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NASA NTRSDate not supplied

Development of Laser Powder Bed Fusion NASA HR-2 for Hydrogen Sensitive Liquid Rocket Engine Applications

The National Aeronautics and Space Administration (NASA) has been involved in the development and maturation of metal additive manufacturing (AM) for space applications since the late 2000’s. AM has provided new design and manufacturing opportunities to reduce cost and schedule, consolidate parts, and optimize performance. Laser Powder Bed Fusion (L-PBF) is one of the most commonly used AM processes to fabricate components that have complex shape and need fine feature resolution. Due to exposure to high pressure gaseous hydrogen, mechanical property degradation caused by hydrogen environment embrittlement (HEE) is a critical concern for many materials in liquid hydrogen propulsion systems. NASA has identified the need to develop and advance new materials in unique engine applications using liquid hydrogen as a propellant. One such material being developed at NASA Marshall Space Flight Center is L-PBF NASA HR-2 (Hydrogen Resistant-2), a high-strength Fe-Ni-based superalloy resistant to HEE. The chemistry of NASA HR-2 was formulated to meet requirements for key liquid rocket engine (LRE) components that operate in high-pressure hydrogen environments. Initial development and material characterization found NASA HR-2 has excellent L-PBF printability and its microstructure evolves well after heat treatment. This new alloy has undergone fundamental metallurgical evaluations, heat treatment studies, detailed microstructure characterization, and mechanical testing across a range of temperatures. Tensile testing was performed in pressurized gaseous hydrogen (GH2) environment to assess its resistance to HEE. L-PBF NASA HR-2 has an average yield stress of 95 ksi, ultimate tensile stress of 165 ksi, and very high fracture elongation at 34 - 36% when tested in a 5 ksi high pressure hydrogen environment. The tensile property data confirms hydrogen has little influence on its ductility, strength, and fracture behavior. L-PBF NASA HR-2 is a promising option for many hydrogen sensitive LRE components that require exceptional resistance to HEE. The development of L-PBF NASA HR-2 is funded under the grants provided by Jacobs TIPI program and the Liquid Engine Office at NASA Marshall Space Flight Center. This paper will provide an overview of the L-PBF process development, material characterization, mechanical and thermophysical properties, and LRE hardware development for NASA HR-2.

NASA HR-2↗

NASA NTRSDate not supplied

Development of Laser Powder Bed Fusion NASA HR-2 for Hydrogen Sensitive Liquid Rocket Engine Applications

The National Aeronautics and Space Administration (NASA) has been involved in the development and maturation of metal additive manufacturing (AM) for space applications since the late 2000s. AM has provided new design and manufacturing opportunities to reduce cost and schedule, consolidate parts, and optimize performance. Laser Powder Bed Fusion (L-PBF) is one of the most commonly used AM processes to fabricate components of complex shape requiring fine feature resolution. Due to exposure to high-pressure gaseous hydrogen, mechanical property degradation caused by hydrogen environment embrittlement (HEE) is a critical concern for many materials in liquid hydrogen propulsion systems. NASA has identified the need to develop and advance new materials in unique engine applications using liquid hydrogen as a propellant. One such material being developed at NASA Marshall Space Flight Center is L-PBF NASA HR-2 (Hydrogen Resistant-2), a high-strength Fe-Ni-based superalloy resistant to HEE. The chemistry of NASA HR-2 was formulated to meet requirements for key liquid rocket engine (LRE) components that operate in high-pressure hydrogen environments. Initial development and material characterization found that NASA HR-2 has excellent L-PBF printability, and its microstructure evolves well after heat treatment. This new alloy has undergone fundamental metallurgical evaluations, heat treatment studies, detailed microstructure characterization, and mechanical testing across a range of temperatures. Tensile testing was performed in a pressurized gaseous hydrogen (GH2) environment to assess its resistance to HEE. L-PBF NASA HR-2 has an average yield stress of 95 ksi, an ultimate tensile stress of 165 ksi, and a very high fracture elongation of 34 - 36% when tested in a high-pressure (5 ksi) hydrogen environment. The tensile property data confirms hydrogen has little influence on HR-2’s ductility, strength, and fracture behavior. L-PBF NASA HR-2 is a promising option for many hydrogen-sensitive LRE components that require exceptional resistance to HEE. This paper will provide an overview of the L-PBF process development, material characterization, mechanical and thermophysical properties, and LRE hardware development for NASA HR-2.

Po S Chen↗

NASA NTRS2003

An Experimental Investigation of Hypergolic Ignition Delay of Hydrogen Peroxide with Fuel Mixtures

An experimental investigation of hypergolicity and ignition delay of fuel mixtures with hydrogen peroxide is presented. Example results of high speed photography and schleiren from drop tests are shown. Also, a discussion of the sensitivity to experimental parameters such as drop size and subsequent uncertainty considerations of ignition delay results is presented. It is shown that using the described setup on the mixtures presented, the precision uncertainty is on the order of 6% of average ignition delay and 5% of average decomposition delay. This represents sufficient repeatability for first order discrimination of ignition delay for propellant development and screening. Two mixtures, each using commonly available amines and transition metal compounds, are presented as examples that result in ignition delays on the order of 10 milliseconds.

Blevins, John A.↗

NASA NTRS2007

Mars Science Laboratory: Entry, Descent, and Landing System Performance

In 2010, the Mars Science Laboratory (MSL) mission will pioneer the next generation of robotic Entry, Descent, and Landing (EDL) systems, by delivering the largest and most capable rover to date to the surface of Mars. To do so, MSL will fly a guided lifting entry at a lift-to-drag ratio in excess of that ever flown at Mars, deploy the largest parachute ever at Mars, and perform a novel Sky Crane maneuver. Through improved altitude capability, increased latitude coverage, and more accurate payload delivery, MSL is allowing the science community to consider the exploration of previously inaccessible regions of the planet. The MSL EDL system is a new EDL architecture based on Viking heritage technologies and designed to meet the challenges of landing increasing massive payloads on Mars. In accordance with level-1 requirements, the MSL EDL system is being designed to land an 850 kg rover to altitudes as high as 1 km above the Mars Orbiter Laser Altimeter defined areoid within 10 km of the desired landing site. Accordingly, MSL will enter the largest entry mass, fly the largest 70 degree sphere-cone aeroshell, generate the largest hypersonic lift-to-drag ratio, and deploy the largest Disk-Gap-Band supersonic parachute of any previous mission to Mars. Major EDL events include a hypersonic guided entry, supersonic parachute deploy and inflation, subsonic heatshield jettison, terminal descent sensor acquisition, powered descent initiation, sky crane terminal descent, rover touchdown detection, and descent stage flyaway. Key performance metrics, derived from level-1 requirements and tracked by the EDL design team to indicate performance capability and timeline margins, include altitude and range at parachute deploy, time on radar, and propellant use. The MSL EDL system, which will continue to develop over the next three years, will enable a notable extension in the advancement of Mars surface science by delivering more science capability than ever before to the surface of Mars. This paper describes the current MSL EDL system performance as predicted by end-to-end EDL simulations, highlights the sensitivity of this baseline performance to several key environmental assumptions, and discusses some of the challenges faced in delivering such an unprecedented rover payload to the surface of Mars.

Way, David W.↗

NASA NTRS2006

Mars Science Laboratory: Entry, Descent, and Landing System Performance

Way, David W.↗

NASA NTRS1973

Study of dual-channel infrared spectroradiometer systems

A dual channel infrared spectroradiometer system was installed at a high altitude facility where spectral data were obtained from F2/H2, FLOX/CH4, FLOX/B2H6, OF2/CH4, and OF2/B2H6 propellant combinations. A good HF emission spectra was obtained for tests with the OF2B2H6 propellant combination. Subsequently, data were obtained on the LOX/GH2 propellant combination at sea level and altitude conditions that were inconsistent because the calculation of plume properties led to physically unrealistic results. The inconsistencies resulted from measurement accuracy not being within the narrow limits required by the data reduction scheme. Although the anticipated program results were not obtained, the data gathered are useful for guiding future experimental efforts and formulations of less sensitive analytical models.

Wrubel, J. A.↗

NASA NTRS1990

Radiation effect on rocket engine performance

Critical problem areas involving the effect of radiation on the combustion of bipropellants are addressed by formulating a universal scaling law in combination with a radiation-enhanced vaporization combustion model. Numerical algorithms are developed and data pertaining to the Variable Thrust Engine (VTE) and the Space Shuttle Main Engine (SSME) are used to conduct parametric sensitivity studies to predict the principal intercoupling effects of radiation. The analysis reveals that low-enthalpy engines, such as the VTE, are vulnerable to a substantial performance setback due to radiative loss, whereas the performance of high-enthalpy engines such as the SSME are hardly affected over a broad range of engine operation. Combustion enhancement by radiative heating of the propellant has a significant impact on propellants with high absorptivity.

Chiu, Huei-Huang↗

NASA NTRS2000

Potential Subjective Effectiveness of Active Interior Noise Control in Propeller Airplanes

Active noise control technology offers the potential for weight-efficient aircraft interior noise reduction, particularly for propeller aircraft. However, there is little information on how passengers respond to this type of interior noise control. This paper presents results of two experiments that use sound quality engineering practices to determine the subjective effectiveness of hypothetical active noise control (ANC) systems in a range of propeller aircraft. The two experiments differed by the type of judgments made by the subjects: pair comparisons based on preference in the first and numerical category scaling of noisiness in the second. Although the results of the two experiments were in general agreement that the hypothetical active control measures improved the interior noise environments, the pair comparison method appears to be more sensitive to subtle changes in the characteristics of the sounds which are related to passenger preference. The reductions in subjective response due to the ANC conditions were predicted with reasonable accuracy by reductions in measured loudness level. Inclusion of corrections for the sound quality characteristics of tonality and fluctuation strength in multiple regression models improved the prediction of the ANC effects.

Powell, Clemans A.↗

NASA NTRS1986

Mission and vehicle sizing sensitivities

Representative interplanetary space vehicle systems are sized to compare and show sensitivity of the initial mass required in low Earth orbit to one mission mode and mission opportunity. Data are presented to show the requirements for Earth-Mars opposition and conjunction class roundtrip flyby and stopover mission opportunities available during the time period from year 1997 to year 2045. The interplanetary space vehicle consists of a spacecraft and a space vehicle acceleration system. Propellant boil-off for the various mission phases is given for the Lox/LH (Liquid Oxygen/Liquid Hydrogen) propulsion systems. Mission abort information is presented for the 1999 Venus outbound swingby trajectory, transfer profile.

Archie C Young↗

NASA NTRS1988

Conceptual design of a lunar oxygen pilot plant Lunar Base Systems Study (LBSS) task 4.2

The primary objective was to develop conceptual designs of two pilot plants to produce oxygen from lunar materials. A lunar pilot plant will be used to generate engineering data necessary to support an optimum design of a larger scale production plant. Lunar oxygen would be of primary value as spacecraft propellant oxidizer. In addition, lunar oxygen would be useful for servicing nonregenerative fuel cell power systems, providing requirements for life support, and to make up oxygen losses from leakage and airlock cycling. Thirteen different lunar oxygen production methods are described. Hydrogen reduction of ilmenite and extraction of solar-wind hydrogen from bulk lunar soil were selected for conceptual design studies. Trades and sensitivity analyses were performed with these models.

Source record↗

NASA NTRS1988

Radiation effect on rocket engine performance

The effects of radiation on the performance of modern rocket propulsion systems operating at high pressure and temperature were recognized as a key issue in the design and operation of various liquid rocket engines of the current and future generations. Critical problem areas of radiation coupled with combustion of bipropellants are assessed and accounted for in the formulation of a universal scaling law incorporated with a radiation-enhanced vaporization combustion model. Numerical algorithms are developed and the pertaining data of the Variable Thrust Engine (VTE) and Space Shuttle Main Engine (SSME) are used to conduct parametric sensitivity studies to predict the principal intercoupling effects of radiation. The analysis reveals that low enthalpy engines, such as the VTE, are vulnerable to a substantial performance set back by the radiative loss, whereas the performance of high enthalpy engines such as the SSME, are hardly affected over a broad range of engine operation. Additionally, combustion enhancement by the radiative heating of the propellant has a significant impact in those propellants with high absorptivity. Finally, the areas of research related with radiation phenomena in bipropellant engines are identified.

Chiu, Huei-Huang↗

NASA NTRS1972

Evaluation of Skylab IB sensitivity to on-pad winds with turbulence

Computer simulation was performed to estimate displacements and bending moments experienced by the SKYLAB 1B vehicle on the launch pad due to atmospheric winds. The vehicle was assumed to be a beam-like structure represented by a finite number of generalized coordinates. Wind flow across the vehicle was treated as a nonhomogeneous, stationary random process. Response computations were performed by the assumption of simple strip theory and application of generalized harmonic analysis. Displacement and bending moment statistics were obtained for six vehicle propellant loading conditions and four representative reference wind profile and turbulence levels. Means, variances and probability distributions are presented graphically for each case. A separate analysis was performed to indicate the influence of wind gradient variations on vehicle response statistics.

Coffin, T.↗

NASA NTRS2004

Aerocapture Guidance Performance for the Neptune Orbiter

A performance evaluation of the Hybrid Predictor corrector Aerocapture Scheme (HYPAS) guidance algorithm for aerocapture at Neptune is presented in this paper for a Mission to Neptune and the Neptune moon Triton'. This mission has several challenges not experienced in previous aerocapture guidance assessments. These challengers are a very high Neptune arrival speed, atmospheric exit into a high energy orbit about Neptune, and a very high ballistic coefficient that results in a low altitude acceleration capability when combined with the aeroshell LD. The evaluation includes a definition of the entry corridor, a comparison to the theoretical optimum performance, and guidance responses to variations in atmospheric density, aerodynamic coefficients and flight path angle for various vehicle configurations (ballistic numbers). The benefits of utilizing angle-of-attack modulation in addition to bank angle modulation to improve flight performance is also discussed. The results show that despite large sensitivities in apoapsis targeting, the algorithm performs within the allocated AV budget for the Neptune mission bank angle only modulation. The addition of angle-of-attack modulation with as little as 5 degrees of amplitude significantly improves the scatter in final orbit apoapsis. Although the angle-of-attack modulation complicates the vehicle design, the performance enhancement reduces aerocapture risk and reduces the propellant consumption needed to reach the high energy target orbit for a conventional propulsion system.

Masciarelli, James P.↗

NASA NTRS2019

Multipoint Design and Optimization of a Turboshaft Engine for a Tiltwing Turboelectric Vertical Takeoff Landing Air Taxi

This paper describes the design of a turboshaft engine for a tiltwing air taxi application. In this case, the tiltwing air taxi is intended to fly a 400 nm mission with up to fifteen passengers. Engine requirements for the concept engine are taken from aircraft system studies where thrust is produced by four propellers driven by electric motors and powered by a single gas turbine engine. The purpose of this paper is to perform a cycle design optimization that minimizes fuel consumption and weight while respecting current technology limitations to meet mission requirements. To achieve results, the engine overall pressure ratio and maximum temperature at the exit of the combustor are set as the design parameters. Several sensitivity studies are also performed to visualize optimization trends. Results of the optimization study show solutions are heavily dependent on engine cooling flow requirements and exact mission requirements. This engine is intended for use in large system optimization research.

gas turbine↗

NASA NTRS2019

Multi-Point Design and Optimization of a Turboshaft Engine for a Tiltwing Turboelectric VTOL Air Taxi

This paper describes the design of a turboshaft engine for a tiltwing air taxi application. In this case, the tiltwing air taxi is intended to fly a 400-nautical mile mission with up to fifteen passengers. Engine requirements for the concept engine are taken from aircraft system studies where thrust is produced by four propellers driven by electric motors and powered by a single gas turbine engine. The purpose of this paper is to perform a cycle design optimization that minimizes fuel consumption and weight while respecting current technology limitations to meet mission requirements. To achieve results, the engine overall pressure ratio and maximum temperature at the exit of the combustor are set as the design parameters. Several sensitivity studies are also performed to visualize optimization trends. Results of the optimization study show solutions are heavily dependent on engine cooling flow requirements and exact mission requirements. This engine is intended for use in large system optimization research.

Optimization↗

NASA NTRS2015

Advanced Booster Composite Case/Polybenzimidazole Nitrile Butadiene Rubber Insulation Development

The NASA Engineering and Safety Center (NESC) was requested to examine processing sensitivities (e.g., cure temperature control/variance, debonds, density variations) of polybenzimidazole nitrile butadiene rubber (PBI-NBR) insulation, case fiber, and resin systems and to evaluate nondestructive evaluation (NDE) and damage tolerance methods/models required to support human-rated composite motor cases. The proposed use of composite motor cases in Blocks IA and II was expected to increase performance capability through optimizing operating pressure and increasing propellant mass fraction. This assessment was to support the evaluation of risk reduction for large booster component development/fabrication, NDE of low mass-to-strength ratio material structures, and solid booster propellant formulation as requested in the Space Launch System NASA Research Announcement for Advanced Booster Engineering Demonstration and/or Risk Reduction. Composite case materials and high-energy propellants represent an enabling capability in the Agency's ability to provide affordable, high-performing advanced booster concepts. The NESC team was requested to provide an assessment of co- and multiple-cure processing of composite case and PBI-NBR insulation materials and evaluation of high-energy propellant formulations.

Gentz, Steve↗

NASA NTRS1976

Heat transfer evaluation in a plasma core reactor

Numerical evaluations of heat transfer in a fissioning uranium plasma core reactor cavity, operating with seeded hydrogen propellant, was performed. A two-dimensional analysis is based on an assumed flow pattern and cavity wall heat exchange rate. Various iterative schemes were required by the nature of the radiative field and by the solid seed vaporization. Approximate formulations of the radiative heat flux are generally used, due to the complexity of the solution of a rigorously formulated problem. The present work analyzes the sensitivity of the results with respect to approximations of the radiative field, geometry, seed vaporization coefficients and flow pattern. The results present temperature, heat flux, density and optical depth distributions in the reactor cavity, acceptable simplifying assumptions, and iterative schemes. The present calculations, performed in cartesian and spherical coordinates, are applicable to any most general heat transfer problem.

Smith, D. E.↗

NASA NTRS1998

RHETT/EPDM Performance Characterization

The 0.6 kW Electric Propulsion Demonstration Module (EPDM) flight thruster system was tested in a large vacuum facility for performance measurements and functional checkout. The thruster was operated at a xenon flow rate of 3.01 mg/s, which was supplied through a self-contained propellant system. All power was provided through a flight-packaged power processing unit, which was mounted in vacuum on a cold plate. The thruster was cycled through 34 individual startup and shutdown sequences. Operating periods ranged from 3 to 3600 seconds. The system responded promptly to each command sequence and there were no involuntary shutdowns. Direct thrust measurements indicated that steady state thrust was temperature sensitive, and varied from a high of 41.7 mN at 16 C, to a low of 34.8 mN at 110 C. Short duration thruster firings showed rapid response and good repeatability.

Haag, T.↗