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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Cooling Duct Analysis for Transpiration/Film Cooled Liquid Propellant Rocket Engines

The development of a low cost space transportation system requires that the propulsion system be reusable, have long life, with good performance and use low cost propellants. Improved performance can be achieved by operating the engine at higher pressure and temperature levels than previous designs. Increasing the chamber pressure and temperature, however, will increase wall heating rates. This necessitates the need for active cooling methods such as film cooling or transpiration cooling. But active cooling can reduce the net thrust of the engine and add considerably to the design complexity. Recently, a metal drawing process has been patented where it is possible to fabricate plates with very small holes with high uniformity with a closely specified porosity. Such a metal plate could be used for an inexpensive transpiration/film cooled liner to meet the demands of advanced reusable rocket engines, if coolant mass flow rates could be controlled to satisfy wall cooling requirements and performance. The present study investigates the possibility of controlling the coolant mass flow rate through the porous material by simple non-active fluid dynamic means. The coolant will be supplied to the porous material by series of constant geometry slots machined on the exterior of the engine.

Micklow, Gerald J.↗
NASA NTRS2005

Low-Cost Flow Visualization for a Supersonic Ejector

Shadowgraph techniques were applied to the cold flow ejector facility at the Propulsion Research Center at the University of Alabama in Huntsville. The setup for the experiments was relatively simple and was accomplished at very little cost. Series of shadowgraph images were taken of both dual nozzle and single nozzle strut based ejectors operating over a range of chamber pressures. The density gradient patterns in the shadowgraphs were compared to pressure data measured along the top and side walls of the mixing duct. The shadowgraph images showed the presence of barrel shocks emanating from the nozzles which at low pressures terminated in Mach disks and at higher pressures extended beyond the barrel shape and reflected off the walls of the duct. Based on pressure data from previous testing, reflected shocks were expected on the walls of the duct. The shadowgraph images confirmed the locations of these reflected shocks on the top wall of the duct. The shadowgraph images also showed the structure change which correlated to a change in pitch of the ejector noise, and corresponded to a change in trend of the duct wall pressure ratio distributions. The images produced from the setup provided insight into the complex flow behavior inside the ejector duct. In addition, the techniques were a valuable tool as an educational device for students.

Olden, George W.↗
NASA NTRS1972

Propulsion sizing program

Computer program was written to evaluate and define optimum design parameters of low pressure auxiliary propulsion system. APS will provide attitude and translational control of space shuttle vehicles. Program evaluates subsystem design parameters such as engine chamber pressure, mixture ratio, expansion ratio, and subsystem component weight and size.

Kaemming, T. A.↗
NASA NTRS1972

Analysis of a supersonic-combustion rocket concept

A preliminary analysis has been made of a supersonic-combustion rocket engine concept using hydrogen and oxygen propellants. The ejector action of a separate small rocket motor is employed to pump the propellants to high stagnation pressures and supersonic velocities. Therefore complicated heavy turbopumps are eliminated and cooling problems of a sonic throat are reduced. The results of the study show that vacuum specific impulse levels as high as a conventional rocket having the same chamber pressure as the drive motor are possible. The supersonic-combustion rocket would be an attractive alternate for a high-altitude low-thrust conventional rocket operating with a pressure feed propellant system. It would also be a convenient technique for obtaining extremely high thrusts without the need for developing corresponding large turbopumps.

Franciscus, L. C.↗
NASA NTRS1960

Effect of Propellant and Catalyst Bed Temperatures on Thrust Buildup in Several Hydrogen Peroxide Reaction Control Rockets

An investigation was undertaken to determine the effect of chamber and propellant feed temperatures on the starting characteristics of hydrogen peroxide thrust chambers. Start delay times for two types of thrust chamber designs in the 1- to 24-pound-thrust range were obtained over a range of chamber and propellant feed temperatures from 30 to 100 F. Start delay times obtained during the first minute of catalyst bed life and again after 6 minutes of total accumulated running time are presented as a function of chamber and propellant feed temperatures. The initial cold-start delay time of the hydrogen peroxide thrust chambers investigated was approximately 0.150 second to attain 90 percent of steady-state chamber pressure at chamber and propellant feed temperatures of 70 F and above. Both thrust chamber designs could be started at chamber and propellant feed temperatures as low as 30 F; start delay times did, however, generally increase at low temperatures. When the chamber was at an elevated temperature from a preceding firing, the start delay time was reduced to approximately 0.050 second, indicating a marked effect of chamber temperature at constant propellant feed temperatures. Accumulated run time affected the starting characteristics only when both the chamber and propellant feed temperatures were at reduced levels.

Wanhainen, John P.↗
NASA NTRS2010

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

Betts, Erin M.↗
NASA NTRS2022

Test Facilities for SHERLOC Laser Development

The Scanning Habitable Environments with Raman and Luminescence for Organics and Chemicals (SHERLOC) instrument is a deep UV laser based spectrometer that is part of NASA’s Mars Perseverance rover. The laser is a pulsed 248.6 nm NeCu hollow cathode gas discharge laser. The design, development, and testing of lasers and laser power supplies (LPS) were performed by scientists and engineers at the Jet Propulsion Laboratory (JPL) and Photon Systems Inc. (PSI). While these lasers had been used previously in extreme terrestrial environments, before they had to be qualified for operation and functionality over the expected range of environmental situations (temperature cycling, vibration, mechanical shock, low pressure corona emission testing) over the course of mission life time. The SHERLOC laser/LPS testing facilities consisted of custom-tailored environmental test chambers with metrology/control electronics. A custom LabVIEW software package was developed to autonomously operate all test facilities using a multi-threaded, object-oriented programming architecture, tasked with interfacing with many instruments simultaneously for operation and data acquisition.

Houck, Andrew↗
NASA NTRS1993

Test results from a simple, low-cost, pressure-fed liquid hydrogen/liquid oxygen rocket combustor

A simple, low-cost rocket engine was designed, fabricated, and successfully hot fire tested over a wide range of interface conditions and operating parameters. The engine used low enthalpy hydrogen (45 to 70 R, 200 to 390 psia) and oxygen (139 to 163 R, 210 to 480 psia) propellants pressure-fed directly from facility cryogenic tanks. The engine demonstrated excellent performance, with 97% average combustion efficiency, and absence of combustion instabilities. Engine design chamber pressure was 300 psia, yielding about 16,500 pounds thrust at sea level with a 3:1 expansion ration test nozzle. The engine used a fixed-element injector based on TRW's unique coaxial pintle design, but was operated at 60%, 80%, and 100% thrust levels by throttling facility propellant valves. The engine was tested at propellant mixture ratios (O/F) from 5.8 to 8.4; design O/F was 6.6. To document combustion stability, in five tests RDX explosive pulse guns were detonated in radial and tangential directions across the combustion chamber during steady-state operation. The largest disturbance consisted of simultaneous detonation of a 20-grain radial gun and a 40-grain tangential gun. In no case was an instability, either feed system mode or chamber acoustic mode, excited. High-frequency piezoelectric pressure transducers documented stable recovery from disturbance overpressures within 40 milliseconds of peak pressure. A total of 67 firing tests, accumulating 149 seconds of firing time above 10% P(sub c), were performed. Since parametric testing required run durations of only 2 to 3 seconds, a heat sink combustion chamber was employed for most runs. To evaluate the feasibility of a low-cost ablative system for a flight engine design, one 20-second continuous firing was conducted with a silicone rubber chamber/throat/nozzle liner cast in one piece directly into the engine. The ablative engine operated at the equivalent of 309 seconds sea level specific impulse, when adjusted to a 98% efficient 6:1 expansion ration nozzle, and 430 seconds vacuum specific impulse, when adjusted to a 98% efficient 50:1 expansion ratio nozzle. This engine and test series represent an initial subscale demonstration of a new booster-class engine that eliminates the cost and complexity associated with regenerative cooling and typical engine cycles. This paper presents a description of the engine design and discussion and summary data plots of the performance measured during the parametric testing.

Dressler, G. A.↗
NASA NTRS2015

Zero Boil-Off System Testing

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration due to their high specific impulse for rocket motors of upper stages suitable for transporting 10s to 100s of metric tons of payload mass to destinations outside of low earth orbit and for their return. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for missions with durations greater than several months. These losses can be eliminated by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and modulating the cryocooler to control tank pressure. The active thermal control technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and its integration to the propellant tank through a distributed cooling tubing network coupled to the tank wall. This configuration was recently tested at NASA Glenn Research Center, in a vacuum chamber and cryoshroud that simulated the essential thermal aspects of low Earth orbit, its vacuum and temperature. Testing consisted of three passive tests with the active cryocooler system off, and 7 active tests, with the cryocooler powered up. The test matrix included zero boil-off tests performed at 90 full and 25 full, and several demonstrations at excess cooling capacity and reduced cooling capacity. From this, the tank pressure response with varied cryocooler power inputs was determined. This test series established that the active cooling system integrated with the propellant tank eliminated boil-off and robustly controlled tank pressure.

Plachta, David W.↗
NASA NTRS1987

Carbon deposition characteristics of LO2/HC propellants

The generation and deposition of carbon have been studied using subscale hardware with LO2/RP-1, LO2/propane, and LO2/methane at low mixture ratio conditions. The deposition of carbon on the turbine simulator tubes was evaluated at mixture ratios of 0.20 to 0.60, and at chamber pressures from 720 to 1650 psia. The carbon-deposition rate is a strong function of mixture ratio and a weak function of chamber pressure. There is a mixture ratio that will minimize deposition for LO2/RP-1; a threshold mixture ratio for LO2/propane; and no deposition for LO2/methane at any mixture ratio tested. The turbine drive operating limits were defined for each fuel tested.

Hernandez, Rosemary↗
NASA NTRS1992

Pressure fed thrust chamber technology program

This is the final report for the Pressure Fed Technology Program. It details the design, fabrication and testing of subscale hardware which successfully characterized LOX/RP combustion for a low cost pressure fed design. The innovative modular injector design is described in detail as well as hot-fire test results which showed excellent performance. The program summary identifies critical LOX/RP design issues that have been resolved by this testing, and details the low risk development requirements for a low cost engine for future Expendable Launch Vehicles (ELVi).

Dunn, Glenn M.↗
NASA NTRS1923

Performance of B. M. W. 185-Horsepower Airplane Engine

This report deals with the results of a test made upon a B. M. W. Engine in the altitude chamber of the Bureau of Standards, where controlled conditions of temperature and pressure can be made to simulate those of the desired altitude. A remarkably low value of fuel consumption - 041 per B. H. P. hour - is obtained at 1,200 revolutions per minute at an air density of 0.064 pound per cubic foot and a brake thermal efficiency of 33 per cent and an indicated efficiency of 37 per cent at the above speed and density. In spite of the fact that the carburetor adjustment does not permit the air-fuel ratio of maximum economy to be obtained at air densities lower than 0.064, the economy is superior to most engines tested thus far, even at a density lower than 0.064, the economies superior to most engines tested thus far, even at a density (0.03) corresponding to an altitude of 25,000 feet. The brake mean effective pressure even at full throttle is rather low. Since the weight of much of the engine is governed more by its piston displacement than by the power developed, a decreased mean effective pressure usually necessitates increased weight per horsepower. The altitude performance of the engine is, in general, excellent, and its low fuel consumption is the outstanding feature of merit.

Sparrow, S W↗
NASA NTRS1994

Xenon lighting adjusted to plant requirements

Xenon lamps are available as low and high power lamps with relatively high efficiency and a relatively long lifetime up to several thousand hours. Different construction types of short-arc and long-arc lamps permit a good adaptation to various applications in projection and illumination techniques without substantial changes of the spectral quality. Hence, the xenon lamp was the best choice for professional technical purposes where high power at simultaneously good spectral quality of the light was required. However, technical development does not stand still. Between the luminous efficacy of xenon lamps of 25-50 lm/W and the theoretical limit for 'white light' of 250 lm/W is still much room for improvement. The present development mainly favors other lamp types, like metal halide lamps and fluorescent lamps for commercial lighting purposes. The enclosed sections deal with some of the properties of xenon lamps relevant to plant illumination; particularly the spectral aspects, the temporal characteristics of the emission, and finally the economy of xenon lamps will be addressed. Due to radiation exceeding the natural global radiation in both the ultraviolet (UV) and the infrared (IR) regions, filter techniques have to be included into the discussion referring to the requirements of plant illumination. Most of the presented results were obtained by investigations in the GSF phytotron or in the closed Phytocell chambers of the University of Erlangen. As our experiences are restricted to area plant illumination rather than spot lights our discussion will concentrate on low pressure long-arc xenon lamps which are commonly used for such plant illuminations. As the spectral properties of short-arc lamps do not differ much from those of long-arc lamps most of our conclusions will be valid for high pressure xenon lamps too. These lamps often serve as light sources for small sun simulators and for monochromators which are used for action spectroscopy of plant responses.

Koefferlein, M.↗
NASA NTRS2017

Liquid Nitrogen Zero Boiloff Testing

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASAs future space exploration due to their high specific impulse for rocket motors of upper stages suitable for transporting 10s to 100s of metric tons of payload mass to destinations outside of low earth orbit and for their return. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for missions with durations greater than several months. These losses can be eliminated by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and modulating the cryocooler to control tank pressure. The active thermal control technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and its integration to the propellant tank through a distributed cooling tubing network coupled to the tank wall. This configuration was recently tested at NASA Glenn Research Center, in a vacuum chamber and cryo-shroud that simulated the essential thermal aspects of low Earth orbit, its vacuum and temperature. Testing consisted of three passive tests with the active cryo-cooler system off, and 7 active tests, with the cryocooler powered up. The test matrix included zero boil-off tests performed at 90 full and 25 full, and several demonstrations at excess cooling capacity and reduced cooling capacity. From this, the tank pressure response with varied cryocooler power inputs was determined. This test series established that the active cooling system integrated with the propellant tank eliminated boil-off and robustly controlled tank pressure.

Cryogenic fluid management↗
NASA NTRS1980

Evaluation of flow quality in two large NASA wind tunnels at transonic speeds

Wind tunnel testing of low drag airfoils and basic transition studies at transonic speeds are designed to provide high quality aerodynamic data at high Reynolds numbers. This requires that the flow quality in facilities used for such research be excellent. To obtain a better understanding of the characteristics of facility disturbances and identification of their sources for possible facility modification, detailed flow quality measurements were made in two prospective NASA wind tunnels. Experimental results are presented of an extensive and systematic flow quality study of the settling chamber, test section, and diffuser in the Langley 8 foot transonic pressure tunnel and the Ames 12 foot pressure wind tunnel. Results indicate that the free stream velocity and pressure fluctuation levels in both facilities are low at subsonic speeds and are so high as to make it difficult to conduct meaningful boundary layer control and transition studies at transonic speeds.

Harvey, W. D.↗
NASA NTRS1968

High-voltage pulse generator developed for wide-gap spark chambers

Low-inductance, high-capacitance Marx pulse generator provides for minimization of internal inductance and suppression of external electromagnetic radiation. The spark gaps of the generator are enclosed in a pressurized nitrogen atmosphere which allows the charging voltage to be varied by changing the nitrogen pressure.

Keller, L. P.↗