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Today, the least expensive means of delivering payloads into space is by using expendable rocket launch vehicles like the United States’ Delta, Europe’s Ariane, Russia’s Proton, China’s Long March and India’s Geosynchronous Satellite Launch Vehicle (GSLV). The biggest disadvantage of these vehicles is that they are for one time use and have to be rebuilt for each launch. NASA built the Space Shuttle Orbiter as a partially reusable vehicle with the capabilities to deliver and bring back men and materials from space. It can also be used as a short-term space laboratory, deliver payloads to low earth orbit, and repair and retrieve malfunctioning satellites.

However, it amounts to enormous cost to refurbish the solid rocket boosters, build a new external fuel tank and maintain the Shuttle between flights. In order to overcome the challenges ∗ These estimates are for “Western” launch vehicles and the variations in cost of payloads are due to different vehicles and different orbits into which the payloads are delivered.

associated with the Space Shuttle, NASA along with US industry aimed at developing a fully reusable single-stage-to-orbit (SSTO) vehicle, the VentureStar. However, this vehicle development has been stopped due to various budgetary and technological constraints.

In 1993, NASA made a study “Access to Space” [2], which concluded that an SSTO reusable launch vehicle (RLV) can bring down the cost of access to space by an order of magnitude. Thus, despite enormous challenges and setbacks, reusable or partially reusable launch vehicles are still something to strive for in future, at least for manned missions to the space, Moon and Mars.

When a spacecraft enters any planetary atmosphere, it usually does so at hypersonic speeds of over Mach 20 and up to Mach 40. High energy, high velocity flights through planetary atmospheres lead to extremely severe aerodynamic heating and pressures on the vehicle exteriors. The vehicle structure needs to be protected from damage due to this heating. Design features incorporated into a vehicle to withstand this aerodynamic heating and protect it from damage are known as Thermal Protection Systems or TPS.

Early thermal protection systems used on lunar vehicles, like Mercury, Gemini and Apollo, were single-use ablators, in which material on the surface ablates by absorbing the heat and thus prevents the heat from entering the capsule [3–9]. This concept is still in use on the Soyuz capsules used by Russia for manned missions. The Space Shuttle Orbiter uses thermal protection systems made of high-temperature-resistant ceramic tiles and blankets [10,11], which primarily function as insulators and prevent heat from reaching the vehicle interiors. These tiles make the Space Shuttle exterior very brittle and susceptible to damage even due to small impact loads. In order to overcome these difficulties, scientists at NASA developed a fully metallic ARMOR TPS [12–16] for the VentureStar program. The ARMOR TPS uses metallic honeycomb sandwich panels as the outer surface, to withstand pressure loads, and high efficiency fibrous insulation materials to prevent heat from reaching the vehicle interior. However, the load-bearing capabilities of this TPS are limited, and large loads like in-plane inertial loads cannot be accommodated under this design and are taken by the structure of the vehicle.

TPS occupies a huge acreage on the vehicle exteriors and forms a major part of the launch weight. Therefore, it is imperative that apart from making the TPS suitable for thermal protection purposes, it should also be made lightweight in order to keep the launch costs down. TPS is also required to present a robust external surface for the vehicle. One approach proposed towards achieving this goal it to combine the functions of load-bearing and thermal protection into one structure.

Efforts are on to develop a robust load-bearing TPS structure called Integral Thermal Protection System, ITPS. The ITPS has both the thermal protection and load-bearing capabilities integrated into one structure. This is unlike the Space Shuttle TPS in which thermal protection and load-bearing functions are performed by completely different members. The biggest challenge of an ITPS is that the requirements of a load-bearing member and a TPS are contradictory to one another. A TPS is required to have low conductivity and high service temperatures. Materials satisfying these conditions are ceramic materials, which are also plagued by poor structural properties like low impact resistance, low tensile strength and low fracture toughness. On the other hand, a robust load-bearing structure needs to have high tensile strength and fracture toughness and good impact resistance. Materials that satisfy these requirements are metals and metallic alloys, which have relatively high conductivity and low service temperatures.

A structure with good load-bearing properties is usually a dense structure. A dense structure presents a larger heat conduction path between the outer surface and vehicle interiors. On the contrary, a good TPS is usually made of low-density material. The challenge is to develop a structure that can combine these two functions into one single structure and this is the objective of the current research effort.

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TPS forms the external surface of a vehicle and is exposed to a variety of environmental conditions at different times during its flight. The specific times and nature of these conditions differ for different missions and vehicle type. For example, the Space Shuttle TPS encounters different types of environments during both launch and reentry, while space capsules like Apollo and Soyuz encounter severe environments only during reentry, as they are well shielded during launch. The aerodynamic heating and pressures experienced by a TPS are also dependent on the vehicle shape and the trajectories taken by the vehicle. A set of general requirements of a TPS is presented in this section.

Heat Load. The primary function of a TPS is to regulate the heat flow to and from the vehicle. In most cases, the major design driver for TPS is the aerodynamic heating during vehicle reentry into Earth’s atmosphere. Depending on the vehicle design, the TPS may be subjected to considerable plume heating. Plume heating is the heating of the TPS due to the combustion process of propellants that generate considerable heat upon exiting the rocket nozzles. In some case like the VentureStar, the TPS may act as insulation to the cryogenic tanks. In this case, the function of a TPS is to prevent the formation of liquid oxygen or ice on the vehicle exteriors as well as to limit the amount of cryogenic fuel lost to boil-off. Solar heating and radiative heat loss from the outer surface to space may be factors in design of TPS for vehicles that spend considerable amount of time in space. Apart from regulating the heat flow, the materials used for TPS should be able to withstand the high temperatures without substantial degradation in material properties.

Mechanical Loads. Mechanical loads on the TPS include transverse aerodynamic pressure loads, in-plane inertial loads, and acoustic and dynamic loads. These loads vary widely depending on the vehicle shape and the position of the TPS on the vehicle. In the Space Shuttle, the nose cone, and wing and tail leading edges experience tremendous aerodynamic pressure.

Aerodynamic pressure loads are also very high for windward side of a space capsule. In contrast, the leeward side of a space capsule and most other areas on the Space Shuttle are subjected to slight negative aerodynamic pressure loads, which may not be major design drivers. In-plane inertial loads on a vehicle are extremely high for TPS on the rear part of the vehicle than compared to the front portion. In addition, there may be acoustic and dynamic loads on the TPS, usually generated by the propulsion system. TPS must be able to withstand these loads without failure in the form of yielding, buckling or fracture.

Deflection Limits. TPS forms the external surface of the vehicle and, therefore, dictates the aerodynamic profile presented by the vehicle. Thermal and mechanical loads could lead to considerable amount of deflection of the top surface. It is required that the top surface deflections of the TPS be below certain limit in order to maintain the smooth aerodynamic profile of the vehicle. Excessive local deflection, such as a top surface dimpling, can lead to severe local aerodynamic heating [12,13], which may lead to catastrophic failures.

Impact Loads. Good impact resistance is another desirable property of a TPS, which can be subjected to various types of impact. During installation and maintenance, the TPS can be subjected to various handling damages and low speed impacts like accidental dropping of the panels on a hard surface or accidentally dropping of tools on the TPS. During launch and landing, it may be subjected to impact from runway debris. During flight, it may encounter small object impact like hail, bird-strike, rain, snow, and dust. In the space, it may be subjected to high-speed impact from space debris and meteorites. A TPS should be able to withstand these small-object-high-velocity impact loads so that they do not lead to a catastrophic failure. Impact resistance of TPS is beyond the scope of this research work.

Chemical Deterioration. During reentry the high temperatures on the top surface may make the TPS susceptible to chemical attack such as oxidation. TPS properties may be altered due to water absorption and spills of various substances during maintenance. TPS should be able to resist these chemical attacks. Chemical stability of TPS is not considered in the this research effort.

Low Cost Operability. Apart from initial fabrication and installation costs, the life cycle cost of a TPS also includes its maintenance throughout its life. TPS that is robust and operable leads to low maintenance costs. A robust TPS is one that is not easily damaged by its design environment and can withstand damage to a certain extent without requiring immediate repair.

An operable TPS is one that can be inspected and maintained easily, removed, replaced or repaired, if necessary.

Light Weight. TPS occupies a huge acreage on space vehicles and constitutes a major portion of the launch weight. Increase in launch loads could lead to increased fuel requirements and/or decrease in the payload mass. Therefore, TPS should be designed for minimum mass to perform its various functions.

The large number of often-contradictory requirements makes the TPS design a formidable task. It requires a good understanding of many technical disciplines. TPS design involves transient heat transfer analysis, stress and deflection analysis, and mechanical and thermal buckling analysis. In this research, the heat load on a TPS is assumed to be available. Calculation of aerodynamic heating and pressures is out of the scope of this research, as it requires a large amount of information like vehicle shape, trajectory and velocity at different times during its flight.

ITPS design can be considered to be an extension of the ARMOR TPS design. It is aimed at including design features to carry large in-plane loads (typical of Space Shuttle) and substantially higher pressure loads, apart from providing the insulation properties similar to that of the ARMOR TPS. Previous ITPS attempts like multi-wall TPS (details in Section 2.2) faced difficulties like manufacturability and evacuation of the gaps in between the walls.

Manufacturability can be a potential concern for the current ITPS design as well; however, addressing this concern is beyond the scope of this research work. The use of high-efficiency insulation materials, like SAFFIL, is considered to be the major improvement over multi-wall TPS, as the insulation material blocks the radiation from the top surface and decreases the thermal conductivity of the structure substantially. Some of the demonstrated manufacturing capabilities for the ARMOR TPS can be potentially exploited for ITPS fabrication as well.

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The objective of this research work is to develop a multidisciplinary design optimization procedure to design an Integral Thermal Protection System by reconciling the various conflicting requirements of a TPS and a load-bearing member into one single structure. The goal is to demonstrate the feasibility of such a design concept by comparing it to the other TPS performance and weight.

Another objective is to study the influence of variation of input parameters on the ITPS weight. Some of the parameters studied are surface emissivity, heat loss from ITPS to the vehicle interiors, initial temperature of the structure before reentry, thermomechanical properties of the materials used for the ITPS, different displacement/rotation boundary conditions on the panels and the magnitude of pressure loads and in-plane loads. These would help formulate a general set of guidelines for ITPS design.

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1.4.1 The Optimization Problem Various conflicting requirements of a TPS and a load-bearing structure need to be reconciled to find a feasible solution. This is possible by formulating the design process as an optimization problem. The objective of the process is to make the structure as light as possible while fulfilling all the functions required of an ITPS. Therefore, the obvious choice for an objective function would be mass per unit area of the ITPS panel, M. Various functions of the panel were formulated in the form of constraints for the optimization problem. Thus, the optimization problem can be stated as “minimize mass per unit area of the panel while satisfying all the constraints” such as temperatures, stresses and deflections limits in various parts of the structure, and withstand global/local buckling.

Critical functions of an ITPS that have a potential influence on the design were taken into account [12,13].

The following are the four critical constraints taken into account for the design (see Section 2.5 for more details):

1. Maximum temperature of the bottom surface of the panel must be below certain limit.

2. Panel must be able to withstand global/local buckling due to mechanical and thermal forces.

3. Maximum stresses in the various members of the panel must be within allowable limits.

4. Maximum deflection of the top surface of the panel must be below acceptable limit.

The mathematical statement for the optimization problems is as follows

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