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At least 703 records · Page 39
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Improvements to Thermal Protection System Design of Aerocapture Systems for Uranus Orbiters

The National Academies Planetary Science and Astrobiology Decadal Survey identified Uranus and Neptune - called Ice Giants - as the priority destinations for science. The survey assessed both a mission to Uranus through the Uranus Orbiter and Probe (UOP) concept, and Neptune through the Neptune-Triton Odyssey concept and determined that Uranus is the highest priority for a Flagship class mission. The UOP mission concept planned to deliver an in situ probe and conduct a multi-year orbital tour of the system to meet the science objectives. While the Uranus mission is currently viable with launch windows starting in 2031 using existing launch vehicles, the mission has a cruise phase of at least 12 years and would require more than half of its weight in fuel propellant to achieve the change in velocity necessary for orbital insertion. Aerocapture uses aerodynamic forces generated on a vehicle by the planet's atmosphere to modulate a spacecraft's trajectory, decreasing spacecraft velocity, and allowing mission designers to target the final orbital state. Aerocapture reduces the time-of-flight from Earth to Uranus over a fully propulsive solution, opening up more launch opportunities to arrive in the 2040's to the mission's science objectives. Aerocapture also allows a payload mass increase by mitigating the need for fuel to retropropulsively insert the payload into orbit, thereby increase the science that can be performed. For an aerocapture mission structure using a traditional aeroshell to deliver the UOP scientific payload to Uranus, Conformal Phenolic Impregnated Carbon Ablator (C-PICA) was determined to be the best-performing forebody thermal protection system (TPS) and other candidate aftbody TPS options were presented as feasible. This paper focuses on A) evaluating C-PICA as a forebody TPS using stressing entry conditions associated with a large range of potential Uranus flagship launch vehicles and interplanetary trajectories, B) widening the aftbody TPS candidates for new mass-efficient and cost-efficient solutions, and C) a summary list of actions remaining to provide a technically feasible and supply-robust set of TPS for an aerocapture vehicle to the Ice Giants is presented.

C-PICA↗
NASA NTRSDate not supplied

Improvements to Thermal Protection System Design of Aerocapture Systems for Uranus Orbiters

The National Academies Planetary Science and Astrobiology Decadal Survey identified Uranus and Neptune - called Ice Giants - as the priority destinations for science. The survey assessed both a mission to Uranus through the Uranus Orbiter and Probe (UOP) concept, and Neptune through the Neptune-Triton Odyssey concept and determined that Uranus is the highest priority for a Flagship class mission. The UOP mission concept planned to deliver an in situ probe and conduct a multi-year orbital tour of the system to meet the science objectives. While the Uranus mission is currently viable with launch windows starting in 2031 using existing launch vehicles, the mission has a cruise phase of at least 12 years and would require more than half of its weight in fuel propellant to achieve the change in velocity necessary for orbital insertion. Aerocapture uses aerodynamic forces generated on a vehicle by the planet's atmosphere to modulate a spacecraft's trajectory, decreasing spacecraft velocity, and allowing mission designers to target the final orbital state. Aerocapture reduces the time-of-flight from Earth to Uranus over a fully propulsive solution, opening up more launch opportunities to arrive in the 2040's to the mission's science objectives. Aerocapture also allows a payload mass increase by mitigating the need for fuel to retropropulsively insert the payload into orbit, thereby increase the science that can be performed. For an aerocapture mission structure using a traditional aeroshell to deliver the UOP scientific payload to Uranus, Conformal Phenolic Impregnated Carbon Ablator (C-PICA) was determined to be the best-performing forebody thermal protection system (TPS) and other candidate aftbody TPS options were presented as feasible. This paper focuses on A) evaluating C-PICA as a forebody TPS using stressing entry conditions associated with a large range of potential Uranus flagship launch vehicles and interplanetary trajectories, B) widening the aftbody TPS candidates for new mass-efficient and cost-efficient solutions, and C) a summary list of actions remaining to provide a technically feasible and supply-robust set of TPS for an aerocapture vehicle to the Ice Giants is presented.

C-PICA↗
NASA NTRSDate not supplied

Thermal Protection System Design of Aerocapture Systems for Uranus Orbiters

The National Academies Planetary Science and Astrobiology Decadal Survey recently identified Uranus and Neptune - called Ice Giants - as the priority destinations for science[1]. The survey assessed both a mission to Uranus through the Uranus Orbiter and Probe (UOP) concept, and Neptune through the Neptune-Triton Odyssey concept and determined that Uranus is the highest priority for a Flagship class mission. The UOP mission will deliver an in situ probe and conduct a multi-year orbital tour of the system to meet the science objectives. While the Uranus mission is currently viable with launch windows starting in 2031 using existing launch vehicles, the mission has a long cruise time to destination (between 12 and 15 years) and would require more than half of its weight in fuel propellant to achieve the change in velocity necessary for orbital insertion. Aerocapture is a method of orbital control that uses aerodynamic forces generated on a vehicle by the planet’s atmosphere to modulate a spacecraft’s trajectory, allowing mission designers to target the final orbital state. For the Uranus mission, using aerocapture for orbital insertion can decrease not only the cruise time to the destination by 2 - 3 years, but the propellant required to achieve orbital insertion (by more than 40%) which would, in turn, increase the available science payload and reduce the timeline for retrieving data vital to the mission’s science objectives[2]. Achieving orbital insertion via aerocapture requires novel algorithms for Guidance, Navigation and Control[3], and mass-efficient Thermal Protection Systems (TPS) performing in a new atmosphere. This paper will focus on the selection and tailoring of the Thermal Protection Systems for the forebody and aftbody heatshields of an aerocapture mission to Uranus. While preliminary results indicate that multiple systems in NASA’s repertoire are capable of performing in the predicted aerothermal environment there are unique aspects like the inert environment that affect ablation efficiency, and the heatload for aerocapture trajectories to the outer planets are among the highest of any mission to-date[4]. These two factors may impose operational requirements to heatshield separation in order to minimize thermal soak to the payload, and may demand TPS thickness and configurations that have not yet been demonstrated. This paper will discuss the updated maturity, manufacturing, and performance capabilities of candidate thermal protection systems, with specific areas of need highlighted to make thermal protection systems viable for use in the recommended Uranus Orbiter and Probe mission.

Uranus↗
NASA NTRS1991

Lunar mission design using Nuclear Thermal Rockets

The NERVA-class Nuclear Thermal Rocket (NTR), with performance nearly double that of advanced chemical engines, has long been considered an enabling technology for human missions to Mars. NTR engines address the demanding trip time and payload delivery needs of both cargo-only and piloted flights. But NTR can also reduce the Earth launch requirements for manned lunar missions. First use of NTR for the Moon would be less demanding and would provide a test-bed for early operations experience with this powerful technology. Study of application and design options indicates that NTR propulsion can be integrated with the Space Exploration Initiative scenarios to deliver performance gains while managing controlled, long-term disposal of spent reactors to highly stable orbits.

Stancati, Michael L.↗
NASA NTRS1990

Remtech SSME nozzle design TPS

Thermal damage to the Space Shuttle Main Engine (SSME) aft manifold Thermal Protection System (TPS) has been observed for flights STS-8 through STS-13. This damaged area is located on the ME2 and ME3 and extends over a region of approximately one square foot. Total failure or burn-through of the TPS could lead to severe thermal damage of the SSME manifold and loss of an engine nozzle necessitating nozzle replacement causing significant schedule delays and cost increases. Thermal damage to the manifold can be defined as a situation where the manifold temperature becomes greater than 1300 F; thereby causing loss of heat treatment in the nozzle. Results of Orbiter/nozzle wind tunnel tests and Hot Gas Facility tests of the TPS are presented. Aerothermal and thermal analysis models for the SSME aft manifold are discussed along with the flight predictions, design trajectory and design environment. Finally, the TPS design concept and TPS thermal response are addressed.

Bancroft, Steven A.↗
NASA NTRS1974

Design manual: Oxygen Thermal Test Article (OTTA)

The characteristics of a cryogenic tank for storing liquid hydrogen, nitrogen, oxygen, methane, or helium for an extended period of time with minimum losses are discussed. A description of the tank and control module, assembly drawings and details of major subassemblies, specific requirements controlling development of the system, thermal concept considerations, thermal analysis methods, and a record of test results are provided. The oxygen thermal test article thermal protection system has proven that the insulation system for cryogenic vessels is effective.

Chronic, W. L.↗
NASA NTRS1979

ATLAS, an integrated structural analysis and design system. Volume 6: Design module theory

The automated design theory underlying the operation of the ATLAS Design Module is decribed. The methods, applications and limitations associated with the fully stressed design, the thermal fully stressed design and a regional optimization algorithm are presented. A discussion of the convergence characteristics of the fully stressed design is also included. Derivations and concepts specific to the ATLAS design theory are shown, while conventional terminology and established methods are identified by references.

Backman, B. F.↗
NASA NTRS1984

Thermal and cryogenic design study for space infrared telescope facility (SIRTF)

A study was conducted to determine the ability of an all superfluid helium design to meet the performance requirements of background limited to 200 micrometer, and a two year lifetime for a one meter class free flying infrared observatory. Both a 98 deg and 28.5 deg inclination orbits were examined, and aperture shade designs were developed for both orbits. A unique forebaffle cooling design significantly reduces the sensitivity to aperture heat loads. With certain restrictions on observing modes, the study determined that an all superfluid helium Dewar will meet the temperature and lifetime requirements. A dual cryogen SFHe/SH2 system was also investigated for the 28.5 deg orbit and found to provide a more constant forebaffle temperature but with only a slight improvement in lifetime.

Urbach, A. R.↗
NASA NTRS1981

Manufacturing cost analysis of a parabolic dish concentrator (General Electric design) for solar thermal electric power systems in selected production volumes

The manufacturing cost of a General Electric 12 meter diameter concentrator was estimated. This parabolic dish concentrator for solar thermal system was costed in annual production volumes of 100 - 1,000 - 5,000 - 10,000 - 50,000 100,000 - 400,000 and 1,000,000 units. Presented for each volume are the costs of direct labor, material, burden, tooling, capital equipment and buildings. Also presented is the direct labor personnel and factory space requirements. All costs are based on early 1981 economics.

Source record↗
NASA NTRS1992

Space Station external thermal control system design and operational overview

The driving design requirements and resulting design characteristics for a two-phase ammonia system of the Space Shuttle are reviewed with particular attention given to operational and station assembly issues related to system activation and steady-state operation. Design is described at an overall system level and an orbit replaceable unit level. It is concluded that a system flow network must be designed and ammonia pressures must be controlled to acquire waste heat efficiently from all contributing heat sources at a controlled temperature.

Raetz, John↗
NASA NTRS1999

Thermal Protection System Design and Development for the X-38

The purpose of the X-38 program is to greatly reduce 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 (Ground Testing, Atmospheric Testing and Space Flight Testing).

T John Kowal↗
NASA NTRS2014

Orion MPCV Continuum RCS Heating Augmentation Model Development

The reaction control system jets of the Orion Multi Purpose Crew Vehicle can have a significant impact on the magnitude and distribution of the surface heat flux on the leeside of the aft-body, when they are fired. Changes in surface heating are expressed in terms of augmentation factor over the baseline smooth body heating. Wind tunnel tests revealed heating augmentation factors as high as 13.0, 7.6, 2.8, and 5.8 for the roll, pitch down, pitch up, and yaw jets respectively. Heating augmentation factor models, based almost exclusively on data from a series of wind tunnel tests have been developed, for the purposes of thermal protection system design. The wind tunnel tests investigated several potential jet-to-freestream similarity parameters, and heating augmentation factors derived from the data showed correlation with the jet-to-freestream momentum ratio. However, this correlation was not utilized in the developed models. Instead augmentation factors were held constant throughout the potential trajectory space. This simplification was driven by the fact that ground to flight traceability and sting effects are not well understood. Given the sensitivity of the reaction control system jet heating augmentation to configuration, geometry, and orientation the focus in the present paper is on the methodology used to develop the models and the lessons learned from the data. The models that are outlined in the present work are specific to the aerothermal database used to design the thermal protection system for the Exploration Flight Test 1 vehicle.

Heating Augmentation↗