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Thermal Protection System (TPS) Design and Optimization – A Case Study

Amarshi A. Bhungalia Amarshi.bhungalia@wpafb.af.mil

(937)255-8335 (Phone) (937)656-4945 (Fax) Structures Division

Air Vehicles Directorate Air Force Research Laboratory

Dr. Timothy Fry

Timothy.Fry@UDRI.UDayton.edu University of Dayton Research Institute

Jess Sponable and Daniel Tejtel

Aerospace Sciences Division Air Vehicles Directorate

Air Force Research Laboratory

Abstract: The Air Vehicle Directorate within the Air Force Research Laboratory (AFRL)

supports technology and configuration development for military reusable launch vehicle (RLV) concepts. RLVs have the potential to reduce cost, increase mission flexibility, and dramatically improve reliability over expendable launch vehicles. In the past, efforts like the joint National Aeronautics and Space Administration and Air Force 120 Day Study of operationally responsive spacecraft and National Aerospace Initiative (NAI) examined methods associated with acquiring RLVs and their technologies. However, the cost of implementing their recommendations was much higher than existing budget constraints permit.

Scientists and engineers at the Air Vehicles Directorate are working closely with the Defense Advanced Research Projects Agency (DARPA) and the Air Force Space Command (AFSPC) to find a practical, lower-cost RLV alternative. In support of the AFSPC Operationally Responsive Spacelift requirements, the directorate has been leading a series of in-house and contracted design studies. These studies began with directorate scientists developing an initial concept for a demonstrator vehicle capable of carrying 1,000 lbs of payload to orbit. These concepts were later refined by industry studies with costs estimated in the $500-800 million range. Although these demonstration costs are much lower than past options, additional cost reduction can be achieved by designing smaller, lower cost “Micro-X” flight demonstration vehicles. These Micro-X vehicles would demonstrate the feasibility of critical technologies, identify operations issues, and give a better understanding of the time and money required to make military RLV concepts a reality prior to embarking on an expensive acquisition program.

Space 200530 August - 1 September 2005, Long Beach, California

AIAA 2005-6809

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Introduction: Previous work with thermal protection systems (TPS) has shown that it is a major

roadblock to the development of rapidly reusable launch vehicles. The reasons for this TPS reusability roadblock are the long maintenance hours required to inspect and replace the relatively fragile thermal tiles used on the NASA Shuttle or ablation effects that change the aerodynamic properties of the vehicle. The current direction of RLV design is to reduce the post-flight maintenance requirements to the point where a vehicle can fly between 10 and 100 missions between routine TPS maintenance. This improvement requires that the TPS maintain its shape for aerodynamic purposes and be resistant to likely impact damage modes such as rain and micrometeorites while providing non-degrading thermal protection.

The NASA Shuttle thermal protection system consists of various materials applied externally to the outer structural skin of the orbiter to maintain skin temperature within an acceptable range, primarily during the reentry phase of the mission. The orbiter’s outer structural skin is constructed primarily of aluminum. During reentry from earth orbit, the TPS materials assure the orbiter’s aluminum skin does not exceed 350 oF. These materials perform in temperature ranges from minus 250 oF in the cold soak of space to entry temperatures that reach nearly 3,000 oF. The TPS also accommodates the forces and deflections of the orbiter airframe as it responds to the various aerodynamic and maneuvering loads along with thermal deflections.

Because the thermal protection system is installed on the outside of the orbiter skin, it establishes the surface responsible for the aerodynamic performance of the vehicle. The TPS performs this function in addition to its primary function of managing the thermal state of the reentry vehicle. If ablation or surface degradation is pronounced, the ability to control the vehicle and maintain a desired trajectory over several, or even one, mission is compromised. The need for military RLVs to alter their trajectory at will only increases the need to maintain the integrity of the aerodynamic surfaces of the vehicle.

TPS Materials: It should be pointed out that there is ongoing research in actively cooled TPS

including, phase change fluids, transpiration cooling, and heat pipes. However, the TPS considered in this paper are all passive and require no working fluids. These TPS concepts are selected for their stability at high temperatures and weight efficiency. Some of these materials are discussed below.

Both reinforced carbon-carbon (RCC) and advanced carbon-carbon (ACC) are used on the Micro-X wing leading edges; nose cap, including an area immediately aft of the nose cap on the lower surface (chine panel); and the immediate area around the forward orbiter/external tank structural attachment. RCC protects areas where temperatures exceed 2,300 oF during entry. Ongoing research is underway within AFRL to develop ACC materials suitable for use in cases where the leading edge temperature is in excess of 3,000 oF. The use of RCC or ACC leading edges generally requires the designer provide a radiation gap behind the leading edge to protect the internal wing structure. In some cases, the radiation gap alone is insufficient to protect the internal structure and

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additional solid or multilayer insulation may be required to optimize the leading edge TPS.

Black high-temperature reusable surface insulation (HRSI) tiles are used in areas on the Micro-X upper forward fuselage, including around the forward fuselage windows; the entire underside of the vehicle where RCC and ACC are not used; portions of the orbital maneuvering system and reaction control system pods; the leading and trailing edges of the vertical stabilizer; wing glove areas; elevon trailing edges; areas adjacent to the carbon-carbon on the upper wing surface; the base heat shield; the interface with wing leading edge carbon-carbon; and the upper body flap surface. The HRSI tiles protect areas where temperatures are below 2,300 oF. These tiles have a black surface coating to increase the surface emissivity of the HRSI. The added emissivity of the black surface helps the TPS radiate some of its absorbed heat back into space.

Another external insulation system used on reentry vehicles is called CRI. CRI, which stands for conformable reusable insulation, consists of an inner insulation layer covered by a high temperature Nextel ceramic fiber fabric. The CRI panels are stitched, much like a quilt, to maintain their dimensional stability and to prevent the fabric from separating from the internal insulation. CRI can be used on both the leeward and windward sides of the RLV. A potential major advantage of the CRI system is that these blankets can be fabricated in larger areas than HRSI tiles, which reduces the number of inter-blanket seams/seals (Figure 1) and potentially lowers post-flight maintenance requirements. Advanced flexible reusable surface insulation (AFRSI) is another blanket acreage TPS to be considered for use where external temperatures are below approximately 700 oF.

Figure 1 - Detail of Seal Between CMC (discussed below) Wrapped Tiles

Micro-X Concept: The Micro-X concept, shown in Figures 2 and Figure 3, is being jointly developed by

AFRL and Conceptual Research Corporation of Playa del Rey, California. The University of Dayton Research Institute acts as a contracting agent and performs TPS research and analysis for the program. This paper focuses on Micro-X’s role as a test bed for TPS. There are many other TPSAs (technologies, processes, and systems approaches) of the Micro-X that will not be discussed.

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The idea behind the Micro-X concept is to provide a demonstrator for advanced TPS technologies and rapidly reusable launch operations in a package considerably smaller and less expensive than that offered by traditional means. The Micro-X flies through the full aero-thermal flight envelope of other reentry vehicles, but in contrast to its larger cousins, the Micro-X is not designed to place a payload in orbit. The lack of payload-to-orbit capability allows the demonstrator to be very small, while still providing useful research data. The demonstrator is accelerated to high speeds either using a bimese configuration or using either a Minuteman or a Peacekeeper booster.

Figure 2 - Micro-X Demonstrator Concept

Figure 3 - Micro-X Demonstrator Concept

The vehicle considered in this case study is approximately 29 feet in length, weighs

8,750 pounds empty, and has a gross liftoff weight (GLOW) of 30,000 pounds. Flown in a standalone mode from the ground the Micro-X achieves a speed of Mach 4 and an altitude of 250,000 feet with a 3,000 lb payload. Mach 7 flight is attainable if no payload is flown. Later flights as a reusable second stage will expand the envelope over the full range of a two-stage-to-orbit (TSTO) RLV. Prior studies confirmed that to minimize weight and size, the vehicle should take off vertically and use liquid oxygen/hydrocarbon (LOX/HC) rockets. Similarly, because AFRL is developing the technology for a TSTO

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system, the vehicle shape is constrained to be cylindrical which enables simple modular mating of multiple stages. The shape also lends itself to a modular strategy that could demonstrate the technology of both a vertical and a horizontal landing capability. To keep the “logistics footprint” small and to minimize thermal environments, the vehicle is designed to be as short and squat to the ground as possible. The thermal reductions (to be discussed later) stemmed from being able to fly this kind of short/squat vehicle at higher angles of attack during reentry, potentially as high as 70 degrees.

Numerous TPS concepts exist in various stages of maturity, each with its own set of distinct advantages and disadvantages. The primary TPS concept selected for this case study is one of HRSI ceramic matrix composite (CMC) wrapped tile (Figure 4) with alumina enhanced thermal barrier (AETB) core on the fuselage; ACC with radiation gaps and insulation on the leading edges; and CRI blankets on the windward side. Future work will analyze the use of CRI on both the leeward and windward sides of the demonstrator in areas away from the stagnation flow. The philosophy of the Micro-X TPS was that off the shelf technologies, or those at a very high maturity level, would be used to protect the vehicle from expected thermal loads. Then, in addition to the baseline system, key areas of the demonstrator would be overlaid with experimental TPS to evaluate its effectiveness. This conservative philosophy is selected to prevent loss of the demonstrator in the event of failure in one of the experimental TPS locations.

Figure 4 - CMC Wrapped Tile

In order to quantify assessments of the individual TPS technology effectiveness, the

Micro-X vehicle, configured for a TSTO mission, was used to size and compare the TPS design and optimization. A thermal model was used to size the TPS for various zones on a baseline vehicle flying a reentry profile. The windward side and leading edge areas of Micro-X vehicle are discussed in this paper. Future papers will describe the leeward side analysis and optimization once that contractual effort is complete.

As was alluded to previously, the Micro-X is to fly a reentry trajectory at very high angle-of-attack to reduce its thermal loading. The present work evaluates the validity of that hypothesis and investigates advantages and disadvantages of a vehicle entering the atmosphere at 30 degree, 50 degree, and 70-degree angle-of-attack (AOA). This study reveals that the TPS weight for vehicle entering the atmosphere at 30 degree AOA is much higher than the vehicle entering the atmosphere at 70 degree AOA. This study also reveals the effects of different parameters on the TPS design of the vehicle.

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MINIVER: NASA’s MINIVER program was the tool selected for performing the thermal

analysis of the Micro-X during its various reentry trajectories with different TPS configurations. MINIVER, which stands for MINiature VERsion of the JA70 General Aerodynamic Heating Code, is an aero-thermal analysis program that allows engineers to rapidly determine the integrated thermal response of a TPS throughout an entire trajectory. MINIVER has the capability to analyze the heat transfer of a vehicle in the presence of shockwaves, stagnation flow, flat plate flow, swept cylinders, turbulent condition, laminar conditions, and transitioning flow, using either real or ideal gas models along with radiation effects. This code was developed by McDonnell Douglas Aeronautics Co. (MDAC) in the early 1970’s and continues to be maintained, enhanced, and validated by NASA at the Langley Research Center.

Unlike a computational fluid dynamics (CFD) solution, which provides a single-point solution, MINIVER is well suited to summing the effects of the entire flight path. While CFD may eventually be able to “fly” the entire trajectory, it is currently impractical to employ it for that task. The output of a MINIVER solution includes laminar and turbulent heat transfer rates, flow transition locations and the temperature history at several points through the TPS thickness as functions of time and calculated boundary conditions. MINIVER calculates the thermal boundary conditions based on a user-supplied trajectory, angle-of-attack, and geometry and correlations to a variety of basic flow conditions. MINIVER uses both wind tunnel and CFD point solutions and interpolates with them to obtain full trajectory solutions across dozens of user specified points on the vehicle surface in a matter of seconds. The set-up time, which is considerable, is offset by the rapidity of solutions and TPS thickness optimizations.

MINIVER is a versatile engineering code that uses various well-known approximate heating methods, together with simplified flow fields and geometric shapes to model the vehicle and perform 1-dimensional transient heat transfer calculations through the thickness of the TPS. Post-shock and local flow properties based on normal-shock or sharp-cone entropy conditions are determined in MINIVER through user selection of the various shock shape and pressure options. Angle-of-attack effects are simulated either by an equivalent tangent-cone or by an approximate cross flow option. The flow can be calculated for either two- or three-dimensional surfaces. However, the three-dimensional effects are available only with the Mangler transformation for flat-plate to sharp-cone conditions. MINIVER has been used extensively as a preliminary design tool in government and industry and has demonstrated excellent agreement with more detailed solutions for stagnation and windward acreage areas on a wide variety of vehicle configurations, including the Space Shuttle orbiter, HL-20, X-33, X-34, X-37, X-43, and National Air Space Plane (NASP). The principle advantage of this engineering code over some of the more detailed methods is the speed with which the analyses can be performed for each flow condition along a trajectory. Its strength lies in its ability to quickly provide the time-dependent thermal environments required for TPS analysis and sizing. Various different output parameters of MINIVER that are used in TPS design will be discussed in detail in a moment.

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The implementation of MINIVER used at AFRL uses MINIVER-1991 combined with a graphical user interface (GUI) developed by Technosoft, Inc. to aid in generating the body geometry, flow curves, and analysis points. The Technosoft GUI, called interactive missile design (IMD), is written using Technosoft’s adaptive modeling language (AML), and uses a combination of IGES geometry import and user generated geometry to generate a MINIVER input deck. Once the user has selected the appropriate flow curves, specified the analysis points of interest, TPS configuration, temperature-dependent material properties, flow conditions, and trajectory, IMD calls MINIVER and performs a complete thermal analysis for each analysis point. The user interrogates the results file to determine if any temperature limits are exceeded or are too conservative and reruns the analysis with a different TPS configuration as appropriate. The analysis/optimization cycle takes only a few moments once the initial setup is complete. Depending on the complexity of the vehicle and the number of analysis points the user wishes to use, the initial setup may take several days.

So far, in this section the purpose and method has been discussed. It is worthwhile to spend a little time discussing how a “MINIVER analysis” is performed. The MINIVER code consists of the following three sub-programs: PREMIN, the preprocessor used to set up the input; LANMIN (LANgley MINiver), used to compute the aero-thermal environments; and EXITS, used to predict the thermal response of the TPS. As mentioned in the previous paragraph, MINIVER has been integrated with the AML to ease pre- and post-processing.

The time advantage of the AML GUI has been quantified by AFRL engineers. When the GUI is used to formulate PREMIN input files it takes 2-3 days to generate several dozen analysis points, complete with geometry, flow curves, and thermal-fluid input parameters. When MINIVER is used in a stand-alone manner, it may take 2 to 3 days to formulate analysis for one point. Use of AML integration saves 2.75 days for each analysis point beyond the first one. Without the use of the AML GUI, each minor change to the analysis input requires the process start from the beginning, which adds tremendous time to the analysis. A final benefit to the AML GUI is that it provides the user with rudimentary error-diagnostics of the input data, which is in contrast to no error checking/diagnostics in the stand-alone version of MINIVER.

AML GUI integration with MINIVER also allows PREMIN, LANMIN, and EXITS data preparation at the same time as input for combined aero-thermal analysis. The LANMIN module output gives flow condition and thermal data for the outer surface of the vehicle. The LANMIN output, along with some other data, creates the needed input data such as number of layers, material thickness, and material properties for EXITS. The EXITS output data consist of transient temperature, radiation heat flux, convection heat flux, and TPS weight per square feet surface area.

MICRO-X TPS Design: The Micro-X serves several technical objectives. These include demonstration of

TPS concepts, development of high rate operations, system reusability, and low development costs. The vehicle size (29 feet, 8,750 pounds dry, and 30,000 pounds gross launch weight) is dictated primarily by the objective to define the minimum size vehicle that addresses key TPSAs. The vehicle designer, Dan Raymer, has provided design

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flexibility to allow mating two Micro-X vehicles in a bimese configuration to demonstrate stage separation and all some in increase overall speed and altitude performance.

The TPS, whose analysis is described here, requires a 350 oF maximum cold-side temperature owing the graphite-epoxy material used in the Micro-X construction. Subsequent analyses will explore TPS weight savings as a function of increased maximum cold-side temperature. It is estimated that some next generation composite materials may be able to sustain 700 oF with acceptable life-limit degradation. The wing and fuselage TPS wetted area for this demonstrator is 267-ft2. Subsequent work will address the leeward side.

Preliminary investigation by AFRL and the vehicle designer indicated it would be advantageous to perform a high angle-of-attack reentry. Quantifying this suggestion required running thermal analyses of the Micro-X for 30o, 50o, and 70o AOA reentry trajectories. Schafer Corporation provided the requisite POST trajectory results for inclusion in the MINIVER runs. The three trajectories are shown in Figure 5, Figure 6, and Figure 7.

AOA 30 Dg, Trajectory (Alt. & Vel.)

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Figure 5 – Micro-X 30o AOA Reentry Trajectory

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AOA 50 Trajectory (Alt. & Vel.)

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Figure 6 – Micro-X 50o AOA Reentry Trajectory

AOA 70 Trajectory (Alt. & Vel.)

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Figure 7 - Micro-X 70o AOA Reentry Trajectory

Using POST (Program to Optimize Simulated Trajectories), Schafer engineers ran each trajectory by fixing the AOA at a particular value and accepting whatever heating and g-loading that results. The altitude oscillations seen in the 30o and 50o trajectories are a result of the vehicle generating excess lift as it plunges into the thickening atmosphere.

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At the top of each oscillation, the excess lift is reduced to the point where the vehicle can resume its descent into the atmosphere.

The general process for a TPS thermal analysis and optimization is presented schematically in Figure 8. In the case of the AFRL implementation of MINIVER, the definition of the “MINIVER case” with PREMIN is greatly simplified and partially automated with the Technosoft AML GUI.

Figure 8 - TPS Thermal Analysis and Optimization Using MINIVER

Following the process flow schematic on Figure 8, it can be seen that PREMIN and LANMIN components of MINIVER setup the problem and solve the heat transfer problem for the calculated boundary conditions. In the AFRL implementation, the MINPLT (plotting of results) is omitted in favor of plotting results in Excel. Finally, the EXITS module calculates the transient 1-dimensional heat transfer through the TPS as it “flies” through the trajectory. This step is where MINIVER shines in comparison to CFD. The EXITS solution is the integrated result of many steps of quasi-steady thermal boundary conditions acting on the TPS, whose current-step initial condition is the result of the previous step solution.

Each trajectory analysis consists of 48 “analysis-points” (analysis-points is an AML GUI reserved word) distributed over the vehicle nose, leading edges, windward fuselage, and windward wing surfaces. The choice of analysis points belongs to the analyst and in the case of the Micro-X, 48 was sufficient to cover the vehicle with a reasonable analysis grid, without spending excessive time doing model setup. During post processing, a rectangle is centered on each analysis point, whose semi-length is half that of adjacent

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chordwise analysis points and whose semi-width is half that of adjacent spanwise analysis points. Recall that MINIVER performs a 1D analysis with no in plane heat transfer. Subject to that constraint, the temperature and TPS thickness are assumed constant within the results rectangle. Figure 9 illustrates this idea.

Figure 9 - Analysis Points, Flow Curves, and Results Rectangle

The heat transfer solution and optimized TPS configuration at each analysis point is

applied to its “results rectangle” and the TPS area weight (a MINIVER result) is multiplied by the rectangle area. The TPS weights for all rectangles are summed to determine the total TPS weight for the vehicle. The weights for the Micro-X TPS as a function of reentry angle-of-attack are shown in Table 1.

Table 1 - Micro-X TPS Weights vs. Reentry Angle-of-attack

TPS Wt. Lbs.(AOA 30)

TPS Wt. Lbs.(AOA 50)

TPS Wt. Lbs. (AOA 70)

Windward Fuselage 822.1 687.9 538.9 Left + Right Wing 178.2 150.7 121.4

Total Windward TPS 1000.3 838.6 660.3 The key result of this work is to validate and quantify the designer’s hypothesis that

very high angle-of-attack reentry yields a less demanding thermal condition for the Micro-X demonstrator. In quantifiable terms, the 70o AOA trajectory requires a thermal protection system that is 34% lighter than the TPS for the 30o AOA trajectory. Another way to view the benefits of the 70o trajectory is that if the vehicle were to be designed for the 30o there would be considerable thermal margin during the majority of flights when reentry at higher AOA is acceptable.

The TPS configuration for the leading edges and nose stagnation area is based on the

lay-up drawn in Figure 10. The configuration is based upon a nose radius of 1.5 feet at

Results Rectangle

Flow Curve

Analysis Point

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the stagnation point and wing leading edge radii ranging from 0.7-1.2 feet. In this configuration, working from the hot-side to the cold-side, the vehicle is protected by a layer of ACC, followed by a radiation gap, backed by LI-2200 insulation, affixed by RTV (room temperature vulcanizing), to a thermal strain isolation slab. Using the LI-2200 insulation thickness as the single optimization variable, the thickness is adjusted to attain a cold-side temperature of 350 oF at each analysis point. The adjustment is performed manually, but the subsequent analysis at the new thickness takes only a second or two on late-model PC.

ACC((N) thin skin 0.02000 in. Max Temp = 2800 Deg F

0.02000 in. ZIRCONIUM Max Temp = 2900 Deg F ------------------------------------------------------------ radiation gap 0.50000 in. ------------------------------------------------------------ 0.02000 in. ZIRCONIUM Max Temp = 2900 Deg F

LI-2200 (N) slab 1.31000 in. Max Temp = 2700 Deg F

RTV-560 thin skin 0.00800 in. Max Temp = 550 Deg F

17 LB SIP .16 IN slab 0.16000 in. Max Temp = 600 Deg F

Figure 10 - Nose and Wing Stagnation Point TPS Configuration

The TPS lay-up for the windward acreage areas is shown in Figure 11. In this

configuration, working from the hot-side to the cold-side, the vehicle is protected by a layer of HRSI, followed by a CMC tile wrap, backed by AETB LI-900 insulation, affixed by RTV, to a thermal strain isolation slab. Using the AETB LI-900 insulation thickness as the single optimization variable, the thickness is adjusted to attain a cold-side temperature of 350 oF at each analysis point.

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g g g y p=== 1. o==================================================== ----- HRSI COAT thin skin 0.05000 in. Max Temp = 2300 Deg F === 2. o==================================================== ----- CMC WRP TILE DENSI slab 0.05000 in. Max Temp = 2300 Deg F === 3. o==================================================== ----- 4. o 5. o 6. o 7. o 8. o 9. o 10. o 11. o 12. o 13. o 14. o 15. o 16. o AETB Mod LI-900 slab 3.58600 in. Max Temp = 2300 Deg F 17. o 18. o (This Layer Thikkness varies from pont to point) 19. o 20. o 21. o 22. o 23. o 24. o 25. o 26. o 27. o 28. o 29. o 30. o 31. o 32. o 33. o 34. o 35. o 36. o 37. o 38. o ===39. o==================================================== ----- RTV-560 (Bonding Epoxy) thin skin 0.00800 in. Max Temp = 550 Deg F ===40. o==================================================== ----- 41. o 42. o 17 LB SIP .16 in slab 0.16000 in. 43. o ===44. o==================================================== -----

Figure 11 - Windward Acreage TPS Configuration

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In addition to the optimized weights shown above, many readers will be interested in the temperature and heat flux vs. time histories of the Micro-X TPS at the three angles of attack. The “hump” in the 30° trajectory is due to the “plunge and rise” oscillation the vehicle makes as it enters the atmosphere.

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Figure 12 - Fuselage Hot-Side TPS Temperature vs. Time at Three AOA Trajectories

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High AOA Vehicle, Fuselage Flux

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Figure 13 - Fuselage Hot-Side TPS Heat Flux vs. Time at Three AOA Trajectories

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Figure 14 – Wing Hot-Side TPS Temperature vs. Time at Three AOA Trajectories

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High AOA Vehicle, Wing Flow Curve 1 Flux

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Figure 15 - Wing Hot-Side TPS Heat Flux vs. Time at Three AOA Trajectories

Summary: A key requirement of the Micro-X demonstrator is that it be able to demonstrate a

number of technologies, not the least of which is experimental TPS, without loss of the vehicle if the experimental technology fails. The work described in this paper shows that MINIVER and the Technosoft GUI used by AFRL and UDRI engineers can be used to rapidly analyze a number of TPS configuration on the Micro-X demonstrator. The results of the work show a 34% weight savings (or extra thermal margin) can be achieved by flying a 70o AOA reentry trajectory vs. a 30o AOA trajectory. The next steps for this work are to explore various other TPS configurations, for example CRI in acreage areas and advanced carbon-carbon on the nose and leading edges, to reduce the number of different parts required to cover the vehicle.

References: 1. MINIVER Upgrade for the AVID System, Volume 1: LANMIN User’s

Manual, NASA Contractor Report 172212, Engel, Carl D. and Praharaj, Sarat C., REMTECH, Inc., Huntsville, Alabama, August 1983.

2. MINIVER Upgrade for the AVID System, Volume 2: LANMIN Input Guide,

NASA Contractor Report 172213, Engel, Carl D. and Schmitz, Craig P., REMTECH, Inc., Huntsville, Alabama, August 1983.

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3. MINIVER Upgrade for the AVID System, Volume 3: EXITS User’s and Input

Guide, NASA Contractor Report 172214, Pond, John E. and Schmitz, Craig P., REMTECH, Inc., Huntsville, Alabama, August 1983.

4. Micro-X Rocket-Powered Technology Demonstrator Design Study, Phase 4

Final Report, UDRI Subcontract RSC03002, Raymer, Daniel P., May 2005. 5. MINIVER – A Versatile Aerothermal Analysis/TPS Design Tool, NASA

Langley Research Center, Wurster, Kathryn E., October 2000.

6. Various emails and telephone conversations with Technosoft, Inc. software developers since April 2004.