the present invention will be described in terms of processes for producing components that contain ceramic - based materials , and particularly the incorporation of one or more tip caps that can be used to close one or more internal cavities of a component formed of a ceramic - based material , preferably a cmc material . while various applications are foreseeable and possible , applications of particular interest include high temperature applications , for example , turbine components of gas turbines , including land - based and aircraft gas turbine engines . the cmc turbine blade 10 of fig1 will serve as an example in the following discussion . while the invention is applicable to a wide variety of ceramic - based materials , ceramic - based materials of particular interest to the invention are believed to be cmc materials containing silicon , such as cmc &# 39 ; s containing silicon carbide as the reinforcement and / or matrix material , for example , continuous silicon carbide fibers in a matrix of silicon carbide . however , other ceramic - based materials are also within the scope of the invention , nonlimiting examples of which include fibers and reinforcement materials formed of titanium carbide ( tic ), silicon nitride ( si 3 n 4 ), and / or alumina ( al 2 o 3 ). as known in the art , the airfoil 18 of the blade 10 is an excellent candidate for being produced from a ceramic - based material , and especially a cmc material , because it is directly exposed to the hot combustion gases within the turbine section of a turbomachine , and has a generally linear geometry . on the other hand , the incorporation of an internal cooling cavity , cooling holes and a tip cap results in a more complex geometry , in the sense that the airfoil 18 has a generally linear geometry along its dominant span - wise axis , whereas a tip cap would be a geometric feature oriented transverse to the span - wise direction of the blade 10 . furthermore , the off - axis geometry of a tip cap would be subjected to high mechanical loading during operation of the engine , and therefore require structural interface capabilities that pose substantial challenges to designing , manufacturing and integration with a blade formed of a cmc material . the present invention provides a process for taking advantage of the high - temperature capabilities of cmc materials , while addressing the difficulties of integrating a tip cap into an airfoil component formed of a cmc material . in particular , a preferred aspect of the present invention is the ability to produce a tip cap from plies , and to fully integrate the tip cap as part of an airfoil formed from plies utilizing a lay - up process . fig2 a , 2 b , 3 a , 3 b , and 4 - 7 schematically represent views of the tip region of the blade 10 of fig1 , and represent the integration of tip caps 22 from plies 24 in accordance with various non - limiting embodiments of the present invention . the airfoil 18 and tip cap 22 can be fabricated from ceramic - based materials produced using known processes , for example , with the use of prepregs . as a particular example , the airfoil 18 and its cap 22 can each be fabricated using a prepreg melt - infiltration ( mi ) process of a type previously described , wherein multiple prepregs are formed to contain one or more desired reinforcement materials and a precursor of the cmc matrix material , as well as one or more binders . the prepregs undergo lay - up , are debulked and cured while subjected to elevated pressures and temperatures , and may undergo various other processing steps to form a laminate preform . thereafter , the laminate preform is heated ( fired ) in a vacuum or an inert atmosphere to decompose the binders and produce a porous preform , which then preferably undergoes melt infiltration . if the cmc material comprises a silicon carbide reinforcement material in a ceramic matrix of silicon carbide ( a sic / sic cmc material ), molten silicon or a silicon alloy is typically used to infiltrate and fill the porosity and , in preferred embodiments , react with a carbon constituent ( carbon , carbon source , or carbon char ) within the matrix to form silicon carbide . however , it will be apparent from the following discussion that the invention also applies to other types and combinations of ceramic and cmc materials . furthermore , it is foreseeable that the unitary airfoil 18 and cap 22 could be fabricated with the use of materials other than prepregs , for example , cloth - reinforced cmcs , such as chemical vapor infiltrated ( cvi ) sic reinforced with carbon fiber cloth ( c / sic ), cvi / slurry cast / melt infiltrated sic / sic , and cvi sic reinforced with sic cloth . polymer infiltration and pyrolysis ( pip ) processes can also be used to deposit the matrix into a cloth reinforced preform , in which case a sic or carbon cloth can be used . according to a preferred aspect of the invention , the fabrication of the tip cap 22 entails steps intended to fully integrate the tip cap 22 into the linear geometry of the airfoil 18 . fig2 a and 2b represent an example of a blade tip region of the blade airfoil 18 during the fabrication of the tip cap 22 , which according to a preferred aspect of the invention can be entirely formed of a cmc material and produced by a cmc process as described above . as represented in fig2 a and 2b , the tip cap 22 is fabricated from multiple prepreg plies 24 . fig2 a and 2b represent the plies 24 as disposed within a cavity 30 defined by and between the convex ( suction ) and concave ( pressure ) walls 26 and 28 of the airfoil 18 , which as represented in fig2 b are also fabricated from multiple plies 34 . fig2 a and 2b further represent the plies 24 as extending in the chord - wise direction of the airfoil 18 . as previously noted , each of the plies 24 and 34 preferably contains a desired reinforcement material and a suitable precursor of a desired ceramic matrix material . the reinforcement material and ceramic matrix material of the tip cap plies 24 are preferably , though not necessarily , the same as those for the airfoil plies 34 . it should be appreciated that various numbers of plies 24 could be incorporated into the construction of the tip cap 22 of the blade 10 . to build up a suitable thickness for the tip cap 22 that completely fills the portion of the cavity 30 within the blade tip region of the airfoil 18 , most of the plies 24 are represented as having roughly equal chord - wise lengths ( fig2 a ). in addition , most of the plies 24 are represented as having roughly equal span - wise lengths ( fig2 b ), such that the tip cap 22 is substantially flush with the end of each wall 26 and 28 of the airfoil 18 . however , certain plies 24 are represented as being intentionally shorter than others in the chord - wise direction ( fig2 a ) to accommodate a varying width of the cavity 30 , and certain plies 24 are also represented as being intentionally shorter than others in the span - wise direction ( fig2 b ). it should be understood that the lengths and widths of the plies 24 can vary , for example , as a result of increasing or decreasing in length and / or width to yield what may be referred to as a stepped formation . accordingly , shapes and sizes of the plies 24 other than the particular shapes and sizes represented in fig2 are foreseeable and within the scope of the invention . according to a preferred aspect of the invention , shorter plies 24 in the span - wise direction are utilized to create a wedge - shaped profile 32 at the radially - inward end of the tip cap 22 . as seen in fig2 b , the wedge - shaped profile 32 of the cap 22 engages complementary notches 33 formed in the interior surfaces of the convex and concave walls 26 and 28 of the airfoil 18 . the wedge - shaped profile 32 of the tip cap 22 and the notch 33 within the airfoil cavity 30 cooperate to interlock the tip cap 22 within the cavity 30 , particularly after the plies 24 of the tip cap 22 are fired and melt infiltrated , enabling the tip cap 22 to withstand high centrifugal forces that exist during the operation of the blade 10 . to complete the manufacturing of the blade 10 and its tip cap 22 , the laid - up prepreg plies 24 and 34 are preferably debulked prior to undergoing curing , followed by firing during which binders are burned - off and a ceramic precursor is converted to the desired ceramic matrix material for the reinforcement material . suitable debulking , curing and firing processes , as well as any additional processes necessary to achieve the final desired shape and properties of the blade 10 , are known in the art and therefore will not be described further . whereas the plies 24 of the tip cap 22 are represented in fig2 a and 2b as being oriented in the span - wise and chord - wise directions of the airfoil 18 , fig3 a and 3b represent another embodiment in which the plies 24 are oriented in the thickness - wise and chord - wise directions of the airfoil 18 . aside from the difference in orientation of the plies 24 , the tip cap 22 can be fabricated and interlocked with the airfoil 18 in essentially the same manner as described for the embodiment of fig2 a and 2b . fig4 through 7 represent additional configurations for tip caps 22 that can be fabricated in accordance with various aspects of the invention . whereas in fig2 a , 2 b , 3 a and 3 b , a single tip cap 22 is represented as filling a single cavity 30 in an airfoil 18 , fig4 represents a single tip cap 22 as closing multiple cavities ( cooling passages ) 30 within an airfoil 18 , fig5 represents separate tip caps 22 as individually filling and closing each of multiple cavities ( cooling passages ) 30 within an airfoil 18 , and fig6 represents the incorporation of pins 38 to help secure a tip cap 22 used to close multiple cavities ( cooling passages ) 30 within an airfoil 18 . in the embodiment of fig4 , the tip cap 22 is fabricated on top of all blade cavities 30 to seal off cooling passage air flow at the blade tip . plies 34 of the airfoil walls 26 and 28 are represented in fig4 as wrapped around the cooling cavities 30 and the plies 24 of the tip cap 22 , and the tip cap 22 is bonded to the interior surfaces of the airfoil walls 26 and 28 during curing of both the airfoil 18 and tip cap 22 . in fig5 , each cooling cavity 30 is individually sealed off by a separate tip cap 24 . the reinforced embodiment of fig6 is intended to increase the aerodynamic and centrifugal loading capability of the tip cap 22 . in fig6 , a single tip cap 22 ( shown in the upper blade tip of fig6 , but omitted in the lower image to reveal the cavities 30 and pins 38 ) is represented as closing multiple cavities ( cooling passages ) 30 within an airfoil 18 , though it should be understood that separate tip caps 22 that individually fill and close multiple cavities 30 could also be reinforced in the same or similar manner . the embodiment represented in fig6 entails additional steps between the lamination and cure processes . in a particular example , holes ( not shown ) are drilled through the airfoil walls 26 and 28 and tip cap 22 , for example , using an ultrasonic needling process , and then the pins 38 , for example , formed of prepregs , are inserted into the holes to create an interlocking connection between the airfoil walls 26 and 28 and tip cap 22 following curing . finally , fig7 represents the incorporation of holes 40 in the tip cap 22 that are fluidically connected to one or more cavities ( not shown ) within the airfoil 18 . as known in the art of blades formed of metallic materials , tip cap purge holes have been utilized to regulate internal cavity pressures within blades , which in turn determines the cooling air flow rates through the cooling passages and cooling holes of the blades . holes 40 of the type represented in fig7 can be formed by drilling after melt infiltration , when the cmc plies 24 of the tip cap 22 have been fully processed . drilling techniques that can be used include electrodischarge machining ( edm ), ultrasonic machining , or another traditional machining technique . while the invention has been described in terms of specific embodiments , it is apparent that other forms could be adopted by one skilled in the art . for example , the number of tip cap plies 24 required to close a particular cavity 30 of a blade 10 can be modified , for example , by increasing the thickness of either or both airfoil walls 26 and 28 . furthermore , the composition of the tip cap 22 can vary from that described above , for example , discontinuous ( chopped ) fiber reinforcement materials could be used in place of continuous fiber reinforcement materials , and in doing so could potentially eliminate the need for multiple laminated plies 24 to form the tip cap 22 . in addition , welding or fusing techniques could be adapted to bond the tip cap 22 to the airfoil 18 after melt infiltration , avoiding the process of forming the tip cap 22 as part of the initial composite laminate . therefore , the scope of the invention is to be limited only by the following claims .