«DESIGN OF AN INTEGRAL THERMAL PROTECTION SYSTEM FOR FUTURE SPACE VEHICLES By SATISH KUMAR BAPANAPALLI A DISSERTATION PRESENTED TO THE GRADUATE SCHOOL ...»
Different insulation materials originally used on the Space Shuttle include reinforced carbon-carbon (RCC), two types of ceramic reusable surface insulation (RSI) tiles, and a limited amount of non-reusable ablative material . RCC has a reuse temperature of 2900 F  and is used at the stagnation regions such as nose cap, wing and tail leading edges, where the heating rates are extremely high. The ceramic RSI tiles cover a major portion of the Shuttles and are of two varieties high and low temperature reusable surface insulation (HRSI, LRSI). In a few areas the HRSI tiles have later been replaced by stronger fibrous refractory composite insulation (FRCI) [11: p.30]. Most of the LRSI tiles on the leeward side were also replaced by an advanced flexible reusable surface insulation, AFRSI [11: p.31]. Figure 2-3 shows the TPS distribution on the Space Shuttle Orbiter.
Figure 2-3. Distribution of TPS on the Space Shuttle Orbiter. (Obtained from Reference ) The biggest disadvantage of RSI tiles is that they are highly brittle . They cannot withstand even the slightest of impact loads and also have a very low strain to failure. Therefore, they cannot be used as structural elements. The tiles are bonded to a sub-structure made of aluminum alloys. The aluminum substructure bears all the mechanical loads. The ceramic tiles also have a low coefficient of thermal expansion when compared to aluminum. If the tiles were bonded directly to aluminum, any thermal or mechanical strains in the substructure may cause considerable tensile strain in the tiles, which could lead to cracking. In order to avoid this, strain isolation pads (SIPs), which have a very low shear and extensional modulli, are used to separate the tiles from the aluminum substructure, Figure 2-4. Further, the tiles are also made of dimensions 6 inches or less in length and width . Gaps are left between tile edges to allow for relative motion when the aluminum skin expands or contracts, Figure 2-4. The tiles also require silicon polymer water proofing before every flight. Thus, even though the Space Shuttle TPS provides excellent thermal protection for the vehicle, it is not robust and is highly susceptible to catastrophic damage. Tile maintenance between flights requires approximately 40,000 man-hours [11: p.2] and this has increased the turnaround times considerably when compared to the original estimates . These factors have contributed to the enormous launch costs of Space Shuttle flights besides making it structurally not very robust.
The X33 was an experimental vehicle aimed at developing a completely reusable launch vehicle (RLV), the VentureStar RLV. The program began in mid 1990s and was cancelled in
2001. It was proposed to be a single-stage-to-orbit (SSTO) launch vehicle with a Boeing Linear Spike engine. The concept was similar to the Space Shuttle’s vertical-take-off-horizontallanding. However, it had no external boosters or fuel tanks. All the fuel tanks were embedded in the vehicle. This meant that the size of the VentureStar would be much larger than the Space Shuttle, which in turn implies a much larger surface area to be lined with TPS. The VentureStar program was stopped in 2001 due to various technological challenges, and cost and schedule overruns. However, there were some useful technologies developed for this program and one of them is the TPS that will be discussed here.
Figure 2-4. Schematic of tile attachment to aluminum structure.
One of the most important goals of TPS development for the VentureStar program was to overcome the numerous problems relating to the ceramic tiles on the Space Shuttle and to provide a robust TPS. A TPS of fully metallic construction was proposed in the 1990s [12–15].
This TPS was christened ARMOR TPS, where ARMOR stands for adaptable, robust, metallic, operable, and reusable . Major portion of the space inside the VentureStar was to be occupied by the cryogenic fuel tanks. Therefore, the ARMOR TPS was designed to be attached to the cryogenic tank structure. Figure 2-5 schematically illustrates the design of the ARMOR TPS.
The top surface was made of a metallic honeycomb sandwich panel, which can withstand small aerodynamic pressure loads on the TPS. The primary job of this panel was to function as a hot structure and re-radiate most of the heat incident on the TPS. The top face sheet of the honeycomb panels was extended to partially overlap the adjacent panels to seal the panel-topanel gaps from ingress of hot gases during reentry. The top panel was insulated from the bottom using a thick layer of high temperature alumina fiber insulation, Saffil®. The insulation material was contained on the sides by bulged, compliant sides made of thin gauge metal foil and at the bottom by a thin gauge metal foil backing. The bulged, compliant sides also helped in blocking the radiation between panel-to-panel gaps. The insulation material was vented to the vehicle internal pressure at the bottom, through the titanium foil, so that the aerodynamic pressure loads were borne by the top honeycomb panel. A picture frame made of thin gauge titanium box beams formed the TPS boundary on the bottom of the panel. This frame can be used to attach the TPS to the structure. The pressure loads on the top surface were transferred to the titanium box beam through four corner support brackets. The support brackets had low bending stiffness to allow the top face sheet to expand almost freely and thus prevent large thermal stresses from developing in the top panel. The ARMOR TPS panels were mounted on a TPS Support Structure (TPSS). The TPSS connects the TPS panels to the cryogenic tank structure. The top honeycomb panel and the compliant sides were made of titanium alloy, Ti 1100, for temperatures up to 1100 F. For temperatures ranging from 1100 to 2000 F the top panel was made of Inconel 617 and the compliant sides were made of Inconel 718 .
ARMOR TPS combines the two concepts of hot structure and insulated structure to design a lightweight and robust TPS. In fact, it also acts as a heat sink and absorbs a portion of the heat.
Since it is made of fully metallic construction, it can be expected to be damage and impact resistant to a certain extent, something that was entirely missing in the Space Shuttle TPS. It can also withstand light aerodynamic pressure loads and acoustic loads produced by the propulsion system. However, it has no provision to withstand in-plane loads and also cannot withstand large aerodynamic pressure loads.
C Figure 2-5. ARMOR TPS construction. A) Cross-sectional view. B) ARMOR TPS mounted on TPSS and cryongenic tank wall. C) Full-view. (Figures obtained from Reference ) Other TPS designs studied for reusable launch vehicles were mainly metallic TPS
concepts. Prepackaged superalloy honeycomb panel was a predecessor to ARMOR TPS [12, 7:
p.54] and with similar design features, as shown in Figure 2-6. It was developed and tested for the X33 during the 1990s. The top surface was made of Inconel honeycomb sandwich panel, as were the side closures. Inside, it was filled with high-temperature insulation and the bottom was closed out with a titanium foil backing. The panel has provisions to be attached with mechanical fasteners to a smooth substructure. The panels are also vented to the local atmospheric pressure so that the aerodynamic pressure loads are borne by the substructure rather than the honeycomb panel that forms the outer surface. The prepackaged superalloy honeycomb concept eventually gave way to a more weight efficient ARMOR TPS.
Figure 2-6. Prepackaged superalloy honeycomb panel. (Figure obtained from Reference ) Titanium and super alloy multiwall TPS [18,19] were explored in the 1970s and1980s.
These designs consisted of a number of alternate layers of flat and dimpled metallic foils stacked on top of each other and joined only at the dimple peaks. The multiple layers prevent the radiated heat from the top surface from reaching the bottom. In some designs, the gaps in between the plates were also evacuated to prevent heat transfer through gas conduction and convection [11 p.52]. The heat conduction through the metallic foils would have to follow a long path and this lead to a structure with low conductivity through the panel thickness. The fabrication of the multi-wall TPS proved difficult and the evacuation of the gaps was impractical. The conductivity of the multi-wall TPS was also found to be twice that of the Space Shuttle tiles .
Another metallic TPS concept tested in the 1970s was a radiative metallic TPS . In this concept the top surface was made of radiative panels, which radiated most of the heat. The panel was corrugated to allow free thermal expansion in the direction transverse to the corrugations and also provide stiffness to resist panel flutter.
Functionally graded metal foam-core sandwich panels were studied for TPS applications [21–23]. It was found that low relative density foam close to the top face sheet and high density foam close to the bottom face sheet provided a good insulation from high temperature. However, the foam-core TPS did not provide any significant advantages when compared to other TPS concepts and there are difficulties in manufacturing of functionally-graded foams as well are attachment of foams to face sheets.
The first TPS concept for the human space flight used ablative TPS for space capsules Mercury, Gemini and Apollo. This was followed by the Space Shuttle Orbiter, which used ceramic tiles for thermal insulation. A fully metallic TPS concept called ARMOR TPS was proposed for the VentureStar program. In between, many other TPS concepts like titanium multiwall TPS and prepackaged honeycomb superalloy TPS were developed and tested. The common feature among all these different concepts was that all of them were designed as add-ons to the structure of the space vehicle. The TPS was not part of the structural design calculations. The function of the TPS was restricted to merely regulating the heat flowing in and out of the vehicle.
The thickness of the ablator or the ceramic tile or the insulation thickness was calculated based only on the heat transfer analysis. Of course, that is the reason they have been known as thermal protection systems.
There are many problems associated with these add-on concepts, the chief among them being the compromise on the robustness of the outer surface of the vehicles, usually arising due to incompatibility of the TPS and the load-bearing structure. For example, the Space Shuttle ceramic tiles needed to be separated from the structure of the vehicle using strain isolation pads.
Apart from the brittleness of the tiles, this relatively weak bonding between the TPS and the structure also exposed the vehicle to failures like loosening and separation of tiles from the structure. The ablator TPS never caused catastrophic failures during the space capsule flights.
However, these TPS designs were based on approximate analytical calculations and subsequent experimental observations. The analytical models were modified after each test flight to understand the ablation and structure response phenomena. The performance of the TPS and the structure were designated “acceptable” or “satisfactory” . These design practices probably lead to very conservative designs . In other words, the TPS was probably much heavier than actually required. The ablators were also beset with other problems like cracking and debonding from the structure, primarily due to mismatch in properties.
In order to overcome the problems associated with the add-on TPS concepts an Integral Thermal Protection System, ITPS, is required. The basic idea behind this concept is to combine the load-bearing structure and the TPS into one single structure. The ARMOR TPS design implemented this concept “partially”, in the sense that small aerodynamic pressure loads were borne by the TPS . But the structural design never really took these limited load-bearing capabilities of the TPS into account. Further there was no provision for in-plane loads in the ARMOR TPS design. The ITPS design, on the other hand, would have provisions to withstand considerable in-plane loads apart and large aerodynamic pressure loads. Thus, the ITPS can be an integral element of the spacecraft skin rather than a mere add-on.
The proposed ITPS design concept is illustrated in Figure 2-7. ITPS panels have a corrugated-core construction with face sheets on top and bottom joined by a corrugated-core in between. The empty spaces in the corrugated-core can be filled with high-temperature insulation material like Saffil™. This design is expected to efficiently combine the three passive TPS concepts of hot structure, insulated structure and heat sink (discussed in Section 2.1) into one single structure. The top face sheet acts as a hot structure and radiates out a large portion of the incident heat. The insulation material in the core leads to low thermal conductivity of the structure. Although the whole TPS panel acts as a heat sink for the heat absorbed, preliminary analyses indicate that the most effective means of adding thermal mass to the ITPS is at the bottom face sheet. Thus, the bottom face sheet has been illustrated as a heat sink in Figure 2-7.
The corrugated-core provides a structural connection between the top and bottom surfaces of the ITPS.
It is an easy design practice to de-link the TPS and the structural designs, that is, first obtain a structural design and then size the TPS to fit the requirements. In this way the technologies for the two designs can be developed independent of one another. Delays in overcoming technical challenges in one design will not lead to delays in the development of the other. When the two are assembled to form the space vehicle, some relatively minor design changes can be implemented to successfully develop the final vehicle assembly. For example, in the design of the Space Shuttle the structural design process and the ceramic TPS tile development were two different tasks. When they were assembled, however, there was a problem of compatibility, which was solved by using SIPs and tiles of small dimensions. In the case of ITPS design, the structural design and the TPS design are very much interlinked. This makes the vehicle design process even more complicated. This could also be one of the reasons why this route of TPS development was not pursued in the past.