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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Advanced small rocket chambers. Option 3: 110 1bf Ir-Re rocket, volume 2

This is the second part of a two-part report that describes the AJ10-221, a high performance iridium-coated rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000 F) (2200 C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units were welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated which, when proven to be capable of withstanding launch vibration, is ready for flight qualification.

Jassowski, Donald M.↗
NASA NTRS1999

Long-Term Cyclic Oxidation Behavior of Uncoated and Coated Re-108 and In-939 at 980 and 870 C

Very long-term cyclic oxidation behavior of Re-108 and ln-939 with and without a protective coating was evaluated at 980 and 870 C, respectively. Re-108 and ln-939 without a protective coating began to show rapid weight loss at 3000 h due to scale spallation, indicating the need for an oxidation protective coating for longer than thousands of hours of oxidative life. NiAl-base coatings of a vapor phase aluminide (VPA), a pack aluminide (CODEP), and a slurry paint aluminide (SERMALOY J) were applied on Re-108 and ln-939. VPA and CODEP on Re-108 and all three coatings on ln-939 showed excellent cyclic oxidation resistance out to 10000 hr. Coated alloys were annealed in an inert atmosphere to determine the loss of Al from the coating into the alloy substrate through diffusion. The Al loss from the coating through diffusion was twice as great as the Al loss through oxidation after 10000 h of cyclic exposure. Oxidation life of VPA-coated Re-108 was estimated by calculating the amount of Al initially available for protective oxidation and the amount of Al lost through oxidation and diffusion.

Lee, K. N.↗
NASA NTRS1999

FTIR Study of Vapor Offgassing from Orbiter Tile Re-Waterproofing Materials

The work presented in this paper was performed to identify and quantify the offgassing behavior of alternative re-waterproofmg materials under investigation for application to Shuttle Orbiter Thermal Protection System (TPS) tile and blanket materials. The purpose was to determine whether the new materials would cause a problem with the operational analysis of residual vapors using the current portable vapor analyzer, a Miran 203. The materials investigated were limited to dimethylethoxysilane (DMES) and proposed solvent selected as suitable for use in re-waterproofing Orbiter TPS. The solvent was selected in another phase of the overall project. Obiter TPS tiles were injected with the re-waterproofing materials under constant conditions of temperature, relative humidity and air flow. The vapor concentrations of offgassing materials were monitored using Fourier Transform Infrared (FTIR) multi-component analysis, and with the Miran 203 instruments. The procedure was to record the time dependent concentrations of offgassing materials as analyzed by the FTIR, and the time response of the Miran 203 to the materials under consideration. The FTIR was calibrated for vapor phase DMES, tetramethyldisiloxane (TMDS), ethanol and the hydrocarbon solvents to be used to dilute the DMES for application to the TPS tile. The Miran 203 was calibrated for the operational measurement of DMES airborne vapors. The FTIR data, shows for the first time that the principal product which offgases from the tile after the first hour is not DMES, but TMDS and ethanol. The Miran 203 response to TMDS is the primary reading after the re-waterproofing operation is completed. The operational use of the Miran 203 to measure DMES vapors after re-waterproofmg operations has been responding to TMDS. The results of this study suggest that the historical complaints that have contributed to the low threshold limit value (TLV) for DMES concentrations, as read with the Miran 203, are actually based on instrumental responses to TMDS, for which no other toxicology information is available at this time. While there is some interference, the use of the tested hydrocarbon solvents does not adversely affect the response of the Miran 203.

Mattson, C. B.↗
NASA NTRS2000

Trajectory Design and Control for the Compton Gamma Ray Observatory Re-Entry

The Compton Gamma Ray Observatory (CGRO) controlled re-entry operation was successfully conducted in June of 2000. The surviving parts of the spacecraft landed in the Pacific Ocean within the predicted footprint. The design of the maneuvers to control the trajectory to accomplish this re-entry presented several challenges. These challenges included timing and duration of the maneuvers, fuel management, post maneuver position knowledge, collision avoidance with other spacecraft, accounting for the break-up of the spacecraft into several pieces with a wide range of ballistic coefficients, and ensuring that the impact footprint would remain within the desired landing area in the event of contingencies. This paper presents the initial re-entry trajectory design and the evolution of the design into the maneuver sequence used for the re-entry. The paper discusses the constraints on the trajectory design, the modifications made to the initial design and the reasons behind these modifications. Data from the re-entry operation are presented.

Hoge, Susan↗
NASA NTRS2001

Effectiveness of Diffusion Barrier Coatings for Mo-Re Embedded in C/SiC and C/C

Advanced high-temperature cooling applications may often require the elevated-temperature capability of carbon/silicon carbide or carbon/carbon composites in combination with the hermetic capability of metallic tubes. In this paper, the effects of C/SiC and C/C on tubes fabricated from several different refractory metals were evaluated. Though Mo, Nb, and Re were evaluated in the present study, the primary effort was directed toward two alloys of Mo-Re, namely, arc cast Mo-41Re and powder metallurgy Mo-47.5Re. Samples of these refractory metals were subjected to either the PyC/SiC deposition or embedding in C/C. MoSi2(Ge), R512E, and TiB2 coatings were included on several of the samples as potential diffusion barriers. The effects of the processing and thermal exposure on the samples were evaluated by conducting burst tests, microhardness surveys, and scanning electron microscopic examination (using either secondary electron or back scattered electron imaging and energy dispersive spectroscopy). The results showed that a layer of brittle Mo-carbide formed on the substrates of both the uncoated Mo-41Re and the uncoated Mo-47.5Re, subsequent to the C/C or the PyC/SiC processing. Both the R512E and the MoSi2(Ge) coatings were effective in preventing not only the diffusion of C into the Mo-Re substrate, but also the formation of the Mo-carbides. However, none of the coatings were effective at preventing both C and Si diffusion without some degradation of the substrate.

Glass, David E.↗
NASA NTRS2003

Cardiovascular effects of anti-G suit and cooling garment during space shuttle re-entry and landing

BACKGROUND: Many cardiovascular changes associated with spaceflight reduce the ability of the cardiovascular system to oppose gravity on return to Earth, leaving astronauts susceptible to orthostatic hypotension during re-entry and landing. Consequently, an anti-G suit was developed to protect arterial pressure during re-entry. A liquid cooling garment (LCG) was then needed to alleviate the thermal stress resulting from use of the launch and entry suit. METHODS: We studied 34 astronauts on 22 flights (4-16 d). Subjects were studied 10 d before launch and on landing day. Preflight, crewmembers were suited with their anti-G suits set to the intended inflation for re-entry. Three consecutive measurements of heart rate and arterial pressure were obtained while seated and then again while standing. Three subjects who inflated the anti-G suits also donned the LCG for landing. Arterial pressure and heart rate were measured every 5 min during the de-orbit maneuver, through maximum G-loading (max-G) and touch down (TD). After TD, crew-members again initiated three seated measurements followed by three standing measurements. RESULTS: Astronauts with inflated anti-G suits had higher arterial pressure than those who did not have inflated anti-G suits during re-entry and landing (133.1 +/- 2.5/76.1 +/- 2.1 vs. 128.3 +/- 4.2/79.3 +/- 2.9, de-orbit; 157.3 +/- 4.5/102.1 +/- 3.6 vs. 145.2 +/- 10.5/95.7 + 5.5, max-G; 159.6 +/- 3.9/103.7 +/- 3.3 vs. 134.1 +/- 5.1/85.7 +/- 3.1, TD). In the group with inflated anti-G suits, those who also wore the LCG exhibited significantly lower heart rates than those who did not (75.7 +/- 11.5 vs. 86.5 +/- 6.2, de-orbit; 79.5 +/- 24.8 vs. 112.1 +/- 8.7, max-G; 84.7 +/- 8.0 vs. 110.5 +/- 7.9, TD). CONCLUSIONS: The anti-G suit is effective in supporting arterial pressure. The addition of the LCG lowers heart rate during re-entry.

NASA Center JSC↗
NASA NTRS2008

Ongoing Capabilities and Developments of Re-Entry Plasma Ground Tests at EADS-ASTRIUM

During re-entry, spacecrafts are subjected to extreme thermal loads. On mars, they may go through dust storms. These external heat loads are leading the design of re-entry vehicles or are affecting it for spacecraft facing solid propellant jet stream. Sizing the Thermal Protection System require a good knowledge of such solicitations and means to model and reproduce them on earth. Through its work on European projects, ASTRIUM has developed the full range of competences to deal with such issues. For instance, we have designed and tested the heat-shield of the Huygens probe which landed on Titan. In particular, our plasma generators aim to reproduce a wide variety of re-entry conditions. Heat loads are generated by the huge speed of the probes. Such conditions cannot be fully reproduced. Ground tests focus on reproducing local aerothermal loads by using slower but hotter flows. Our inductive plasma torch enables to test little samples at low TRL. Amongst the arc-jets, one was design to test architecture design of ISS crew return system and others fit more severe re-entry such as sample returns or Venus re-entry. The last developments aimed in testing samples in seeded flows. First step was to design and test the seeding device. Special diagnostics characterizing the resulting flow enabled us to fit it to the requirements.

Jullien, Pierre↗
NASA NTRS2011

STS-134 Re-Rendezvous Design History

In preparation to provide the capability for the Orion spacecraft to rendezvous with the International Space Station (ISS), a new suite of relative navigation sensors are in development and will be tested on one of the final Space Shuttle missions to ISS. The National Aeronautics and Space Administration (NASA) commissioned a flight test of prototypes of the instruments on STS-134, in order to test their performance in the space environment during the nominal rendezvous and docking, as well as a re-rendezvous dedicated to testing the prototype sensors following the undocking of the Space Shuttle Orbiter at the end of the mission. Unlike the initial rendezvous and docking, the re-rendezvous profile would replicate the newly designed Orion coelliptic approach trajectory, something never before attempted with the Shuttle Orbiter. Therefore, there were a number of new parameters that needed to be conceived of, designed, and tested for this re-rendezvous to make the flight test successful. And all of this work had to be integrated with the normal operations of the ISS and Shuttle and had to conform to the constraints of the mission and vehicles. The result of this work is a separation and re-rendezvous trajectory design that will prove not only the design of the relative navigation sensors for the Orion vehicle, but also will serve as a proof of concept for the Orion rendezvous trajectory itself. This document presents the analysis and decision making process involved in attaining the final STS-134 re-rendezvous design.

Stuit, Timothy D.↗
NASA NTRS2011

Improved Mo-Re VPS Alloys for High-Temperature Uses

Dispersion-strengthened molybdenum- rhenium alloys for vacuum plasma spraying (VPS) fabrication of high-temperature-resistant components are undergoing development. In comparison with otherwise equivalent non-dispersion-strengthened Mo-Re alloys, these alloys have improved high-temperature properties. Examples of VPS-fabricated high-temperature-resistant components for which these alloys are expected to be suitable include parts of aircraft and spacecraft engines, furnaces, and nuclear power plants; wear coatings; sputtering targets; x-ray targets; heat pipes in which liquid metals are used as working fluids; and heat exchangers in general. These alloys could also be useful as coating materials in some biomedical applications. The alloys consist of 60 weight percent Mo with 40 weight percent Re made from (1) blends of elemental Mo and Re powders or (2) Re-coated Mo particles that have been subjected to a proprietary powder-alloying-and-spheroidization process. For most of the dispersion- strengthening experiments performed thus far in this development effort, 0.4 volume percent of transition-metal ceramic dispersoids were mixed into the feedstock powders. For one experiment, the proportion of dispersoid was 1 volume percent. In each case, the dispersoid consisted of either ZrN particles having sizes <45 m, ZrO2 particles having sizes of about 1 m, HfO2 particles having sizes <45 m, or HfN particles having sizes <1 m. These materials were chosen for evaluation on the basis of previously published thermodynamic stability data. For comparison, Mo-Re feedstock powders without dispersoids were also prepared.

Hickman, Robert↗
NASA NTRS1972

Studies of the performance of W-Re type thermocouples.

The effect of exposure of bare-wire and BeO-insulated commercial W-Re thermocouple materials to high temperatures in gaseous environments was investigated. The temperature range of interest was primarily 2000 to 2400 K, and the investigations were confined to thermocouple wires of 0.25 mm diameter. Thermoelements were examined for metallurgical structural changes and chemical changes by conventional methods. In the temperature range of interest, grain growth was inhibited in the chemically doped W-3% Re alloy, and excellent room-temperature ductility was retained subsequent to the exposure. The compatibility of high purity (in excess of 99.8%), sintered BeO insulators with the thermoelements differed, depending upon whether the BeO-insulated thermoelement assemblies were self-heated electrically or heated in a furnace; in tests in argon environments, highly reliable performance occurred when the assemblies were heated in a furnace. BeO-insulated W-3% Re vs W-25% Re thermocouples, constructed with degassed and aged materials, exhibited drifts equivalent to about 3 mK/hr during 1000 hours of exposure at 2073 K in argon while in the presence of tantalum.

Burns, G. W.↗
NASA NTRS1995

RE J1255+266: Detection of an extremely bright EUV transient

During a pointed ROSAT observation in the direction of the Coma cluster of galaxies an exceptionally bright EUV source, RE J1255+266, was detected serendipitously. The source is located close to the Galactic North Pole, at b(sub II) is approximately or equal to 89 deg. Its observed EUV flux (62 - 110 eV) at the time of the detection was of order 7 x 10(exp -9) ergs/sq cm, making RE J1255+266 temporarily one of the brightest EUV sources on the sky. The EUV flare of RE J1255+266 has a light curve with a decay time of about 0.86 days. With respect to earlier non-detections, the source brightened by a factor of approximately 7000. Such a behavior has not been observed before. Thus, it is unclear what type of source RE J1255+266 might be. Up to now no positive identification with any known source could be obtained. Emission at the position of the source was previously only detected in the 1987 Green Bank radio continuum survey. Simultaneous observations with Compton Gamma Ray Observatory/Burst and Transient Source Experiment (CGRO/BATSE) resulted in non-detections of the source in the 8 - 50 keV energy range.

Dahlem, M.↗
NASA NTRS1995

Re-scheduling as a tool for the power management on board a spacecraft

The scheduling of events on board a spacecraft is based on forecast energy levels. The real time values of energy may not coincide with the forecast values; consequently, a dynamic revising to the allocation of power is needed. The re-scheduling is also needed for other reasons on board a spacecraft like the addition of new event which must be scheduled, or a failure of an event due to many different contingencies. This need of rescheduling is very important to the survivability of the spacecraft. In this presentation, a re-scheduling tool will be presented as a part of an overall scheme for the power management on board a spacecraft from the allocation of energy point of view. The overall scheme is based on the optimal use of energy available on board a spacecraft using expert systems combined with linear optimization techniques. The system will be able to schedule maximum number of events utilizing most energy available. The outcome is more events scheduled to share the operation cost of that spacecraft. The system will also be able to re-schedule in case of a contingency with minimal time and minimal disturbance of the original schedule. The end product is a fully integrated planning system capable of producing the right decisions in short time with less human error. The overall system will be presented with the re-scheduling algorithm discussed in detail, then the tests and results will be presented for validations.

Albasheer, Omar↗
NASA NTRS2000

Long-Term Cyclic Oxidation Behavior of Uncoated and Coated Re-108 and In-939 at 980 and 870 C

Very long-term cyclic oxidation behavior of Re108 and In939 with and without a protective coating was evaluated at 980 and 870 C, respectively. Re-108 and In-939 without a protective coating began to show a rapid weight loss at 3000 h due to scale spallation, indicating the need for an oxidation protective coating for longer than thousands of hours of oxidative life. NiAl-base coatings of a vapor phase aluminide (VPA), a pack aluminide (CODEP), and a slurry paint aluminide (SERMALOY J) were applied on Re-108 and In-939. The VPA and CODEP on Re-108 and all three coatings on In-939 showed excellent cyclic oxidation resistance out to 10,000 h. Coated alloys were annealed in an inert atmosphere to determine the loss of Al from the coating into the alloy substrate through diffusion. The Al loss from the coating through diffusion was twice as great as the Al loss through oxidation after 10,000 h of cyclic exposure. The oxidation life of VPA-coated Re-108 was estimated by calculating the amount of Al initially available for protective oxidation and the amount of Al lost through oxidation and diffusion.

Lee, K. N.↗
NASA NTRS2001

Trajectory Design and Control for the Compton Gamma Ray Observatory Re-Entry

The Compton Gamma Ray Observatory (CGRO) controlled re-entry operation was successfully conducted in June of 2000. The surviving parts of the spacecraft landed in the Pacific Ocean within the nominal impact target zone. The design of the maneuvers to control the trajectory to accomplish this re-entry presented several challenges. These challenges included the timing and duration of the maneuvers, propellant management, post-maneuver state determination, collision avoidance with other spacecraft, accounting for the break-up of the spacecraft into several pieces with a wide range of ballistic coefficients, and ensuring that the impact footprint would remain within the desired impact target zone in the event of contingencies. This paper presents the initial re-entry trajectory design and traces the evolution of that design into the maneuver sequence used for the re-entry. The paper also discusses the spacecraft systems and operational constraints imposed on the trajectory design and the required modifications to the initial design based on those constraints. Data from the reentry operation are also presented.

Hoge, Susan↗
NASA NTRS2002

Investigations of Control Surface Seals for Re-entry Vehicles

Re-entry vehicles generally require control surfaces (e.g., rudders, body flaps) to steer them during flight. Control surface seals are installed along hinge lines and where control surface edges move close to the vehicle body. These seals must operate at high temperatures and limit heat transfer to underlying structures to prevent them from overheating and causing possible loss of vehicle structural integrity. This paper presents results for thermal analyses and mechanical testing conducted on the baseline rudder/fin seal design for the X-38 re-entry vehicle. Exposure of the seals in a compressed state at the predicted peak seal temperature of 1900 F resulted in loss of seal resiliency. The vertical Inconel rudder/fin rub surface was re-designed to account for this loss of resiliency. Room temperature compression tests revealed that seal unit loads and contact pressures were below limits set to protect Shuttle thermal tiles on the horizontal sealing surface. The seals survived an ambient temperature 1000 cycle scrub test over sanded Shuttle tiles and were able to disengage and re-engage the tile edges during testing. Arc jet tests confirmed the need for seals in the rudder/fin gap location because a single seal caused a large temperature drop (delta T = 1710 F) in the gap.

Dunlap, Patrick H., Jr.↗
NASA NTRS2003

Re-Os Systematics and HSE Distribution in Metal from Ochansk (H4) Chondrite

Previous studies of the Re-Os systematics of chondrites have documented considerable variation in the Re/Os ratios of whole rock samples. For some whole rock chondrites, Re-Os systematics display large deviations from the primitive isochron that are considerably larger than deviations in other isotope systems. Possible interpretation of these facts is that the Re-Os system in chondrites is particularly sensitive to post-formation alteration processes, thus providing a useful tool to examine such processes. Significant variations that have been detected in highly siderophile element (HSE) patterns for ordinary chondrites support this conclusion. We report Re-Os isotope data for metal separates from the Ochansk H4 chondrite coupled with abundance data for Ru, Pd, Ir, and Pt, determined in the same samples by isotope dilution. We chose this meteorite mainly because it is an observed fall with minimal signs of weathering, and its low metamorphic grade (H4) and shock stage (S3).

Smoliar, M. I.↗
NASA NTRS2004

NASA GEOS-3/TRMM Re-analysis: Capturing Observed Tropical Rainfall Variability in Global Analysis for Climate Research

Understanding climate variability over a wide range of space-time scales requires a comprehensive description of the earth system. Global analyses produced by a fixed assimilation system (i.e., re-analyses) - as their quality continues to improve - have the potential of providing a vital tool for meeting this challenge. But at the present time, the usefulness of re-analyses is limited by uncertainties in such basic fields as clouds, precipitation, and evaporation - especially in the tropics, where observations are relatively sparse. Analyses of the tropics have long been shown to be sensitive to. the treatment of cloud precipitation processes, which remains a major source of uncertainty in current models. Yet, for many climate studies it is crucial that analyses can accurately reproduce the observed rainfall intensity and variability since a small error of 1 mm/d in surface rain translates into an error of approx. 30 W/sq m in energy (latent heat) flux. Currently, discrepancies between the observed and analyzed monthly-mean rain rates averaged to 100 km x 100 km resolution can exceed 4 mm/d (or 120 W/sq m ), compared to uncertainties in surface radiative fluxes of approx. 10-20 W/sq m . Improving precipitation in analyses would reduce a major source of uncertainty in the global energy budget. Uncertainties in tropical precipitation have also been a major impediment in understanding how the tropics interact with other regions, including the remote response to El Nino/Southern Oscillation (ENSO) variability on interannual time scales, the influence of Madden-Julian Oscillation (MJO) and monsoons on intraseasonal time scales. A global analysis that can replicate the observed precipitation variability together with physically consistent estimates of other atmospheric variables provides the key to breaking this roadblock. NASA Goddard Space Flight Center has been exploring the use of satellite-based microwave rainfall measurements in improving global analyses and has recently produced a multi-year, 1 x 1 TRMM re-analysis , which assimilates 6-hourly TMI and SSM/I surface rain rates over tropical oceans using a ID variational continuous assimilation (VCA) procedure in the GEOS-3 global data assimilation system. The analysis period extends from 1 November 1997 through 3 1 December 2002. The goal is to produce a multi-year global analysis that is dynamically consistent with available tropical precipitation observations for the community to assess its utility in climate applications and identify areas for further improvements. A distinct feature of the GEOS-3RRMh4 re-analysis is that its precipitation analysis is not derived from a short-term forecast (as done in most operational systems) but is given by a time- continuous model integration constrained by precipitation observations within a 6-h analysis window, while the wind, temperature, and pressure fields are allowed to directly respond to the improved precipitation and associated latent heating structures within the same analysis window. In this talk, I will assess the impact VCA precipitation assimilation on analyses of climate signals ranging from a few weeks to interannual time scales and compare results against other operational and reanalysis products.

Hou, Arthur Y.↗
NASA NTRS2002

Rudder/Fin Seals Investigated for the X-38 Re-Entry Vehicle

NASA is developing the X-38 vehicle that will demonstrate the technologies required for a potential crew return vehicle for the International Space Station. This vehicle would serve both as an ambulance for medical emergencies and as an evacuation vehicle for the space station. Control surfaces on the X-38 (body flaps and rudder/fin assemblies) require high temperature seals to limit hot gas ingestion and the transfer of heat to underlying low temperature structures. Working with the NASA Johnson Space Center, the Seals Team at the NASA Glenn Research Center completed a series of tests to further characterize baseline seal designs for the rudder/fin interfaces of the X-38. The structures of the rudder/fin assembly and its associated seals are shown in the the preceding illustration. Tests performed at Glenn indicated that exposure of the seals in a compressed state at simulated seal re-entry temperatures resulted in a large permanent set and loss of seal resiliency. This could be of concern because the seals are required to maintain contact with the sealing surfaces while the vehicle goes through the maximum re-entry heating cycle to prevent hot gases from leaking past the seals and damaging interior low-temperature structures. To simulate conditions in which the seals may become unloaded during use, such as when they take on a large permanent set, Glenn researchers performed room temperature flow and compression tests to determine seal flow rates, resiliency, and unit loads under minimal loads. Flow rates through an unloaded (i.e., 0-percent compression) double seal arrangement were twice those of a double seal compressed to the 20-percent design compression level. These flow rates are being used in thermal analyses to predict the effect of flow through the seals on over-all seal temperatures. Compression test results showed that seal unit loads and contact pressures were below the limits that Johnson had set as goals for the seals. In the rudder/fin seal location, the seals are in contact with shuttle thermal tiles and are moved across the tiles as the rudder is rotated during re-entry. Low seal unit loads and contact pressures are required to limit the loads on these tiles and minimize any damage that the seals could cause. A series of tests were performed on these seals in NASA Ames Research Center's arc jet facility. The arc jet facility approximates relevant thermal environments that a seal or other structure would be subjected to during extreme heating conditions such as those experienced during space vehicle re-entry. Eleven tests were completed, including one test in which no seal was installed in the gap to examine the flow of heat down into the gap. The seal was compressed between stationary insulation tiles and a movable elevon that was rotated during the test to deflect the arc jet exhaust into the seal gap. Peak seal temperatures as high as 2000 F were reached during the 5-min tests. Results of these tests indicate satisfactory performance of the seal for single-use (e.g., X-38) applications. The results of these tests were shared with the NASA Johnson Space Center and are being used to validate aerothermostructural analysis codes that predict seal temperatures under these conditions. The tests performed at Glenn have provided valuable information to Johnson about the performance of the seals that they are considering using in the rudder/fin location of the X-38 vehicle. Glenn and Johnson are currently defining what additional work needs to be done to develop the final rudder/fin seal design for the X-38 vehicle.

Dunlap, Patrick H., Jr.↗