SEARCH · Engineering Papers
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
A sound data base was established by analytically and experimentally generating basic regenerative cooling, combustion performance, combustion stability, and combustion chamber heat transfer parameters for LOX/HC propellants, with specific application to second generation orbit maneuvering and reaction control systems (OMS/RCS) for the Space Shuttle Orbiter.
In 1969 a combustion-heated direct-connect test facility was assembled at the NASA Langley Research Center and used to support NASA’s Hypersonic Research Engine program. Later, the facility was used to support NASA’s generic hypersonic program, assessing the mixing, ignition, flame holding, and combustion characteristics of ramjet and scramjet combustors. Over the years, the facility has been updated and modified to conduct increasingly complicated tests. Today, the facility is known as the Direct Connect Supersonic Combustion Test Facility (DCSCTF) and is part of the NASA Langley Scramjet Test Complex. Recently, the facility has been used by external customers to test fuel-cooled and uncooled, flight-like test articles. Tests have been conducted at nominal Mach 5 flight conditions for up to 15 minutes of continuous operation. The facility remains a viable and useful tool in the support of various Department of Defense projects.
A program is underway to demonstrate the practical feasibility of thermally-choked combustor technology with particular emphasis on rocket propulsion applications. Rather than induce subsonic to supersonic flow transition in a geometric throat, the goal is to create a thermal throat by adding combustion heat in a diverging nozzle. Such a device would have certain advantages over conventional flow accelerators assuming that the pressure loss due to heat addition does not severely curtail propulsive efficiency. As an aid to evaluation, a generalized one-dimensional compressible flow analysis tool was constructed. Simplified calculations indicate that the process is fluid dynamically and thermodynamically feasible. Experimental work is also being carried out in an attempt to develop, assuming an array of practical issues are surmountable, a practical bench-scale demonstrator using high flame speed H2/O2 combustibles.
The use of throat film coolant as a method to reduce peak combustion heat flux was investigated. Hot-fire tests were performed using a subscale LOX/RP-1 combustor at a nominal chamber pressure of 2000 psia. A special injection manifold was fabricated to introduce the film coolant into the chamber. Four tests were conducted with RP-1 film coolant injected 4.4 inches upstream of the throat. In order to determine the effect of chamber length on the film coolant, four additional tests were performed with the injection point moved 16.4 inches upstream of the throat. Results from the test program are discussed. Methods for incorporating throat film cooling into a flightweight combustion chamber are also discussed.
The heavy-duty transportation sector has primarily relied on conventional diesel combustion engines given their reliability and high thermal efficiency relative to spark ignition engines, but increased focus on reducing greenhouse gas emissions has led to investigation into alternative fuels. Gaseous hydrogen fuel has garnered a great deal of recent interest in the engine community given it has zero carbon, but hydrogen is not available at the scale and cost that petroleum fuels are currently available, and this is a barrier to adoption for industries that are looking to decarbonize their operations. Because of the fuel flexibility provided, dual fuel technology offers a pathway for some industries to adopt hydrogen as a fuel source while maintaining sufficient flexibility in times and locations where the new fuel is not yet available. This computational study investigates dual fuel combustion in a large bore locomotive engine architecture using direct injected diesel and port injected gaseous hydrogen fuel. With an optimal port fuel injection configuration from previous work, simulations of varying substitution ratio, compression ratio, manifold air temperature, diesel injection timing, and diesel injection pressure were performed to understand their effect on combustion performance. Results indicated that both increased substitution ratio and higher intake air temperature accelerates hydrogen flame propagation and can result in high peak cylinder pressures. Additionally, diesel injection timing and injection pressure were demonstrated as effective methods for controlling dual fuel combustion heat release rates.
Venus and mars atmosphere entry simulated by combustion heated hypervelocity wind tunnel
Selected aerospace technology conference for petroleum industry representatives - combustion, heat transfer, surface physics and chemistry, lubricants and bearings, hydrodynamics
Tabulated data on heats of combustion and heats of formation of organic compounds of biological interest
A thermal energy storage (TES) system developed for use with the Stirling engine as an automotive power system has gravimetric and volumetric storage densities which are competitive with electric battery storage systems, meets all operational requirements for a practical vehicle, and can be packaged in compact sized automobiles with minimum impact on passenger and freight volume. The TES/Stirling system is the only storage approach for direct use of combustion heat from fuel sources not suitable for direct transport and use on the vehicle. The particular concept described is also useful for a dual mode TES/liquid fuel system in which the TES (recharged from an external energy source) is used for short duration trips (approximately 10 miles or less) and liquid fuel carried on board the vehicle used for long duration trips. The dual mode approach offers the potential of 50 percent savings in the consumption of premium liquid fuels for automotive propulsion in the United States.
The liquid rocket propulsion technology needs to support anticipated future space vehicles were examined including any special action needs to be taken to assure that an industrial base in substained. Propulsion system requirements of Earth-to-orbit vehicles, orbital transfer vehicles, and planetary missions were evaluated. Areas of the fundamental technology program undertaking these needs discussed include: pumps and pump drives; combustion heat transfer; nozzle aerodynamics; low gravity cryogenic fluid management; and component and system life reliability, and maintenance. The primary conclusion is that continued development of the shuttle main engine system to achieve design performance and life should be the highest priority in the rocket engine program.
A planned modification of the NASA Langley 8-Foot High Temperature Tunnel to make it a unique national research facility for hypersonic air-breathing propulsion systems is described, and some of the ongoing supporting research for that modification is discussed. The modification involves: (1) the addition of an oxygen-enrichment system which will allow the methane-air combustion-heated test stream to simulate air for propulsion testing; and (2) supplemental nozzles to expand the test simulation capability from the current nominal Mach number to 7.0 include Mach numbers 3.0, 4.5, and 5.0. Detailed design of the modifications is currently underway and the modified facility is scheduled to be available for tests of large scale propulsion systems by mid 1988.
A planned modification of the NASA Langley 8-Foot High Temperature Tunnel to make it a unique national research facility for hypersonic air-breathing propulsion systems is described, and some of the ongoing supporting research for that modification is discussed. The modification involves: (1) the addition of an oxygen-enrichment system which will allow the methane-air combustion-heated test stream to simulate air for propulsion testing; and (2) supplemental nozzles to expand the test simulation capability from the current nominal Mach number to 7.0 include Mach numbers 3.0, 4.5, and 5.0. Detailed design of the modifications is currently underway and the modified facility is scheduled to be available for tests of large scale propulsion systems by mid 1988.
A new method is described for deducing the thermal-acoustic efficiency spectrum (defined as the fraction of combustion heat release converted to acoustic energy at a given frequency) of a long turbulent burner from the sound spectrum measured in the far field. The method, which is based on a one-dimensional model of the unsteady flow in the burner, is applied to a tubular diffusion-flame hydrogen burner whose length is large compared to its diameter. The results for thermal powers ranging from 4.5 to 22.3 kW show that the thermal-acoustic efficiency is relatively insensitive to the burner power level, decreasing from a value of around 0.0001 at 150 Hz with a slope of about 20 dB per decade. Evidence is presented indicating that acoustic agitation of the flame below 500 Hz, especially in the neighborhood of the resonant frequencies of the burner, is a significant acoustic source.
Five black, electrically-conductive thermal control coatings have been formulated and tested for application on the Galileo spacecraft. The coatings consist of both organic and inorganic systems applied on titanium, aluminum, and glass/epoxy composite surfaces. The coatings were tested under simulated space environment conditions. Coated specimens were subjected to thermal radiation, convective and combustive heating, and cryogenic conditions over a temperature range between -196 C and 538 C. Mechanical, physical, thermal, electrical, and thermooptical properties are presented for one of these coatings. This paper describes the preparation, characteristics, and spraying of iron titanate on titanium and aluminum, and presents performance results.
Consideration is given to the development of testing methods for the National Aerospace Plane (NASP) program, focusing on techniques for simulating and testing scramjet engines and scramjet engine components during flight from Mach 4 to Mach 20. The flight conditions associated with the NASP are outlined and the stagnation temperatures and pressures required to simulate flight up to orbital speeds are presented. The types of test facilities needed for scramjet propulsion testing are discussed, including combustion-heated facilities, electric arc-heated facilities, convection-heated facilities, shock tunnels, and expansion tubes. Issues related to assessing the flow quality in the scramjet ground testing facilities are examined, including uniformity, contamination, dissociation, ionization, nonequilibrium, and turbulence. Direct-connect combustor versus free-jet scramjet engine tests are compared and the capabilities of current and near-future scramjet engine test facilities are evaluated.
A coherent anti-Stokes Raman scattering (CARS) system has been hardened for utilization in a NASA Langley supersonic combustion test cell that can obtain temperature cross sections of the flow at three locations. This system is remotely operated and environmentally protected. Measurements were obtained in a scramjet combustor model consisting of a rearward-facing step, followed by an expansion duct. The freestream conditions were Mach 2, with static pressure that ranged from 0.8 to 1.9 atm, and a static temperature of about 800 K. Data for two different flow conditions were obtained that provided a comparison between reacting and nonreacting mixing of injected hydrogen fuel with the combustion-heated supersonic stream.
Experimental and computational fluid dynamic activities in rocket propulsion were discussed. The workshop was an open meeting of government, industry, and academia. A broad number of topics were discussed including computational fluid dynamic methodology, liquid and solid rocket propulsion, turbomachinery, combustion, heat transfer, and grid generation.
Coherent anti-Stokes Raman spectroscopy (CARS) thermometry has been used to obtain static temperature cross sections in a three-dimensional supersonic combustor flowfield. Data were obtained in three spanwise planes downstream of a single normal fuel injector which was located downstream of a rearward-facing step. The freestream flow was nominally Mach 2 and was combustion heated to a total temperature of 1440 K (yielding a static temperature of about 800 K in the freestream) to simulate the inflow to a combustor operating at a flight Mach number of about 5.4. Since a broadband probe laser was used an instantaneous temperature sample was obtained with each laser shot at a repetition rate of 10 Hz. Thus root-mean-square (rms) temperatures and temperature probability density functions (pdf's) were obtained in addition to mean temperatures.