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Low-density ablative heat shield fabricationLow-density ablative heat shield fabrication description/claimsThe Patent Description & Claims data below is from USPTO Patent Application 20070224407, Low-density ablative heat shield fabrication. Brief Patent Description - Full Patent Description - Patent Application Claims
[0001]This application claims priority to U.S. Provisional Patent Application Ser. No. 60/785,930, titled "Low-Density Ablative Heat Shield Fabrication," filed Mar. 24, 2006 and incorporated herein by reference. [0002]The invention described herein was made by nongovernment employees, whose contributions were done in the performance of work under a NASA contract, and is subject to the provisions of Public Law 96-517 (35 U.S.C. 202). This invention was made with Government support under contract NNA04BC25C awarded by NASA. The Government has certain rights in this invention. BACKGROUND OF THE INVENTION [0003]1. Field of the Invention [0004]The present invention relates to the space program, and more specifically, it relates to processes for making ablative heat shields that provide protection to spacecraft during the severe heating conditions of atmospheric entry. [0005]2. Description of Related Art [0006]The making of large ablative heat shields that provide protection to spacecraft during the severe heating conditions of atmospheric entry is a difficult problem. Limitations due to inherent physical properties and the inability to fabricate large assemblies due to processing and/or machining requirements have prevented the use of the most promising materials in spacecraft design in several instances. In the specific cases of low density and mid-density ablative materials, difficulties in controlling the impregnation of fiber matrix materials with polymeric resins, and difficulty in processing larger fiber matrix substrates within density specifications to achieve uniform material properties have limited the size that billets or blocks of these materials can be made. These limitations have required the use of less capable materials or the use of fabrication and assembly methods that lead to complex and costly final products. A new fabrication process is desired that would overcome several of the problems previously encountered in the making of large, one-piece heat shields. Such process should be applicable to a wide variety of refractory fiber matrix materials or refractory porous substrates and polymeric resins that are known to form efficient ablative heat shield materials. [0007]The type of thermal protection system (TPS) that best protects against high heat flux is the ablative heat shield. The ablative heat shield functions by the energy-absorbing thermal degradation of a polymeric component resulting in the production of a char layer and of gaseous products through a process known as pyrolysis; the absorption of additional energy as these gases flow through the porous degraded material to the heat shield surface; the possible phase change of components from solid to liquid to gaseous, or from solid to gaseous states; and the reduction of the convective heat flux to the heat shield surface by gaseous products as they leave the surface by the thickening and cooling of the boundary layer in a process called blowing. The kinetics and products of the pyrolysis process can be measured in real time using thermogravimetric analysis, so that the ablative performance can be evaluated. Ablation can also provide blockage against radiative heat flux by the introduction of spectrally absorbing gaseous pyrolysis products into the boundary and shock layers in front of an entry spacecraft. Radiative heat flux blockage was the primary thermal protection mechanism of the Galileo Probe TPS material (carbon phenolic). Thermal protection also can be enhanced in some TPS materials through coking. Coking is the process of forming and depositing solid carbon within the char layer of the TPS, resulting in a localized density increase within the char. [0008]The thermal conductivity of a TPS material is proportional to the material's density and dependent on fiber orientation in fiberous substrates. Carbon phenolic is a very effective ablative material but also has high density and resulting high conductivity which is undesirable. If the heat flux experienced by an entry vehicle is insufficient to cause pyrolysis then the TPS material's conductivity could allow heat flux conduction into the TPS attachment and spacecraft structures, thus leading to TPS failure. Consequently for entry trajectories causing lower heat flux, higher density TPS materials such as carbon phenolic are inappropriate and lower density TPS materials may be better design choices. [0009]The concepts presented herein are applicable to a broad range of ablative TPS constructed from refractory fibrous matrix materials or refractory porous substrates and polymeric impregnation resins. An example used herein for purposes of illustration is Phenolic Impregnated Carbon Ablator (PICA) TPS. This material was developed by NASA Ames Research Center and was the primary TPS material for the Stardust Sample Return Capsule aeroshell. Because the Stardust spacecraft was the fastest man-made object to reenter Earth's atmosphere (.about.12.4 km/sec, .about.28,000 mph relative velocity at 135 km altitude), PICA was an enabling technology for the Stardust mission. (For reference, the Stardust reentry was faster than the Apollo Mission capsules and 70% faster than the reentry velocity of the Shuttle.) PICA is a modern TPS material that has the advantages of low density (much lighter than carbon phenolic) coupled with efficient ablative capability at high heat flux. Stardust's heat shield (0.81 m base diameter) was manufactured from a single monolithic piece sized to withstand a nominal peak heating rate of 1200 W/cm.sup.2. PICA is a good choice for ablative applications such as high-peak-heating conditions found on sample-return missions or lunar-return missions. PICA's thermal conductivity is lower than other high-heat-flux ablative materials, such as conventional carbon phenolics. [0010]Small blocks or tiles of both ablative and insulative thermal protection materials have been used in heat shield applications (e.g., Space shuttle Orbiter) because of thermal expansion and processing size limitations. This requires that each tile be machined to a finished size prior to being individually attached to spacecraft structure in a complex and costly operation and, in the case of the Space Shuttle, fabric filler materials were required to fill gaps between tiles that exist before entry heating. In other applications (e.g., Genesis sample return capsule) the choice of an ablative heat shield material was constrained by the inability to make the required one-piece heat shield. For the Apollo capsule heat shield, the required one-piece ablative heat shield was fabricated by injecting an epoxy resin into a phenolic honeycomb in a costly, complex, and hard to control process. SUMMARY OF THE INVENTION [0011]It is an object of the present invention to provide methods for making ablative heat shields. [0012]It is another object to provide a solid, one-piece, monolithic ablative heat shield. [0013]Sill another object is to provide an ablative heat shield of modular pieces of a fiber matrix material or refractory porous substrate material that has been cross-linked together to form a one-piece assembly. [0014]These and other objects will be apparent based on the disclosure herein. [0015]The fabrication of large (larger than 1 meter diameter) spacecraft heat shields as one-piece assemblies using low-density ablative thermal protection materials (TPS) formed from rigid substrates has been constrained by limits on available component matrix material sizes and processing requirements. Methods are provided that allow large uni-piece heat shields to be fabricated for use on future space vehicles that require protection from atmospheric entry heating at severe conditions. These fabrication methods provide such large assemblies by building a heat shield assembly from modular pieces (blocks) formed by conventional ablative TPS impregnation and processing methods and limitations. Once the full-size heat shield is assembled from the modular blocks, appropriate heat treatment is used to bond the individual blocks together by facilitating polymeric cross-linking of impregnant material within and/or between each block. This provides a structurally integral one-piece heat shield assembly that can be further machined to final dimensions and attached directly to spacecraft structure or a carrier panel separately attached to the spacecraft. BRIEF DESCRIPTION OF THE DRAWINGS [0016]The accompanying drawings, which are incorporated into and form a part of the disclosure, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. [0017]FIG. 1 shows an exemplary individual refractory fiber matrix or refractory porous substrate block. [0018]FIG. 2 illustrates the impregnation of a refractory fiber matrix or refractory porous substrate block with resin from all sides. [0019]FIG. 3 shows the assembly on a mandrel of a section of a heat shield having a plurality of fiber matrix or refractory porous substrate blocks. [0020]FIG. 4 shows an oven containing an exemplary heat shield assembly. [0021]FIG. 5 shows a one-piece bonded heat shield attached to a spacecraft structure. Continue reading about Low-density ablative heat shield fabrication... 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