Engineering topics

Curry, Donald M.

Publications and source records attributed to Curry, Donald M..

At least 19 records

NASA NTRS2008

Oxidation Through Coating Cracks of SiC-Protected Carbon/Carbon

The oxidation of SiC-protected carbon/carbon through machined slots and naturally occurring craze cracks in the SiC was studied. The slot and crack geometries were characterized, and the subsurface oxidation of the carbon/carbon substrate at temperatures of 1000 to 1300 C in air was assessed using weight change, x-ray computed tomography, and optical microscopy of sections. Rate constants were derived from these measurements and compared with a two-step diffusion control model of carbon oxidation. Oxidation kinetic measurements on both the specimens with machined slots and with naturally occurring craze cracks showed good agreement with the model.

Jacobson, Nathan S.↗

NASA NTRS2007

Oxidation Through Coating Cracks of SiC-Protected Carbon/Carbon

Jacobson, Nathan S.↗

NASA NTRS2007

Oxidation Through Coating Cracks of SiC-Protected Carbon/Carbon

Jacobson, Nathan S.↗

NASA NTRS2006

Oxidation Microstructure Studies of Reinforced Carbon/Carbon

Laboratory oxidation studies of reinforced carbon/carbon (RCC) are discussed with particular emphasis on the resulting microstructures. This study involves laboratory furnace (500-1500 C deg) and arc-jet exposures (1538 C deg) on various forms of RCC. RCC without oxidation protection oxidized at 800 and 1100 C deg exhibits pointed and reduced diameter fibers, due to preferential attack along the fiber edges. RCC with a SiC conversion coating exhibits limited attack of the carbon substrate at 500, 700 and 1500 C deg. However samples oxidized at 900, 1100, and 1300 C deg show small oxidation cavities at the SiC/carbon interface below through-thickness cracks in the SiC coating. These cavities have rough edges with denuded fibers and can be easily distinguished from cavities created in processing. Arc-jet tests at 1538 C deg show limited oxidation attack when the SiC coating and glass sealants are intact. When the SiC/sealant protection system is damaged, attack is extensive and proceeds through matrix cracks, creating denuded fibers on the edges of the cracks. Even at 1538 C deg, where diffusion control dominates, attack is non-uniform with fiber edges oxidizing in preference to the bulk fiber and matrix.

Jacobson, Nathan S.↗

NASA NTRS2005

Family System of Advanced Charring Ablators for Planetary Exploration Missions

Advanced Ablators Program Objectives: 1) Flight-ready(TRL-6) ablative heat shields for deep-space missions; 2) Diversity of selection from family-system approach; 3) Minimum weight systems with high reliability; 4) Optimized formulations and processing; 5) Fully characterized properties; and 6) Low-cost manufacturing. Definition and integration of candidate lightweight structures. Test and analysis database to support flight-vehicle engineering. Results from production scale-up studies and production-cost analyses.

Congdon, William M.↗

NASA NTRS2004

TPS Ablator Technologies for Interplanetary Spacecraft

This slide presentation reviews the status of Thermal Protection System (TPS) Ablator technologies and the preparation for use in interplanetary spacecraft. NASA does not have adequate TPS ablatives and sufficient selection for planned missions. It includes a comparison of shuttle and interplanetary TPS requirements, the status of mainline TPS charring ablator materials, a summary of JSC SBIR accomplishments in developing advanced charring ablators and the benefits of SBIR Ablator/fabrication technology.

Curry, Donald M.↗

NASA NTRS2003

Response Modeling of Lightweight Charring Ablators and Thermal Radiation Testing Results

Under NASA's In-Space Propulsion/Aerocapture Program, ARA conducted arc-jet and thermal-radiation ablation test series in 2003 for advanced development, characterization, and response modeling of SRAM-20, SRAM-17, SRAM-14, and PhenCarb-20 ablators. Testing was focused on the future Titan Explorer mission. Convective heating rates (CW) were as high as 153 W/sq cm in the IHF and radiation rates were 100 W/sq cm in the Solar Tower Facility. The ablators showed good performance in the radiation environment without spallation, which was initially a concern, but they also showed higher in-depth temperatures when compared to analytical predictions based on arc-jet thermal-ablation response models. More testing in 2003 is planned in both of these facility to generate a sufficient data base for Titan TPS engineering.

Congdon, William M.↗

NASA NTRS2002

X-38 Seal Development

An X-38 Crew Return Vehicle Seal Development is presented. The contents include: 1) X-38 Crew Return Vehicle; 2) X-38 TPS Configuration; 3) X-38 Seal Locations; 4) X-38 Rudder/Fin Seal Assembly; 5) Baseline X-38 Rudder/Fin Seal Design; 6) Rudder/Fin Seal to Bracket Assembly; 7) X-38 Rudder/Fin Vertical Rub Surface Inconel-0.10 inches; 8) X-38 Rudder/Fin Seal Analysis; 9) Seal Analysis Model; and 10) Governing Differential Equations for Equilibrium Thermal Assumption. The X-38 Rudder/Fin Seal temperature and pressure properties are also given.

Curry, Donald M.↗

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 NTRS2001

X-38 Seal Development

This viewgraph presentation provides information on the development of seals for the X-38 spacecraft. This seals will be an important part of the X-38's thermal protection system (TPS) in its service as a crew return vehicle (CRV) associated with the International Space Station (ISS).

Curry, Donald M.↗

NASA NTRS2001

Rudder/Fin Seal Investigations for the X-38 Re-Entry Vehicle

NASA is currently developing the X-38 vehicle that will be used to demonstrate the technologies required for a crew return vehicle (CRV) for the International Space Station (ISS). The CRV will 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 rudders/fins) require high temperature seals to limit hot gas ingestion and transfer of heat to underlying low-temperature structures to prevent over-temperature of these structures and possible loss of the vehicle. NASA's Johnson Space Center (JSC) and Glenn Research Center (GRC) are working together to develop and evaluate seals for the rudder/fin control surfaces. The specific objectives of this study are to measure seal flow rates, resiliency, and unit loads in as-received and temperature-exposed conditions and compare the measured results to property goals where applicable. Areas for future work would then be identified.

Dunlap, Patrick H., Jr.↗

NASA NTRS2001

Further Investigations of Control Surface Seals for the X-38 Re-Entry Vehicle

NASA is currently developing the X-38 vehicle that will be used to demonstrate the technologies required for a potential crew return vehicle (CRV) 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 transfer of heat to underlying low-temperature structures to prevent over-temperature of these structures and possible loss of the vehicle. NASAs Johnson Space Center (JSC) and Glenn Research Center (GRC) are working together to develop and evaluate seals for these control surfaces. This paper presents results for compression. flow, scrub, and arc jet tests conducted on the baseline X-38 rudder/fin seal design. Room temperature seal compression tests were performed at low compression levels to determine load versus linear compression, preload. contact area, stiffness. and resiliency characteristics under low load conditions. For all compression levels that were tested, unit loads and contact pressures for the seals were below the 5 lb/in. and 10 psi limits required to limit the loads on the adjoining Shuttle thermal tiles that the seals will contact. Flow rates through an unloaded (i.e. 0% compression) double arrangement were twice those of a double seal compressed to the 20% design compression level. The seals survived an ambient temperature 1000 cycle scrub test over relatively rough Shuttle tile surfaces. The seals were able to disengage and re-engage the edges of the rub surface tiles while being scrubbed over them. Arc jet tests were performed to experimentally determine anticipated seal temperatures for representative flow boundary conditions (pressures and temperatures) under simulated vehicle re-entry conditions. Installation of a single seat in the gap of the test fixture caused a large temperature drop (1710 F) across the seal location as compared to an open gap condition (140 F) confirming the need for seals in the rudder/fin gap location. The seal acted as an effective thermal barrier limiting heat convection through the seal gap and minimizing temperature increases downstream of the seal during maximum heating conditions.

Dunlap, Patrick H., Jr.↗

NASA NTRS2000

Rudder/Fin Seal Investigations for the X-38 Re-Entry Vehicle

NASA is currently developing the X-38 vehicle that will be used to demonstrate the technologies required for a crew return vehicle (CRV) for the International Space Station. The X-38 control surfaces require high temperature seals to limit hot gas ingestion and transfer of heat to underlying low-temperature structures to prevent over-temperature of these structures and possible loss of the vehicle. This paper presents results for thermal analyses and flow and compression tests conducted on as-received and thermally exposed seals for the rudder/fin location of the X-38. A thermal analysis of the rudder/fin dual seal assembly based on representative heating rates on the windward surface of the rudder/fin area predicted a peak seal temperature of 1900 F. The temperature-exposed seals were heated in a compressed state at 1900 F corresponding to the predicted peak temperature. Room temperature compression tests were performed to determine load versus linear compression, preload, contact area, stiffness, and resiliency characteristics for the as-received and temperature-exposed seals. Temperature exposure resulted in permanent set and loss of resiliency in these seals. Unit loads and contact pressures for the seals were below the five pounds/inch and ten psi limits set to limit the loads on the Shuttle thermal tiles that the seals seal against in the rudder/fin location. Measured seal flow rates for a double seal were about 4.5 times higher than the preliminary seal flow goal. The seal designs examined in this study are expected to be able to endure the high temperatures that they will be exposed to for a single-use life. Tests performed herein combined with future analyses, arc jet tests, and scrubbing tests will be used to select the final seal design for this application.

Dunlap, Patrick H., Jr.↗

NASA NTRS2000

X-38 TPS Seal Status

This presentation discuss the x-38 crew return vehicle. As an element of the International Space Station (ISS), there are potential problems that are discussed. These include ISS catastrophe, emergency medical evacuation, and period of Space Shuttle unavailability. The x-38 program purpose was also discussed. The Reduction of the costs and schedule for the development of Crew Return Vehicles (CRV's) and Crew Transfer Vehicles (CTV's) through the use of the rapid development methodology associated with an X-project were also presented. With specific attention to ground testing, atmospheric testing, and space flight testing.

Curry, Donald M.↗

NASA NTRS2000

Rudder/Fin Seal Investigations for the X-38 Re-Entry Vehicle

NASA is currently developing the X-38 vehicle that will be used to demonstrate the technologies required for a crew return vehicle (CRV) for the International Space Station. The X-38 control surfaces require high temperature seals to limit hot gas ingestion and transfer of heat to underlying low-temperature structures to prevent over-temperature of these structures and possible loss of the vehicle. This paper presents results for thermal analyses and flow and compression tests conducted on as-received and thermally exposed seals for the rudder/fin location of the X-38. A thermal analysis of the rudder/fin dual seal assembly based on representative heating rates on the windward surface of the rudder/fin area predicted a peak seal temperature of 1900 F. The temperature-exposed seals were heated in a compressed state at 1900 F corresponding to the predicted peak temperature. Room temperature compression tests were performed to determine load versus linear compression, preload, contact area, stiffness, and resiliency characteristics for the as-received and temperature-exposed seals. Temperature exposure resulted in permanent set and loss of resiliency in these seals. Unit loads and contact pressures for the seals were below the 5 lb/in. and 10 psi limits set to limit the loads on the Shuttle thermal tiles that the seals seal against in the rudder/fin location. Measured seal flow rates for a double seal were about 4.5 times higher than the preliminary seal flow goal. The seal designs examined in this study are expected to be able to endure the high temperatures that they will be exposed to for a single-use life. Tests performed herein combined with future analyses, arc jet tests, and scrubbing tests will be used to select the final seal design for this application.

Dunlap, Patrick H., Jr.↗

NASA NTRS2000

Oxidation of Reinforced Carbon-Carbon Subjected to Hypervelocity Impact

This paper presents results from arc jet tests conducted at the NASA Johnson Space Center on reinforced carbon-carbon (RCC) samples subjected to hypervelocity impact. The RCC test specimens are representative of RCC components used on the Space Shuttle Orbiter. The arc jet testing established the oxidation characteristics of RCC when hypervelocity projectiles, simulating meteoroid/orbital debris, impact the RCC material. In addition to developing correlations for use in trajectory simulations, we discuss analytical modeling of the increased material oxidation in the impacted area using measured hole growth data. Entry flight simulations are useful in assessing the increased Space Shuttle RCC component degradation as a result of impact damage and the hot gas flow through an enlarging hole into the wing leading-edge cavity.

Curry, Donald M.↗

NASA NTRS2000

Orbiter Reinforced Carbon-Carbon Advanced Sealant Systems: Screening Tests

Oxidation protection for the Orbiter reinforced carbon-carbon (RCC consists of three components: silicon carbide coating, tetraethyl orthosilicate (TEOS) impregnated into the carbon substrate and a silicon based surface sealant (designated Type A). The Orbiter Type A sealant is being consumed each mission, which results in increased carbon-carbon substrate mass loss, which adversely impacts the mission life of the RCC components. In addition, the sealant loss in combination with launch pad contamination (salt deposit and zinc oxide) results in RCC pinholes. A sealant refurbishment schedule to maintain mission life and minimize affects of pin hole formation has been implemented in the Orbiter maintenance schedule. The objective of this investigation is to develop an advanced sealant system for the RCC that extends the refurbishment schedule by reducing sealant loss/pin hole formation and that can be applied to existing Orbiter RCC components. This paper presents the results of arc jet screening tests conducted on several sealants that are being considered for application to the Orbiter RCC.

Curry, Donald M.↗

NASA NTRS1999

Oxidation of Reinforced Carbon-Carbon Subjected to Hypervelocity Impact

Results of arc-jet tests conducted at the NASA Johnson Space Center (JSC) on Reinforced Carbon-Carbon (RCC) samples subjected to hypervelocity impact are presented. The RCC test specimens are representative of RCC component used on the Space Shuttle Orbiter. The objective of the arc jet testing was to establish the oxidation characteristics of RCC when hypervelocity projectiles, simulating meteoroid/orbital debris (MOD), impact the RCC material. In addition, analytical modeling of the increased material oxidation in the impacted area, using measured hole growth data, to develop correlations for use in trajectory simulations is also discussed.

Curry, Donald M.↗