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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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Coaxial injector response to axial instability under simulated cold-flow

The frequency dependence of the admittances and response factors of various gaseous rocket injector configurations subject to axial instabilities under cold-flow conditions, have been measured using the modified impedance-tube technique. The tested configurations simulate the flows in a gaseous-fuel injector, gaseous-oxidizer injector and a coaxial injector with both fuel and oxidizer elements. Comparison of the measured response data with corresponding data predicted by the Feiler and Heidmann model indicates good agreement between the two sets of data.

Janardan, B. A.↗
NASA NTRS1978

Ion chromatographic determination of sulfur in fuels

The sulfur content of fuels was determined using an ion chromatograph to measure the sulfate produced by a modified Parr bomb oxidation. Standard Reference Materials from the National Bureau of Standards, of approximately 0.2 + or - 0.004% sulfur, were analyzed resulting in a standard deviation no greater than 0.008. The ion chromatographic method can be applied to conventional fuels as well as shale-oil derived fuels. Other acid forming elements, such as fluorine, chlorine and nitrogen could be determined at the same time, provided that these elements have reached a suitable ionic state during the oxidation of the fuel.

Mizisin, C. S.↗
NASA NTRS1977

Quiet Clean Short-haul Experimental Engine (QCSEE) over-the-wing control system design report

A control system incorporating a digital electronic control was designed for the over-the-wing engine. The digital electronic control serves as the primary controlling element for engine fuel flow and core compressor stator position. It also includes data monitoring capability, a unique failure indication and corrective action feature, and optional provisions for operating with a new type of servovalve designed to operate in response to a digital-type signal and to fail with its output device hydraulically locked into position.

Source record↗
NASA NTRS1981

NASA V/STOL Propulsion Control Analysis - Phase I and II program status

NASA-Lewis Research Center initiated the V/STOL Propulsion Control Analysis Program in 1979 in order to take advantage of advanced electronic control technology for the next generation of V/STOL aircraft. The rationale, methods, and criteria developed during Phase I and Phase II of the program are discussed first. The development of an integrated flight and propulsion control system is then described. Most V/STOL aircraft under consideration depend on reaction thrusters or separate lift engines for attitude control and on engine thrust variations for height control. S/CTOL aircraft vector-thrust for pitch control during low speed operation. The baseline propulsion control system provides a thrust level response of 10 rad/sec and a vectoring response of 22 rad/sec. The requirements of the control components include: (1) redundant electronic control computers with minimum software and 100% Fault Coverage; (2) prime reliable electromechanical actuator interfaces; (3) miniature, digitally compatible sensors with easily detected failure modes; and (4) dual element, fail operational fuel pumps.

Miller, R. J.↗
NASA NTRS1987

Combustion chamber heat transfer characterization of LOX/hydrocarbon-type propellants

The gas-side heat transfer rates for LOX/propane and LOX/ethanol are experimentally characterized using a 1000 lb thrust water-cooled calorimeter chamber. The effects of injector element type and fuel film cooling are defined as a function of mixture ratio. The interaction of fuel injected through the resonator cavities on heat transfer and wall soot buildup are displayed as a function of time, axial distance, fuel coolant flow rate, and mixture ratio. Comparisons between clean-burning ethanol and sooting propane show a large difference between the two fuels and significantly higher than expected heat flux levels for ethanol in the throat region.

Schoenman, Leonard↗
NASA NTRS1988

High-pressure calorimeter chamber tests for liquid oxygen/kerosene (LOX/RP-1) rocket combustion

An experimental program was conducted to investigate the rocket combustion and heat transfer characteristics of liquid oxygen/kerosene (LOX/RP-1) mixtures at high chamber pressures. Two water-cooled calorimeter chambers of different combustion lengths were tested using 37- and 61-element oxidizer-fuel-oxidizer triplet injectors. The tests were conducted at nominal chamber pressures of 4.1, 8.3, and 13.8 MPa abs (600, 1200, and 2000 psia). Heat flux Q/A data were obtained for the entire calorimeter length for oxygen/fuel mixture ratios of 1.8 to 3.3. Test data at 4.1 MPa abs compared favorably with previous test data from another source. Using an injector with a fuel-rich outer zone reduced the throat heat flux by 47 percent with only a 4.5 percent reduction in the characteristic exhaust velocity efficiency C* sub eff. The throat heat transfer coefficient was reduced approximately 40 percent because of carbon deposits on the chamber wall.

Masters, Philip A.↗
NASA NTRS2001

Optimization of a GO2/GH2 Impinging Injector Element

An injector optimization methodology, method i, is used to investigate optimal design points for a gaseous oxygen/gaseous hydrogen (GO2/GH2) impinging injector element. The unlike impinging element, a fuel-oxidizer- fuel (F-O-F) triplet, is optimized in terms of design variables such as fuel pressure drop, (Delta)P(sub f), oxidizer pressure drop, (Delta)P(sub o), combustor length, L(sub comb), and impingement half-angle, alpha, for a given mixture ratio and chamber pressure. Dependent variables such as energy release efficiency, ERE, wall heat flux, Q(sub w), injector heat flux, Q(sub inj), relative combustor weight, W(sub rel), and relative injector cost, C(sub rel), are calculated and then correlated with the design variables. An empirical design methodology is used to generate these responses for 163 combinations of input variables. Method i is then used to generate response surfaces for each dependent variable. Desirability functions based on dependent variable constraints are created and used to facilitate development of composite response surfaces representing some, or all, of the five dependent variables in terms of the input variables. Three examples illustrating the utility and flexibility of method i are discussed in detail. First, joint response surfaces are constructed by sequentially adding dependent variables. Optimum designs are identified after addition of each variable and the effect each variable has on the design is shown. This stepwise demonstration also highlights the importance of including variables such as weight and cost early in the design process. Secondly, using the composite response surface which includes all five dependent variables, unequal weights are assigned to emphasize certain variables relative to others. Here, method i is used to enable objective trade studies on design issues such as component life and thrust to weight ratio. Finally, specific variable weights are further increased to illustrate the high marginal cost of realizing the last increment of injector performance and thruster weight.

Tucker, P. Kevin↗
NASA NTRS1975

Estimate of contribution of jet aircraft operations to trace element concentration at or near airports

Samples of ASTM type A jet fuel were analyzed for trace element content by instrumental neutron activation techniques. Forty-nine elements were sought. Only ten, aluminum, gold, indium, lanthanum, titanium, vanadium, barium, dysprosium, tellurium, and uranium, were observed at levels above the detection limits encountered; of these only aluminum, titanium, and barium were present at concentrations greater than 0.1 ppm. Estimates of exhaust gas concentrations are made, and the ambient contribution at or near airports is calculated by using the Los Angeles International Airport dispersion model. It is shown that the ambient contribution is about an order of magnitude below typical urban levels for virtually all elements sought.

Fordyce, J. S.↗
NASA NTRS1992

Water and cheese from the lunar desert: Abundances and accessibility of H, C, and N on the Moon

The Moon has been underrated as a source of H, N, C, and other elements essential to support life and to provide fuel for rockets. There is enough of these elements in each cubic meter of typical lunar soil to provide a substantial lunch for two, if converted to edible forms. The average amount of C per square meter of the lunar surface to a depth of 2 m is some 35 percent of the average amount per square meter tied up in living organisms on Earth. The water equivalent of H in the upper 2 m of the regolith averages at least 1.3 million liters per square kilometer. Mining of H from a small fraction of the regolith would provide all the rocket fuel needed for thousands of years. These elements can be removed from the soil by heating it to high temperature. Some favor the unproven resources of Phobos, Deimos, or near-Earth asteroids instead of the Moon as a source of extraterrestrial material for use in space, or Mars over the Moon as a site for habitation, partly on the basis that the chemical elements needed for life support and propellant are readily abundant on those bodies, but not on the Moon. Well, the Moon is not as barren of H, C, and N as is commonly perceived. In fact, the elements needed for life support and for rocket fuel are plentiful there, although the ore grades are low. Furthermore, the proximity of the Moon and consequent lower cost of transportation and shorter trip and communication times favor that body as the logical site for early acquisition of resources and extraterrestrial living.

Haskin, L. A.↗
NASA NTRS1978

Study of fuel systems for LH2-fueled subsonic transport aircraft, volume 1

Several engine concepts examined to determine a preferred design which most effectively exploits the characteristics of hydrogen fuel in aircraft tanks received major emphasis. Many candidate designs of tank structure and cryogenic insulation systems were evaluated. Designs of all major elements of the aircraft fuel system including pumps, lines, valves, regulators, and heat exchangers received attention. Selected designs of boost pumps to be mounted in the LH2 tanks, and of a high pressure pump to be mounted on the engine were defined. A final design of LH2-fueled transport aircraft was established which incorporates a preferred design of fuel system. That aircraft was then compared with a conventionally fueled counterpart designed to equivalent technology standards.

Brewer, G. D.↗
NASA NTRS1989

Finite element analysis of the Space Shuttle Main Engine (SSME) high pressure fuel turbopump turbine blade (HPFTP)

Cracking of the turbine blades of the SSME HPFTP is studied using two- and three-dimensional finite element analysis. The development and composition of the two- and three-dimensional models are described. Analyses are conducted under the speed, pressure, and thermal load conditions that occur during the full power level of the engine. The effects of friction on the two-dimensional model are examined. The strain and life cycle data reveal that the LCF cracking in the first stage is not probable unless the effects of fit-up tolerance between the blade and rotor are present, and for the second stage it is predicted that hydrogen assisted LCF cracking will occur under the present thermal environment. Design modifications to alleviate this cracking are discussed.

Lee, H. M.↗
NASA NTRS1974

Estimate of contribution of jet aircraft operations to trace element concentration at or near airports

Samples of ASTM type A jet fuel were analyzed for trace-element content by instrumental neutron activation techniques. Forty-nine elements were sought. Only ten, aluminum, gold, indium, lanthanum, titanium, vandium, barium, dysprosium, tellurium, and uranium, were observed at levels above the detection limits encountered; of these only aluminum, titanium, and barium were present at concentrations greater than 0.1 ppm. Estimates of exhaust gas concentrations are made, and the ambient contribution at or near airports is calculated by using the Los Angeles International Airport dispersion model. It is shown that the ambient contribution is about an order of magnitude below typical urban levels for virtually all elements sought.

Fordyce, J. S.↗
NASA NTRS1996

Structural Evaluation of a Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump Turbine Blade

Thermal and structural finite-element analyses were performed on the first high pressure fuel turbopump turbine blade of the space shuttle main engine (SSME). A two-dimensional (2-D) finite-element model of the blade and firtree disk attachment was analyzed using the general purpose MARC (finite-element) code. The loading history applied is a typical test stand engine cycle mission, which consists of a startup condition with two thermal spikes, a steady state and a shutdown transient. The blade material is a directionally solidified (DS) Mar-M 246 alloy, the blade rotor is forged with waspalloy material. Thermal responses under steady-state and transient conditions were calculated. The stresses and strains under the influence of mechanical and thermal loadings were also determined. The critical regions that exhibited high stresses and severe localized plastic deformation were the blade-rotor gaps.

Abdul-Aziz, Ali↗
NASA NTRS2012

The Single Crew Module Concept for Exploration

Many concepts have been proposed for exploring space. In early 2010 presidential direction called for reconsidering the approach to address changes in exploration destinations, use of new technologies and development of new capabilities to support exploration of space. Considering the proposed new technology and capabilities that NASA was directed to pursue, the single crew module (SCM) concept for a more streamlined approach to the infrastructure and conduct of exploration missions was developed. The SCM concept combines many of the new promising technologies with a central concept of mission architectures that uses a single habitat module for all phases of an exploration mission. Integrating mission elements near Earth and fully fueling them prior to departure of the vicinity of Earth provides the capability of using the single habitat both in transit to an exploration destination and while exploring the destination. The concept employs the capability to return the habitat and interplanetary propulsion system to Earth vicinity so that those elements can be reused on subsequent exploration missions. This paper describes the SCM concept, provides a top level mass estimate for the elements needed and trades the concept against Many concepts have been proposed for exploring space. In early 2010 presidential direction called for reconsidering the approach to address changes in exploration destinations, use of new technologies and development of new capabilities to support exploration of space. Considering the proposed new technology and capabilities that NASA was directed to pursue, the single crew module (SCM) concept for a more streamlined approach to the infrastructure and conduct of exploration missions was developed. The SCM concept combines many of the new promising technologies with a central concept of mission architectures that uses a single habitat module for all phases of an exploration mission. Integrating mission elements near Earth and fully fueling them prior to departure of the vicinity of Earth provides the capability of using the single habitat both in transit to an exploration destination and while exploring the destination. The concept employs the capability to return the habitat and interplanetary propulsion system to Earth vicinity so that those elements can be reused on subsequent exploration missions. This paper describes the SCM concept, provides a top level mass estimate for the elements needed and trades the concept against Constellation approaches for Lunar, Near Earth Asteroid and Mars Surface missions.

Chambliss, Joe↗
NASA NTRS1989

Meteoroid and orbital debris shielding on the Orbital Maneuvering Vehicle

NASA's Orbital Maneuvering Vehicle (OMV) is being designed to withstand a 10-year lifetime in polar and low earth orbits. A large percentage of OMV's lifetime will be spent operating in the vicinity of the Space Shuttle and Space Station or in storage at these manned locations. An extensive analysis has been performed to determine the effects of the meteoroid and orbital debris environments on OMV's external fuel tanks. A finite element model of OMV was constructed using NASTRAN and analyzed with the meteoroid and debris design analysis code BUMPER. The results show that the long design lifetime, and the ever increasing man-made orbital debris environment, will require the use of shielding over the external fuel tanks.

Kirkpatrick, Marc E.↗
NASA NTRS1993

CFD simulation of coaxial injectors

The development of improved performance models for the Space Shuttle Main Engine (SSME) is an important, ongoing program at NASA MSFC. These models allow prediction of overall system performance, as well as analysis of run-time anomalies which might adversely affect engine performance or safety. Due to the complexity of the flow fields associated with the SSME, NASA has increasingly turned to Computational Fluid Dynamics (CFD) techniques as modeling tools. An important component of the SSME system is the fuel preburner, which consists of a cylindrical chamber with a plate containing 264 coaxial injector elements at one end. A fuel rich mixture of gaseous hydrogen and liquid oxygen is injected and combusted in the chamber. This process preheats the hydrogen fuel before it enters the main combustion chamber, powers the hydrogen turbo-pump, and provides a heat dump for nozzle cooling. Issues of interest include the temperature and pressure fields at the turbine inlet and the thermal compatibility between the preburner chamber and injector plate. Performance anomalies can occur due to incomplete combustion, blocked injector ports, etc. The performance model should include the capability to simulate the effects of these anomalies. The current approach to the numerical simulation of the SSME fuel preburner flow field is to use a global model based on the MSFC sponsored FNDS code. This code does not have the capabilities of modeling several aspects of the problem such as detailed modeling of the coaxial injectors. Therefore, an effort has been initiated to develop a detailed simulation of the preburner coaxial injectors and provide gas phase boundary conditions just downstream of the injector face as input to the FDNS code. This simulation should include three-dimensional geometric effects such as proximity of injectors to baffles and chamber walls and interaction between injectors. This report describes an investigation into the numerical simulation of GH2/LOX coaxial injectors. The following sections will discuss the physical aspects of injectors, the CFD code employed, and preliminary results of a simulation of a single coaxial injector for which experimental data is available. It is hoped that this work will lay the foundation for the development of a unique and useful tool to support the SSME program.

Landrum, D. Brian↗
NASA NTRS2018

Performance, Stability and Compatibility of Oxygen/RP-1 Multi-Element Oxidizer-Rich Staged-Combustion Injectors

In 2015 and 2016, the National Aeronautics and Space Administration Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/RP-1 propellant multi-element oxidizer-rich staged-combustion test article. The main objective was to provide thrust chamber combustion stability data as part of the Combustion Stability Tool Development program, although demonstration of performance and compatibility of oxidizer-rich main injectors was also important. Funding was provided by the Air Force Space and Missile Systems Center. Five configurations of main injectors were designed and fabricated, using conventional gas-centered swirl coaxial injector element designs generally similar to those used in oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines. Variations of element features included element size, recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. Ablative combustion chambers were fabricated based on hardware previously used at the NASA MSFC for testing at similar size and pressure. Existing oxygen/RP-1 oxidizer-rich subscale preburner injectors and hot gas ducts from a previous NASA-funded program were modified for use to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Testing of the resulting integrated test article - which included the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was conducted at Test Stand 116 at the East Test Area of the NASA MSFC. The test article was well instrumented with static and dynamic pressure, temperature, and vibration sensors. This paper presents and discusses all the hot-fire test results of the integrated test article thrust chamber. Eighteen successful hot-fire tests of the integrated rig were conducted. Testing was accomplished with all five of the injector element concepts. Main combustion chamber pressures ranged from 710 to 2350 psia, and main combustion chamber mixture ratios ranged from 2.47 to 2.87. A chamber barrier fuel film coolant of about 2% to 4% of the total fuel flow was used for most tests. Characteristic exhaust velocity efficiency excluding the influence of the fuel film cooling ranged from 91% to 98% of theoretical. All tests of the thrust chamber exhibited stable combustion, even down to 40% of nominal operating pressures. Compatibility of the injector face and combustion chamber walls was acceptable. This paper is a follow-on to publication of preliminary test data presented at the 2016 JANNAF Liquid Propulsion Subcommittee meeting.

Hulka, J. R.↗