-structural tools model the combined aero-mechanical and
thermo-structural loads arising from thermal expansion and gradients within
the materials.

In recent years, there has been a push towards higher-fidelity tools and more
physics-based modeling, necessitating multi-scale modeling approaches to
describe the material at smaller and smaller scales. An example of this is the
focus on micro-scale modeling for porous TPS materials by NASA [127]. These
span meso-scale models describing the distinct phases that are present in a
material, down to atomistic models describing the fundamental material
interactions. A challenge is to bridge the gap between the various scales
using a truly physics-based approach, describing the relationship between
materials processing, microstructure, physical properties, and thermal,
structural, and chemical performance. These approaches can be generalized in
an Integrated Computational Materials Engineering (ICME) framework.

Flight tests are prohibitively expensive, and this has historically been a
major barrier in the development of hypersonic vehicles. Dedicated ground
tests provide an alternative way to emulate flight conditions in a controlled
environment. These include both aerothermal and structural tests. Although
aerothermal ground tests seek to recreate flight conditions as accurately as
possible, no facility is able to reproduce the exact flight conditions, and
instead seek to match two or more parameters [128]. Flight parameters of
interest include (but are not limited to) Reynolds number, Mach number, heat
flux, pressure, shear, temperature, chemical environment (e.g. dissociated
air), thermal shock, and exposure time. The freestream enthalpies experienced
during hypersonic flight are huge (in the MJ/kg), which presents another
challenge for ground test facilities. Arc jets, which produce realistic flight
enthalpies, shear, and pressure for up to several minutes, have been
considered the gold-standard of aerothermal testing for decades. However,
these tests are costly and time-consuming to prepare and perform. Other
facilities such as shock tubes/tunnels can match other flight parameters more
accurately, but only for sub-second exposure times, which limits the utility
of these facilities for thermal testing [129].

Similarly, the aim of conventional thermo-mechanical ground testing is to
validate some structural property, feature, or behavior given flight-realistic
mechanical and thermo-structural loads. For external TPS materials,
particularly composites, properties such as the interlaminar and shear stress
strengths are critical to the material performance and can vary greatly with
respect to material processing. Thermo-mechanical testing can span simple
property characterization, coupon-level, sub-scale, up to full vehicle-level
tests [130]. For materials designed for re-usable applications, the lifecycle
of the material under the thermo-structural loads is also critical to
evaluate, including any structural creep mechanisms.

For any flight vehicle, there is a large reliance on heritage materials, i.e.
materials that have flown previously. Thus, there has historically been a
large barrier to flight-test any new material system. In general, the
technology readiness level (TRL) and the manufacturing readiness level (MRL)
must be sufficiently high in order for a material to be considered flight-
ready, depending on the risk tolerance of the flight program. Increasing both
the TRL and MRL is a graduated process that requires extensive testing at
multiple scales, ranging from coupon-level and sub-scale to full-vehicle
tests.

Since the opportunities for flight tests are so limited, the conventional
approach to developing a material system is to conduct as much learning
beforehand via computational models and ground testing, and to treat the
flight test as a final verification of the material/vehicle system. In this
paradigm, the goal of integrated testing is to provide data to inform and
validate models, as well as characterize phenomena that are difficult or
impossible to model. These phenomena can be broadly categorized by the
physical length scales, illustrated in Figure 4. At the macroscopic level, the
fundamental chemical, thermal, and structural material properties need to be
characterized for any computational model. Some statistical basis and
understanding of the variance is generally required as well. The
microstructure of a material is inherently linked to these thermo-physical
properties. In turn, the microstructure and material architecture are strongly
tied to the manufacturing process, especially for composite materials such as
carbon-carbon [131]. At the largest scales, integrated testing is also
critical to understand how the material behaves under flight conditions
combining oxidation/ablation, shear, and thermo-structural loads.

Within a multi-scale modeling framework, there exists a symbiotic relationship
between modeling and experiments. As models become more and more detailed and
span many different length scales, experiments, and diagnostic tools are
needed to provide data for the validation of these models. Concurrently,
models are used to bridge the gap between ground test and flight environments.
In fact, validation of models that occur at smaller length scales is often
achieved by testing at successively larger scales, e.g. a meso-scale material
model may be validated by coupon-level testing in Figure 4. The development of
improved modeling methodologies must be accompanied by test data to validate
the methods and models.

##  6 Conclusion

This perspective offers valuable insights into the critical challenges that
researchers must tackle in order to advance materials technologies for
hypersonic environments. By delving into key sub-systems of hypersonic
vehicles — primary structures, thermal protection, and propulsion systems — we
elucidate the essential material properties required to withstand the severe
thermal and oxidative conditions inherent in such settings. Emerging materials
design strategies are actively being formulated to enhance material properties
and address pivotal challenges within each principal materials group.

1\. Refractory metals, while possessing noteworthy attributes, exhibit
limitations in oxidation resistance and strength at elevated temperatures,
impacting their resilience in high heat flux environments. Multi-principal
element “high entropy” alloys could provide significant benefits over legacy
refractory alloys by decreasing density and improving oxidation kinetics.
However, there remains a gap in our understanding of thermophysical properties
in operation-relevant conditions. Leveraging rapidly evolving manufacturing
methods alongside high-throughput characterization and machine learning
algorithms holds promise for expeditiously uncovering novel compositions.
Integrating these approaches could pave the way for novel advanced cooling
technologies and resilient metallic structural components.

2\. Composites display promising strength-to-weight ratios and elevated
temperature tolerance in inert atmospheres. Nevertheless, uncoated materials
are susceptible to significant oxidation and erosion when subjected to extreme
temperatures. Enhancing performance involves the application of high-
temperature emissive protective coatings and modifying carbonaceous matrices.
Challenges arise from the disparity in material properties between coatings
and matrices, as well as matrices and fibers, often leading to premature
failure. The inherent anisotropic properties of existing fiber-reinforced
materials underscore the importance of advancements in multi-scale
reinforcement strategies and higher-dimensional materials, which represent
compelling avenues for future development.

3\. Refractory ceramics and UHTCs are characterized by exceedingly high
melting points and thermal conductivities, but present challenges due to their
low thermal shock resistance and density when used as monolithic components.
Their optimal application lies in thermal barrier coatings and ceramics matrix
composites. Notably, transition metal carbides incorporating SiC have gained
attention for components exposed to elevated thermal and mechanical loads.
However, the oxidation of refractory compositions containing SiC remains a
formidable hurdle for extended applications beyond approximately 1600°C.
Further research is imperative to exploring ceramics’ structure-processing
property relationships, transformation toughening, oxidation enhancement
through high entropy compositions, and the intricacies of fiber/matrix bonding
for ultrahigh-temperature ceramics.

The adoption of expanded materials design frameworks - from the atomistic to
the macroscopic - is needed to create the next generation of resilient
materials systems that can survive in complex hypersonic environments.

## Acknowldegements

S.C. acknowledges support by DoD (N00014-21-1-2515) and NSF (NRT-HDR
DGE-2022040). This material is based upon work supported by the Air Force
Office of Scientific Research under award number FA9550-22-1-0221. I.M and
C.O. acknowledge support from the Air force Office of Scientific Research, and
discussions with Michael Brupbacher and Tom Magee. T.M.P. acknowledges the
support of a Department of Defense Vannevar Bush Fellowship Grant ONR
N00014-18-1-3031.

## Competing Interests

The authors declare no competing interests.

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