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NASA NTRS2006

Detection of Ketenimine (CH2CNH) in SGRB2(N) Hot Cores

Ketenimine (CH2CNH) has been detected in absorption toward the starforming region Sagittarius B2(N) with the 100-m Green Bank Telescope (GBT) by means of three rotational transitions: 7(sub 16)-8(sub 08) at 41.5 GHz, 8(sub 19)-9(sub 09), at 23.2 GHz, and 9(sub 18)-10(sub 0,10) at 4.9 GHz. Ketenimine has a sparse rotational spectrum below 50 GHz. From transition line strength arguments, the spectral lines found are the ones most likely to be detected and occur in spectral regions that have little possibility of confusion with other molecular species. Partially resolved hyperfine structure is apparent in the 4.9 GHz transition which has energy levels approximately 50 K above ground state level; the absorption seen in this transition appears to be emanating from gas in close proximity to the LMH hot core that has a systemic LSR velocity of +64 kilometers per second. By comparison, the 41.5 GHz and 23.2 GHz transitions have lower energy levels of approximately 33 K and approximately 41 K, respectively; and show absorption against the two star-forming SgrB2(N) hot cores with systematic LSR velocities of +64 (the LMH) and +82 kilometers per second. These ketenimine data show that the hot core at +82 kilometers per second is cooler than the hot core at +64 kilometers per second. Ketenimine is likely formed directly from its isomer methyl cyanide (CH3CN) by tautomerization driven by shocks that pervade the star-forming region.

Lovas, F. J.↗

NASA NTRS2006

Space Shuttle Main Engine Low Pressure Oxidizer Turbo-Pump Inducer Dynamic Environment Characterization through Water Model and Hot-Fire Testing

The Low Pressure Oxidizer Turbopump (LPOTP) inducer on the Block II configuration Space Shuttle Main Engine (SSME) experienced blade leading edge ripples during hot firing. This undesirable condition led to a minor redesign of the inducer blades. This resulted in the need to evaluate the performance and the dynamic environment of the redesign, relative to the current configuration, as part of the design acceptance process. Sub-scale water model tests of the two inducer configurations were performed, with emphasis on the dynamic environment due to cavitation induced vibrations. Water model tests were performed over a wide range of inlet flow coefficient and pressure conditions, representative of the scaled operating envelope of the Block II SSME, both in flight and in ground hot-fire tests, including all power levels. The water test hardware, facility set-up, type and placement of instrumentation, the scope of the test program, specific test objectives, data evaluation process and water test results that characterize and compare the two SSME LPOTP inducers are discussed. In addition, dynamic characteristics of the two water models were compared to hot fire data from specially instrumented ground tests. In general, good agreement between the water model and hot fire data was found, which confirms the value of water model testing for dynamic characterization of rocket engine turbomachinery.

Patrick Arellano↗

NASA NTRS2007

Novel Thin Film Sensor Technology for Turbine Engine Hot Section Components

Degradation and damage that develops over time in hot section components can lead to catastrophic failure of the turbine section of aircraft engines. A range of thin film sensor technology has been demonstrated enabling on-component measurement of multiple parameters either individually or in sensor arrays including temperature, strain, heat flux, and flow. Conductive ceramics are beginning to be investigated as new materials for use as thin film sensors in the hot section, leveraging expertise in thin films and high temperature materials. The current challenges are to develop new sensor and insulation materials capable of withstanding the extreme hot section environment, and to develop techniques for applying sensors onto complex high temperature structures for aging studies of hot propulsion materials. The technology research and development ongoing at NASA Glenn Research Center for applications to future aircraft, launch vehicles, space vehicles, and ground systems is outlined.

Wrbanek, John D.↗

NASA NTRS2008

Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this paper is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and will be discussed briefly.

Glass, David E.↗

NASA NTRS2012

Variability of Jupiter's Five-Micron Hot Spot Inventory

Global upheavals on Jupiter involve changes in the albedo of entire axisymmetric regions, lasting several years, with the last two occurring in 1989 and 2006. Against this backdrop of planetary-scale changes, discrete features such as the Great Red Spot (GRS), and other vortices exhibit changes on shorter spatial- and time-scales. We track the variability of the discrete equatorial 5-micron hot spots, semi-evenly spaced in longitude and confined to a narrow latitude band centered at 6.5degN (southern edge of the North Equatorial Belt, NEB), abundant in Voyager images. Tantalizingly similar patterns were observed in the visible (bright plumes and blue-gray regions), where reflectivity in the red is anti-correlated with 5-microns thermal radiance. Ortiz et al. (1998, GRL, 103) characterized the latitude and drift rates of the hot spots, including the descent of the Galileo probe at the southern edge of a 5-micron hot spot, as the superposition of equatorial Rossby waves, with phase speeds between 99 - 103m/s, relative to System III. We note that the high 5-micron radiances correlate well but not perfectly with high 8.57-micron radiances. Because the latter are modulated primarily by changes in the upper ammonia (NH3) ice cloud opacity, this correlation implies that changes in the ammonia ice cloud field may be responsible for the variability seen in the 5-~m maps. During the NEB fade (2011 - early 2012), however, these otherwise ubiquitous features were absent, an atmospheric state not seen in decades. The ongoing NEB revival indicates nascent 5-~m hot spots as early as April 2012, with corresponding visible dark spots. Their continuing growth through July 2012 indicates the possit.le re-establishment of Rossby waves. The South Equatorial Belt (SEB) and NEB revivals began similarly with an instability that developed into a major outbreak, and many similarities in the observed propagation of clear regions.

Yanamandra-Fisher, Padma A.↗

NASA NTRS2014

Local ISM 3D Distribution and Soft X-ray Background Inferences for Nearby Hot Gas

Three-dimensional (3D) interstellar medium (ISM) maps can be used to locate not only interstellar (IS) clouds, but also IS bubbles between the clouds that are blown by stellar winds and supernovae, and are filled by hot gas. To demonstrate this, and to derive a clearer picture of the local ISM, we compare our recent 3D IS dust distribution maps to the ROSAT diffuse Xray background maps after removal of heliospheric emission. In the Galactic plane, there is a good correspondence between the locations and extents of the mapped nearby cavities and the soft (0.25 keV) background emission distribution, showing that most of these nearby cavities contribute to this soft X-ray emission. Assuming a constant dust to gas ratio and homogeneous 106 K hot gas filling the cavities, we modeled in a simple way the 0.25 keV surface brightness along the Galactic plane as seen from the Sun, taking into account the absorption by the mapped clouds. The data-model comparison favors the existence of hot gas in the solar neighborhood, the so-called Local Bubble (LB). The inferred mean pressure in the local cavities is found to be approx.9,400/cu cm K, in agreement with previous studies, providing a validation test for the method. On the other hand, the model overestimates the emission from the huge cavities located in the third quadrant. Using CaII absorption data, we show that the dust to CaII ratio is very small in those regions, implying the presence of a large quantity of lower temperature (non-X-ray emitting) ionized gas and as a consequence a reduction of the volume filled by hot gas, explaining at least part of the discrepancy. In the meridian plane, the two main brightness enhancements coincide well with the LB's most elongated parts and chimneys connecting the LB to the halo, but no particular nearby cavity is found towards the enhancement in the direction of the bright North Polar Spur (NPS) at high latitude. We searched in the 3D maps for the source regions of the higher energy (0.75 keV) enhancements in the fourth and first quadrants. Tunnels and cavities are found to coincide with the main bright areas, however no tunnel nor cavity is found to match the low-latitude b > or approx. 8deg, brightest part of the NPS. In addition, the comparison between the 3D maps and published spectral data favors a NPS central source region location beyond 230 pc, i.e. at larger distance than usually considered. Those examples illustrate the potential use of more detailed 3D distributions of the nearby ISM for the interpretation of the diffuse soft X-ray background.

Xray Emission↗

NASA NTRS2015

Hot-Fire Testing of 5N and 22N HPGP Thrusters

This hot-fire test continues NASA investigation of green propellant technologies for future missions. To show the potential for green propellants to replace some hydrazine systems in future spacecraft, NASA Marshall Space Flight Center (MSFC) is continuing to embark on hot-fire test campaigns with various green propellant blends.NASA completed hot-fire testing of 5N and 22N HPGP thrusters at the Marshall Space Flight Center’s Component Development Area altitude test stand in April 2015. Both thrusters are ground test articles and not flight ready units, but are representative of potential flight hardware with a known path towards flight application. The purpose of the 5N testing was to perform facility check-outs and generate a small set of data for comparison to ECAPS and Orbital ATK data sets. The 5N thruster performed as expected with thrust and propellant flow-rate data generated that are similar to previous testing at Orbital ATK. Immediately following the 5N testing, and using the same facility, the 22N testing was conducted on the same test stand with the purpose of demonstrating the 22N performance. The results of 22N testing indicate it performed as expected.The results of the hot-fire testing are presented in this paper and presentation.

Burnside, Christopher G.↗

NASA NTRS2015

A Multifunctional Hot Structure Heatshield Concept for Planetary Entry

A multifunctional hot structure heatshield concept is being developed to provide technology enhancements with significant benefits compared to the current state-of-the-art heatshield technology. These benefits can potentially enable future planetary missions. The concept is unique in integrating the function of the thermal protection system with the primary load carrying structural component. An advanced carbon-carbon material system has been evaluated for the load carrying structure, which will be utilized on the outer surface of the heatshield, and thus will operate as a hot structure exposed to the severe aerodynamic heating associated with planetary entry. Flexible, highly efficient blanket insulation is sized for use underneath the hot structure to maintain required operational internal temperatures. The approach followed includes developing preliminary designs to demonstrate feasibility of the concept and benefits over a traditional, baseline design. Where prior work focused on a concept for an Earth entry vehicle, the current efforts presented here are focused on developing a generic heatshield model and performing a trade study for a Mars entry application. This trade study includes both structural and thermal evaluation. The results indicate that a hot structure concept is a feasible alternative to traditional heatshields and may offer advantages that can enable future entry missions.

Walker, Sandra P.↗

NASA NTRS2016

Optimization of a Hot Structure Aeroshell and Nose Cap for Mars Atmospheric Entry

The National Aeronautics and Space Administration (NASA) is preparing to send humans beyond Low Earth Orbit and eventually to the surface of Mars. As part of the Evolvable Mars Campaign, different vehicle configurations are being designed and considered for delivering large payloads to the surface of Mars. Weight and packing volume are driving factors in the vehicle design, and the thermal protection system (TPS) for planetary entry is a technology area which can offer potential weight and volume savings. The feasibility and potential benefits of a ceramic matrix composite hot structure concept for different vehicle configurations are explored in this paper, including the nose cap for a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) and an aeroshell for a mid lift-to-drag (Mid L/D) concept. The TPS of a planetary entry vehicle is a critical component required to survive the severe aerodynamic heating environment during atmospheric en- try. The current state-of-the-art is an ablative material to protect the vehicle from the heat load. The ablator is bonded to an underlying structure, which carries the mechanical loads associated with entry. The alternative hot structure design utilizes an advanced carbon-carbon material system on the outer surface of the vehicle, which is exposed to the severe heating and acts as a load carrying structure. The preliminary design using the hot structure concept and the ablative concept is determined for the spherical nose cap of the HIAD entry vehicle and the aeroshell of the Mid L/D entry vehicle. The results of the study indicate that the use of hot structures for both vehicle concepts leads to a feasible design with potential weight and volume savings benefits over current state-of-the-art TPS technology that could enable future missions.

Langston, Sarah L.↗

NASA NTRS2018

Initial Criticality Assessments to Guide FMEAs on Rocket Engine Hot Fire Testing

Presented is an explanation of the use of the Initial Criticality Assessment (ICA) technique, a triage process for prioritizing required Failure Modes and Effects Analysis (FMEAs), for the European Service Module's main propulsion system's hot-fire test bed at White Sands New Mexico. Rather than instinctively performing many FMEAs of subsystems, or one large system level FMEA where every subcomponent is analyzed, the ICA guided an informed analysis of only the hardware that had a large impact to hazards. The low criticality hardware was documented via the ICA and no FMEA was performed; the work could then focus on the high criticality hardware. Thus a savings of Program resources was achieved. The experiences gained in creating these ICAs for this international collaborative project confirmed that the need for continuous communication across the technical teams is one of the greatest areas of emphasis. The European Space Agency (ESA) is developing the European Service Module (ESM), with its primary contractor, Airbus Defence and Space in Germany, for delivery to the National Aeronautics and Space Administration (NASA). The module will be equipped with a total of 21 engines to support NASA’s Orion spacecraft: one U.S. Space Shuttle Orbital Maneuvering System-Engine (OMS-E), eight auxiliary thrusters and 12 smaller RCS (Reaction Control System) thrusters. The main ESM propulsion system, used for large translational maneuvers, consists of one OMS-E... To qualify the design of the ESM propulsion subsystem (PSS) an all-steel Propulsion Qualification Module (PQM) structure is used to test the propulsion systems on Orion, including “hot firing” of the OMS engine, thrusters, and RCS. The PQM has been developed as a hot-fire test bed to be tested at the NASA White Sands Test Facility (WSTF). One of the objectives of the testing is to assure that the OMS-E can be safely operated with the PQM. Testing will also demonstrate that the PQM can set the proper upstream pressures and temperatures for the OMS-E to operate nominally given the PQM has never been tested in hot-fire operation with OMS-E before. In order to safely conduct the test campaign, hardware such as the engine subassembly, fluid feed lines, valves, electrical power lines, instrumentation, stiff links, installation Ground Support Equipment (GSE), and diffuser [whose objectives are to collect the exhaust of the OMS-E to actively cool down the exhaust gases, reduce thermal exchanges, and create a vacuum at OMS-E level before igniting], had to be analyzed for any hazards and failure modes.

Fault Tolerance↗

NASA NTRS2019

Bimetallic Channel Wall Nozzle Development and Hot-Fire Testing Using Additively Manufactured Laser Wire Direct Closeout Technology

NASA has been developing and advancing regeneratively-cooled channel wall nozzle technology for liquid rocket engines to reduce cost and schedules associated with fabrication. One of the primary methods being advanced is Laser Wire Direct Closeout (LWDC). LWDC was developed to provide an additively manufactured laser deposited closeout of the coolant channels that also forms the structural jacket in-situ. This technique has been previously demonstrated through process development and hot-fire testing on a series of subscale nozzles at NASA Marshall Space Flight Center. The hot-fire test articles were fabricated using monolithic alloys to simplify the fabrication process. Ongoing research is being conducted to further expand use of this process for increased scale and bimetallic or multi-alloy options. The use of multi-alloys is desired to fully optimize the combination of materials in the radial and axial directions to reduce overall weight of the nozzle and allow for higher thermal and structural margins on the channel wall nozzle. NASA recently completed process development and hot-fire testing of a series of channel wall nozzles that incorporate a copper-alloy as the hotwall liner material and a superalloy and combination thereof for the structural jacket using the LWDC technique. The fabrication process was further advanced by using a multi-alloy axial joint using explosive bonding integrating a copper-alloy at the forward end of the nozzle hotwall and a stainless-alloy for the remaining length. A third alloy was then used for the channel closeout using the LWDC process. This paper will describe the process development using the LWDC process for channel closeout utilizing the multi-alloys, hardware design and results from hot-fire testing on subscale multi-alloy LWDC channel cooled nozzles.

Gradl, Paul↗

NASA NTRS2019

Detection of Residual “Hot Spots” in RFI-Filtered SMAP Data

Radio frequency interference (RFI) is a well-documented problem for passive remote sensing of the Earth at L-band even though the measurements are made in the protected band at 1.413 GHz. Consequently, filtering for RFI is an important early step in the processing of measurements made by the SMAP (Soil Moisture Active/Passive) radiometer. However, the filtered data still include regions with suspiciously high antenna temperatures. One possible cause of these “hot spots” is interference not fully detected during RFI filtering. This paper presents evidence supporting this hypothesis and describes an algorithm to identify these “hot spots” so that they can be removed from the measurements. The impact of removing these “hot spots” is generally small, but evidence is presented that the brightness temperature and soil moisture improve when the hot spots are removed.

Soldo, Yan↗

NASA NTRS2019

HD 202772A b: A Transiting Hot Jupiter Around a Bright, Mildly Evolved Star in a Visual Binary Discovered by TESS

We report the first confirmation of a hot Jupiter discovered by the Transiting Exoplanet Survey Satellite (TESS) mission: HD 202772A b. The transit signal was detected in the data from TESS Sector 1, and was confirmed to be of planetary origin through radial velocity (RV) measurements. HD 202772A b is orbiting a mildly evolved star with a period of 3.3 days. With an apparent magnitude of V = 8.3, the star is among the brightest and most massive known to host a hot Jupiter. Based on the 27 days of TESS photometry and RV data from the CHIRON, HARPS, and Tillinghast Reflector Echelle Spectrograph, the planet has a mass of 1.017(Sup +0.070, sub 0.068) M(J) and radius of 1.545(sup +0.052, sub -0.060) R(J), making it an inflated gas giant. HD 202772A b is a rare example of a transiting hot Jupiter around a quickly evolving star. It is also one of the most strongly irradiated hot Jupiters currently known.

Songhu Wang↗

NASA NTRSDate not supplied

Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth’s atmosphere and non-Earth atmospheres. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This presentation will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the “insulated airplane” approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and will be discussed briefly.

ceramic matrix composites↗

NASA NTRSDate not supplied

Hot-Fire Testing Results of Additively Manufactured NASA HR-1 towards Large Scale Implementation of Channel Wall Nozzles

Additive manufacturing (AM) has provided new design and manufacturing opportunities to reduce cost and schedules, consolidate parts, and optimize performance. One technique being evaluated is Laser Powder Directed Energy Deposition (LP-DED), which provides a significant increase in scale compared to Laser Powder Bed Fusion (L-PBF). NASA along with industry partners have been developing the LP-DED process to demonstrate internal channel geometry and development components for use in liquid rocket engine channel cooled nozzles. One material that has been the focus of this development is NASA HR-1, an Iron-Nickel based superalloy optimized for use in high pressure hydrogen environments. Material properties and process characterization has been completed on the high density thin-wall material in addition to advancements of the supply chain. NASA has also completed fabrication of several subscale and full scale channel wall nozzles in LP-DED NASA HR-1 and completed hot-fire testing. This includes refinement of the process to produce thin-walls and various channel geometries to meet the requirements for channel wall nozzle applications. This paper will provide an overview of the LP-DED process development, material characterization and properties, component manufacturing, and hot-fire testing. Hot-fire testing was completed for a lander-class 7K-lbf thrust chamber using Liquid Oxygen (LOX)/Methane. The design overview and results from hot-fire testing will be presented in addition to hardware development for future testing on 2K-lbf and 35k-lbf thrust chambers and large-scale manufacturing technology demonstrators.

Paul Gradl↗

NASA NTRSDate not supplied

Process Development and Hot-fire Testing of Additively Manufactured NASA HR-1 for Liquid Rocket Engine Applications

Additive manufacturing (AM) has provided new design and manufacturing opportunities to reduce cost and schedules, consolidate parts, and optimize performance. One technique being evaluated is Laser Powder Directed Energy Deposition (LP-DED), which provides a significant increase in scale compared to Laser Powder Bed Fusion (L-PBF). NASA along with industry partners have been developing the LP-DED process to demonstrate internal channel geometry and development components for use in liquid rocket engine channel cooled nozzles. Optimized materials in the extreme high pressure and hydrogen environment for liquid rocket engines remains a key challenge. NASA has advanced an enabling material called NASA HR-1 (Hydrogen Resistant -1) as a solution using AM techniques. NASA HR-1 is a high-strength Fe-Ni superalloy designed to resist high pressure, hydrogen environment embrittlement, oxidation, and corrosion. NASA HR-1 meets materials requirements for liquid rocket engine components, including good hydrogen resistance, high conductivity, good low cycle fatigue performance, and high elongation and strength for components in high heat flux environments. Material properties and process characterization have been completed on the high density thin-wall material in addition to advancements of the supply chain. NASA has also completed fabrication of several subscale and full-scale channel wall nozzles in LP-DED NASA HR-1 and completed hot-fire testing. This includes refinement of the process to produce thin-walls and various channel geometries to meet the requirements for channel wall nozzle applications. This paper will provide an overview of the LP-DED process development, material characterization and properties, component manufacturing, and hot-fire testing. Hot-fire testing was completed for a lander-class 7K-lbf thrust chamber using Liquid Oxygen (LOX)/Methane. The design overview and results from hot-fire testing will be presented in addition to hardware development for future testing on 2K-lbf and 35k-lbf thrust chambers and large-scale manufacturing technology demonstrators.

DED↗

NASA NTRSDate not supplied

Process Development and Hot-fire Testing of Additively Manufactured NASA HR-1 for Liquid Rocket Engine Applications

Additive manufacturing (AM) has provided new design and manufacturing opportunities to reduce cost and schedules, consolidate parts, and optimize performance. One technique being evaluated is Laser Powder Directed Energy Deposition (LP-DED), which provides a significant increase in scale compared to Laser Powder Bed Fusion (L-PBF). NASA along with industry partners have been developing the LP-DED process to demonstrate internal channel geometry and development components for use in liquid rocket engine channel cooled nozzles. Optimized materials in the extreme high pressure and hydrogen environment for liquid rocket engines remains a key challenge. NASA has advanced an enabling material called NASA HR-1 (Hydrogen Resistant -1) as a solution using AM techniques. NASA HR-1 is a high-strength Fe-Ni superalloy designed to resist high pressure, hydrogen environment embrittlement, oxidation, and corrosion. NASA HR-1 meets materials requirements for liquid rocket engine components, including good hydrogen resistance, high conductivity, good low cycle fatigue performance, and high elongation and strength for components in high heat flux environments. Material properties and process characterization have been completed on the high density thin-wall material in addition to advancements of the supply chain. NASA has also completed fabrication of several subscale and full-scale channel wall nozzles in LP-DED NASA HR-1 and completed hot-fire testing. This includes refinement of the process to produce thin-walls and various channel geometries to meet the requirements for channel wall nozzle applications. This paper will provide an overview of the LP-DED process development, material characterization and properties, component manufacturing, and hot-fire testing. Hot-fire testing was completed for a lander-class 7K-lbf thrust chamber using Liquid Oxygen (LOX)/Methane. The design overview and results from hot-fire testing will be presented in addition to hardware development for future testing on 2K-lbf and 35k-lbf thrust chambers and large-scale manufacturing technology demonstrators.

Additive Manufacturing↗

NASA NTRSDate not supplied

Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth’s atmosphere and non-Earth atmospheres. The focus of this presentation is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. The presentation will start with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the “insulated airplane” approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the presentation will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and will be discussed briefly.

David E Glass↗