High temperature resistant airfoil apparatus for a hypersonic space vehicle

Abstract

A ruddervator for an aerospacecraft including a monolithic, one-piece, oxide/oxide-based ceramic matrix composite (Oxide-CMC) shell having a hollowed interior area. A graphite composite structural member is inserted into the hollowed interior area and bonded thereto. The Oxide-CMC shell is comprised of one or more plies of Oxide-CMC fabric which are fused over a thick substrate of rigid ceramic foam insulation to form the monolithic shell. An outer mold line ply of the Oxide-CMC shell is further infused with a high-emissivity coating such as reaction-cured glass (RCG) to provide plasma heating re-radiation outward to reduce internal temperatures in the ruddervator. A torque box transition structure is secured to the graphite composite structural member to interface the ruddervator to a fuselage actuator spindle of the aerospacecraft. The ruddervator is reusable and formed from a reduced number of independent component parts, and weighs less than previously developed ruddervators. The manufacturing cost of the ruddervator is reduced by utilizing the reduced tooling complexities of Oxide-CMC over CMC fabrication processes.

What is claimed is:  
     
         1 . A high temperature resistant airfoil apparatus for a hypersonic space vehicle, comprising: 
 a temperature resistant ceramic matrix composite shell having an opening at one end and a hollowed interior area;    a structural member inserted into said hollowed interior area and secured to an interior surface of said shell to form a structurally rigid airfoil assembly; and    a transition structure secured to said structural member for interfacing said airfoil apparatus to a control element of said space vehicle to permit said airfoil apparatus to be controlled by said control element.    
     
     
         2 . The apparatus of  claim 1 , wherein said shell comprises an oxide/oxide-based ceramic matrix composite (Oxide-CMC) shell.  
     
     
         3 . The apparatus of  claim 2 , wherein a top outer mold line (OML) ply of said Oxide-CMC shell is infused with a high-emissivity coating.  
     
     
         4 . The apparatus of  claim 1 , wherein said structural element comprises a graphite composite structural member having a graphite composite facesheet and a honeycomb core element.  
     
     
         5 . The apparatus of  claim 1 , further comprising a plurality of thermal barrier tiles secured over said transition structure.  
     
     
         6 . The apparatus of  claim 2 , wherein said shell comprises an Oxide-CMC fabric fused to an outer surface of a rigid ceramic foam insulation member; and 
 wherein said insulation member is RTV bonded to said structural member.    
     
     
         7 . The apparatus of  claim 1 , wherein said structural member is wedge-shaped when viewed chord-wise to help eliminate air being trapped within said hollowed interior area of said shell as said structural member is inserted into said hollowed interior area during manufacturing of said airfoil.  
     
     
         8 . The apparatus of  claim 1 , wherein said structural member is wedge-shaped when viewed from one side thereof to help eliminate air being trapped within said hollowed interior area of said shell as said structural member is inserted therein during manufacturing of said airfoil.  
     
     
         9 . The apparatus of  claim 1 , wherein a lower end of said shell comprises a serrated edge to minimize high temperature flow to an aft portion of said airfoil assembly.  
     
     
         10 . A high temperature resistant ruddervator apparatus for a hypersonic space vehicle, comprising: 
 a one piece, temperature resistant oxide/oxide-based ceramic matrix composite (Oxide-CMC) shell having an opening at one end and a hollowed interior area, said Oxide-CMC shell comprising an Oxide-CMC fabric fused to a rigid ceramic foam insulation member;    a structural member inserted into said hollowed interior area of said Oxide-CMC shell and bonded to an interior surface of said Oxide-CMC shell to form a structurally rigid ruddervator assembly; and    a transition structure secured to said structural member for interfacing said ruddervator assembly to a control element of said space vehicle to permit said ruddervator assembly to be controlled by said control element.    
     
     
         11 . The apparatus of  claim 10 , wherein said structural member comprises a graphite composite structural member having a graphite/epoxy facesheet secured to a honeycomb core element.  
     
     
         12 . The apparatus of  claim 10 , wherein said Oxide-CMC fabric is comprised of a plurality of plies of Oxide-CMC fabric fused to an outer surface of said rigid ceramic foam insulation member.  
     
     
         13 . The apparatus of  claim 10 , wherein an outer surface of said Oxide-CMC fabric is infused with a high emissivity coating to reduce internal temperatures experienced by said rigid ceramic foam insulation and said structural member.  
     
     
         14 . The apparatus of  claim 13 , wherein said high emissivity coating comprises reaction cured glass (RCG).  
     
     
         15 . A method of manufacturing a ruddervator for a hypersonic space vehicle, comprising the steps of: 
 forming a one piece shell comprised of oxide/oxide-based ceramic matrix composite (Oxide-CMC) material, said Oxide-CMC shell being open at one end and having a hollowed interior area;    inserting a structural member into said one end and into said hollowed interior area of said Oxide-CMC shell, said structural member being shaped generally in accordance with a shape of said hollowed interior area such that said structural member fits snugly with said hollowed interior area; and    securing an interior surface of said Oxide-CMC shell to an outer surface of said structural member.    
     
     
         16 . The method of  claim 15 , further comprising the step of securing a transition structure to a lower end of said structural member, said transition structure being adapted to be secured to a control element of said hypersonic space vehicle.  
     
     
         17 . The method of  claim 15 , further comprising the step of forming said Oxide-CMC shell from an Oxide-CMC fabric fused over a rigid ceramic foam insulation substrate.  
     
     
         18 . The method of  claim 15 , further comprising the steps of: 
 forming said Oxide-CMC shell from an Oxide-CMC fabric and a rigid ceramic foam insulation substrate, wherein said foam insulation substrate comprises said hollowed interior area; and    using an RTV bonding process to secure said foam insulation substrate to said structural member.    
     
     
         19 . The method of  claim 15 , further comprising the step of forming said hollowed interior area and said structural member each with a wedge shape to reduce the chance of air being trapped inside said hollowed interior area during insertion of said structural member therein.