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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

Performance tests of a gas blending system based on mass-flow controllers

The system provides many of the gas mixtures required for calibrating analytical instruments used in engine exhaust gas analysis and is capable of blending from one to four additive gases with either of two carrier gases in concentrations from 20 ppm to 50%. Two mixtures can be flowing simultaneously. Performance tests were made to determine the stability accuracy of the system while it was in limited use for a period of 2 years. The accuracy of the blender was measured by comparing binary mixtures from the blender with National Bureau of Standards standard reference materials. Analytical instruments were used to make these comparisons. The expected accuracy of 2% was obtained in some of the tests, by the majority showed a systematic biAs of -5%. Although these tests revealed subtle instabilities in the flow controllers that contributed to the random scatter of data, the accuracy of wet test meters and bubble flowmeters used for calibration is marginal for this purpose. A simple procedure is recommended that should enable the full potential of the system to be realized.

Evans, A., Jr.↗
NASA NTRS1990

CO-oxidation catalysts: Low-temperature CO oxidation over Noble-Metal Reducible Oxide (NMRO) catalysts

Oxidation of CO to CO2 is an important reaction technologically and environmentally and a complex and interesting reaction scientifically. In most cases, the reaction is carried out in order to remove CO as an environmental hazard. A major application of heterogeneous catalysts is catalytic oxidation of CO in the exhaust of combustion devices. The reaction over catalysts in exhaust gas is fast and often mass-transfer-limited since exhaust gases are hot and O2/CO ratios are high. The main challenges to catalyst designers are to control thermal sintering and chemical poisoning of the active materials. The effect of the noble metal on the oxide is discussed, followed by the effect of the oxide on the noble metal, the interaction of the noble metal and oxide to form unique catalytic sites, and the possible ways in which the CO oxidation reaction is catalyzed by the NMRO materials.

Herz, Richard K.↗
NASA NTRS2019

Design of a Nozzle for the Spyder 2nd Stage Solid Rocket Motor

During the 2019 summer term, the author worked with a group of four interns to complete the preliminary design of a 2nd stage solid rocket motor for Up Aerospace’s Spyder Launch Vehicle. The Spyder vehicle is a four stage, solid fuel rocket designed as part of collaboration between NASA and Up Aerospace to develop a vehicle capable of delivering a 10 kg, 6U CubeSat into a 350 km, circular, low Earth orbit. As part of the agreement, NASA is tasked with designing high performance 2nd, 3rd, and 4th stages while Up Aerospace will provide the 1st stage, derived from the first stage of the company’s sub-orbital Spaceloft XL vehicle. Previous intern teams have designed the 3rd and 4th stages, which left the preliminary design of the 2nd stage motor to be completed this summer. The purpose of this report is to highlight a trade study which the author conducted to determine the nozzle geometry which would most benefit the performance of the 2nd stage motor. In this study, various nozzle parameters such as throat radius (RSI), expansion ratio, mass and their effects on the Isp and Delta V of the 2nd Stage were investigated. From this study, a nozzle geometry providing the necessary performance was chosen and implemented as part of the preliminary design of the 2nd stage motor. To mature the 2nd stage motor design, a trade space was needed to determine the nozzle configuration which would most benefit the performance of the 2nd stage. The trade space established did not only evaluate different expansion ratios for the same throat radius, but also investigated the possible performance gained from decreasing the throat radius to increase the expansion ratio and Isp capable of being delivered by the nozzle. Decreasing the throat radius would cause the chamber pressure to increase, consequently increasing the case and insulation mass required to safely operate a motor at higher pressures. To account for this factor, accurate estimates of inert mass first needed to be established. After doing so, the effects of varying nozzle expansion ratios, exit half angles, and subsequently length and mass were evaluated against motor and nozzle performance factors such as delta V and Isp. For this study, four throat radii ranging from 1.75” to 2.375” and consequently four different chamber pressures ranging 550 psia to 1200 psia were investigated. BACKGROUND To launch into Low Earth Orbit, a payload needs to be accelerated to the orbital velocity necessary to keep it from falling back to Earth. The change in velocity required between launch and orbital insertion is known as Delta V. The Delta V which a rocket or stage can deliver can be calculated using the Ideal Rocket equation, 𝛥𝑉=−𝑔0∗𝐼𝑠𝑝∗ln(𝑀𝑓𝑀𝑖) (3) Where 𝑔0 is the acceleration due to gravity at the earth’s surface, 𝐼𝑠𝑝 is the specific impulse of the rocket, 𝑀𝑖 is the initial mass of the rocket, and 𝑀𝑓 is the final mass of the rocket after burnout. From preliminary calculations beyond the scope of this paper, it was determined that 30500 ft/s of delta V would be required for a payload to be inserted into a 350 km circular orbit around the Earth. Using the known masses and Isp values of the 1st, 3rd, and 4th stages and equation 3, the delta V of each stage was calculated. The delta V required by the 2nd stage could then be found by taking the difference between the total delta V required and the delta V of the 1st, 3rd, and 4th stages. From this, the required delta V of the 2nd Stage was calculated to be 7340 ft/s. Specific impulse is an efficiency factor of the nozzle which defines the impulse delivered by the motor per unit of propellant weight. The main variables of a nozzle’s specific impulse investigated in this trade were exit cone half angle, throat radius, and expansion ratio which is affected by the throat radius. The expansion ratio, ε, of a nozzle is defined as the ratio between the nozzle exit area and throat area, and can be calculated using the equation, ε=𝑅𝑒𝑥𝑖𝑡2𝑅𝑠𝑖2 (2) Where 𝑅𝑒𝑥𝑖𝑡 is the radius of the nozzle’s exit and 𝑅𝑠𝑖 is the radius of the nozzle’s throat. A larger expansion ratio and smaller exit half angle will increase the Isp of a nozzle by allowing the gas to expand more and by allowing more of the exhaust gas to produce thrust in the direction of the motor’s central axis. A cross section view of the 2nd Stage motor with the major components annotated is provided in figure 1.

Bennett, Daniel↗
NASA NTRS1980

Computational models for the viscous/inviscid analysis of jet aircraft exhaust plumes

Computational models which analyze viscous/inviscid flow processes in jet aircraft exhaust plumes are discussed. These models are component parts of an NASA-LaRC method for the prediction of nozzle afterbody drag. Inviscid/shock processes are analyzed by the SCIPAC code which is a compact version of a generalized shock capturing, inviscid plume code (SCIPPY). The SCIPAC code analyzes underexpanded jet exhaust gas mixtures with a self-contained thermodynamic package for hydrocarbon exhaust products and air. A detailed and automated treatment of the embedded subsonic zones behind Mach discs is provided in this analysis. Mixing processes along the plume interface are analyzed by two upgraded versions of an overlaid, turbulent mixing code (BOAT) developed previously for calculating nearfield jet entrainment. The BOATAC program is a frozen chemistry version of BOAT containing the aircraft thermodynamic package as SCIPAC; BOATAB is an afterburning version with a self-contained aircraft (hydrocarbon/air) finite-rate chemistry package. The coupling of viscous and inviscid flow processes is achieved by an overlaid procedure with interactive effects accounted for by a displacement thickness type correction to the inviscid plume interface.

Dash, S. M.↗
NASA NTRS1973

Odor intensity and characterization of jet exhaust and chemical analytical measurements

Odor and chemical analyses were carried out on the exhaust samples from a J-57 combustor can operated over a range of inlet conditions, and with several fuel types and nozzle modifications. The odor characteristics and total intensity of odor for each exhaust were determined over a range of dilutions to allow for a least squares determination of the intensity at 1,000 to 1 dilutions. Analytical measures included the concentration of total hydrocarbons and the concentrations of aromatic organic species and oxygenated organic species from collected samples which were taken concurrently. A correlation was found between the concentration of the odorous oxygenated fraction and the total intensity of aroma. Inlet operating conditions and nozzle modifications which increase the efficiency of combustion as measured by exhaust gas analyses reduce the odor intensity and the quantity of oxygenates in the exhaust. The type of fuel burned altered the intensity of odor in relation to the quantity of oxygenates produced and, in some instances, changed the odor character.

Kendall, D. A.↗
NASA NTRS2011

Alternative Aviation Fuel Experiment (AAFEX)

The rising cost of oil coupled with the need to reduce pollution and dependence on foreign suppliers has spurred great interest and activity in developing alternative aviation fuels. Although a variety of fuels have been produced that have similar properties to standard Jet A, detailed studies are required to ascertain the exact impacts of the fuels on engine operation and exhaust composition. In response to this need, NASA acquired and burned a variety of alternative aviation fuel mixtures in the Dryden Flight Research Center DC-8 to assess changes in the aircraft s CFM-56 engine performance and emission parameters relative to operation with standard JP-8. This Alternative Aviation Fuel Experiment, or AAFEX, was conducted at NASA Dryden s Aircraft Operations Facility (DAOF) in Palmdale, California, from January 19 to February 3, 2009 and specifically sought to establish fuel matrix effects on: 1) engine and exhaust gas temperatures and compressor speeds; 2) engine and auxiliary power unit (APU) gas phase and particle emissions and characteristics; and 3) volatile aerosol formation in aging exhaust plumes

Anderson, B. E.↗
NASA NTRS2011

The Development of a CO2 Test Capability in the NASA JSC ARCJet for Mars Reentry Simulation

The Atmospheric Reentry Materials and Structures Evaluation Facility (ARMSEF) located at NASA Johnson Space Center is used for simulating the extreme environment experienced upon reentry for the development and certification of thermal protection systems (TPS). The facility supports a large variety of programs and was heavily leveraged for the certification and operational support of the TPS for the Orbiter and, more recently, the development of the heat shield for CEV. This paper will provide more detail into the heritage of the facility. Unique attributes of the facility include a modular aerodynamically stabilized arc heater and independently controlled O2 and N2 for the test gases. When combining the O2 and N2 in a 23:77 mass ratio respectively the Earth s atmosphere is accurately simulated and via modification of this ratio the investigation of the effects of atomic oxygen on a material s response is possible. In the summer of 2010 a development effort was started to add CO2 as a third independently controlled test gas such that, when combined with N2, opens up the possibility of accurately simulating a Martian reentry environment. This paper will discuss the test facility, especially the arc heater, in more detail. Initial testing involved relatively low concentrations of CO2 combined with N2 for the primary purpose of gathering data to answer two pressing safety concerns. The first being the rate of production of carbon monoxide (CO) within the ejector vacuum system. The main concern was that CO can be flammable and possibly explosive at high enough concentrations and pressures. The hazard control during the development phase involved the use of injecting N2 inside the test chamber diffuser to dilute and reduce the concentration of any and all CO present. A residual gas analyzer (RGA) was used to determine the relative amount of CO in the exhaust gas and provide a conversion rate of CO2 to CO. This paper will discuss in more detail the results of the RGA data and the calculated conversion rate. The second safety concern addressed is the possible formation of hydrogen cyanide (HCN) and cyanide (CN). HCN would primarily be present in the cooling water while the CN would most probably condense onto the interior surfaces of the test chamber. Water samples and wipes of the test chamber surfaces were analyzed by an industrial hygienist for the presence of HCN and CN. His paper will discuss these results in more detail. Throughout this development effort measurements of the CO2:N2 flowfield were made with heat flux and pressure probes and with laser induced fluorescence (LIF) of the atomic oxygen. This paper will discuss these results

DelPapa, Steven V.↗
NASA NTRS1993

Canceling Torque Caused By Boiloff Of Cryogen

Proposed technique for cancellation of small torques caused by venting of cryogens from tanks in spacecraft adaptable to use in making small corrections in orientations of terrestrial scientific instruments. Torque eliminated by directing exhaust gas through "T" vent. To cancel remaining small torque, that side of "T" from which exhaust produces smaller thrust heated just enough to equalize opposing thrusts. Concept useful in situations in which conventional electromagnetic actuators not suitable because of constraints related to temperature, radiation, or vibration and equipment required to be robust, simple, reliable, light in weight, and without moving parts.

Bhandari, Pradeep↗
NASA NTRS1931

A Method for Reducing the Temperature of Exhaust Manifolds

This report describes tests conducted at the Langley Memorial Aeronautical Laboratory on an "air-inducting" exhaust manifold for aircraft engines. The exhaust gases from each cylinder port are discharged into the throat of an exhaust pipe which has a frontal bellmouth. Cooling air is drawn into the pipe, where it surrounds and mixes with the exhaust gases. Temperatures of the manifold shell and of the exhaust gases were obtained in flight for both a conventional manifold and the air-inducting manifold. The air-inducting manifold was installed on an engine which was placed on a test stand. Different fuels were sprayed on and into the manifold to determine whether the use of this manifold reduced the fire hazard. The flight tests showed reductions in manifold temperatures of several hundred degrees, to values below the ignition point of aviation gasoline. On the test stand when the engine was run at idling speeds fuels sprayed into the manifold ignited. It is believed that at low engine speeds the fuel remained in the manifold long enough to become thoroughly heated, and was then ignited by the exhaust gas which had not mixed with cooling air. The use of the air-inducting exhaust manifold must reduce the fire hazard by virtue of its lower operating temperature, but it is not a completely satisfactory solution of the problem.

Schey, Oscar W↗
NASA NTRS1951

Full-scale Investigation of Cooling Shroud and Ejector Nozzle for a Turbojet Engine : Afterburner Installation

A full-scale ejector cooling investigation was made on a turbojet engine - afterburner installation in the NACA Lewis altitude wind tunnel. Ejector performance was studied at primary exhaust-gas temperatures from 2700 degrees to 3400 degrees R (corresponding to ejector temperature ratios from 2.0 to 5.0), primary pressure ratios from 1.79 to 3.4, secondary air flows up to 29 percent of the primary gas flow, and for diameter ratios from 1.08 to 1.42 and spacing ratios from 0.04 to 1.16. In addition, variations were made in the primary exhaust-nozzle area. Ejectors with large diameter ratios permit the attainment of high gas flow ratios, but the jet-thrust losses become prohibitive as the spacing ratio is increased from 0 to 0.16. As the ejector diameter is reduced, the obtainable gas-flow ratio and the thrust loss are reduced. Previous results showing that data obtained at a temperature ratio of 1.0 could not be extrapolated to determine ejector performance at high temperature ratios by the application of the temperature ratio factor to the gas-flow ratios are substantiated by the present investigation.

AFTERBURNERS↗
NASA NTRS2002

Turbine Engine Clearance Control Systems: Current Practices and Future Directions

Improved blade tip sealing in the high pressure compressor (HPC) and high pressure turbine (HPT) can provide dramatic reductions in specific fuel consumption (SFC), time-on-wing, compressor stall margin, and engine efficiency as well as increased payload and mission range capabilities. Maintenance costs to overhaul large commercial gas turbine engines can easily exceed $1M. Engine removal from service is primarily due to spent exhaust gas temperature (EGT) margin caused mainly by the deterioration of HPT components. Increased blade tip clearance is a major factor in hot section component degradation. As engine designs continue to push the performance envelope with fewer parts and the market drives manufacturers to increase service life, the need for advanced sealing continues to grow. A review of aero gas turbine engine HPT performance degradation and the mechanisms that promote these losses are discussed. Benefits to the HPT due to improved clearance management are identified. Past and present sealing technologies are presented along with specifications for next generation engine clearance control systems.

Lattime, Scott B.↗
NASA NTRS1942

The Tracer Gas Method of Determining the Charging Efficiency of Two-stroke-cycle Diesel Engines

A convenient method has been developed for determining the scavenging efficiency or the charging efficiency of two-stroke-cycle engines. The method consists of introducing a suitable tracer gas into the inlet air of the running engine and measuring chemically its concentration both in the inlet and exhaust gas. Monomethylamine CH(sub 3)NH(sub 2) was found suitable for the purpose as it burns almost completely during combustion, whereas the "short-circuited" portion does not burn at all and can be determined quantitatively in the exhaust. The method was tested both on four-stroke and on two-stroke engines and is considered accurate within 1 percent.

Schweitzer, P H↗
NASA NTRS1957

Experimental Study of Heat Transfer to Small Cylinders in a Subsonic, High-temperature Gas Stream

A Nusselt-Reynolds number relation for cylindrical thermocouple wires in crossflow was obtained from the experimental determination of time constants. Tests were conducted in exhaust gas over a temperature range of 2000 to 3400 R, a Mach number range of 0.3 to 0.8, and a static-pressure range from 2/3 to 1-1/3 atmospheres, yielding a Reynolds number range of 450 to 3000. The correlation obtained is Nu=(0.428 plus or minus 0.003) times the square root of Re* with average deviations of a single observation of 8.5 percent. This relation is the same as one previously reported for room-temperature conditions.

GASES, PROPERTIES↗
NASA NTRS1998

Effects of Radiative Emission and Absorption on the Propagation and Extinction of Premixed Gas Flames

Premixed gas flames in mixtures of CH4, O2, N2, and CO2 were studied numerically using detailed chemical and radiative emission-absorption models to establish the conditions for which radiatively induced extinction limits may exist independent of the system dimensions. It was found that reabsorption of emitted radiation led to substantially higher burning velocities and wider extinction limits than calculations using optically thin radiation models, particularly when CO2, a strong absorber, is present in the unburned gas, Two heat loss mechanisms that lead to flammability limits even with reabsorption were identified. One is that for dry hydrocarbon-air mixtures, because of the differences in the absorption spectra of H2O and CO2, most of the radiation from product H2O that is emitted in the upstream direction cannot be absorbed by the reactants. The second is that the emission spectrum Of CO2 is broader at flame temperatures than ambient temperature: thus, some radiation emitted near the flame front cannot be absorbed by the reactants even when they are seeded with CO2 Via both mechanisms, some net upstream heat loss due to radiation will always occur, leading to extinction of sufficiently weak mixtures. Downstream loss has practically no influence. Comparison with experiment demonstrates the importance of reabsorption in CO2 diluted mixtures. It is concluded that fundamental flammability limits can exist due to radiative heat loss, but these limits are strongly dependent on the emission-absorption spectra of the reactant and product -gases and their temperature dependence and cannot be predicted using gray-gas or optically thin model parameters. Applications to practical flames at high pressure, in large combustion chambers, and with exhaust-gas or flue-gas recirculation are discussed.

Ju, Yiguang↗
NASA NTRS1947

Temperatures and Stresses on Hollow Blades For Gas Turbines

The present treatise reports on theoretical investigations and test-stand measurements which were carried out in the BMW Flugmotoren GMbH in developing the hollow blade for exhaust gas turbines. As an introduction the temperature variation and the stress on a turbine blade for a gas temperature of 900 degrees and circumferential velocities of 600 meters per second are discussed. The assumptions onthe heat transfer coefficients at the blade profile are supported by tests on an electrically heated blade model. The temperature distribution in the cross section of a blade Is thoroughly investigated and the temperature field determined for a special case. A method for calculation of the thermal stresses in turbine blades for a given temperature distribution is indicated. The effect of the heat radiation on the blade temperature also is dealt with. Test-stand experiments on turbine blades are evaluated, particularly with respect to temperature distribution in the cross section; maximum and minimum temperature in the cross section are ascertained. Finally, the application of the hollow blade for a stationary gas turbine is investigated. Starting from a setup for 550 C gas temperature the improvement of the thermal efficiency and the fuel consumption are considered as well as the increase of the useful power by use of high temperatures. The power required for blade cooling is taken into account.

Pollmann, Erich↗
NASA NTRS2004

Exhaust Nozzle for a Multitube Detonative Combustion Engine

An improved type of exhaust nozzle has been invented to help optimize the performances of multitube detonative combustion engines. The invention is applicable to both air-breathing and rocket engines used to propel some aircraft and spacecraft, respectively. In a detonative combustion engine, thrust is generated through the expulsion of combustion products from a detonation process in which combustion takes place in a reaction zone coupled to a shock wave. The combustion releases energy to sustain the shock wave, while the shock wave enhances the combustion in the reaction zone. The coupled shockwave/reaction zone, commonly referred to as a detonation, propagates through the reactants at very high speed . typically of the order of several thousands of feet per second (of the order of 1 km/s). The very high speed of the detonation forces combustion to occur very rapidly, thereby contributing to high thermodynamic efficiency. A detonative combustion engine of the type to which the present invention applies includes multiple parallel cylindrical combustion tubes, each closed at the front end and open at the rear end. Each tube is filled with a fuel/oxidizer mixture, and then a detonation wave is initiated at the closed end. The wave propagates rapidly through the fuel/oxidizer mixture, producing very high pressure due to the rapid combustion. The high pressure acting on the closed end of the tube contributes to forward thrust. When the detonation wave reaches the open end of the tube, it produces a blast wave, behind which the high-pressure combustion products are expelled from the tube. The process of filling each combustion tube with a detonable fuel/oxidizer mixture and then producing a detonation repeated rapidly to obtain repeated pulses of thrust. Moreover, the multiple combustion tubes are filled and fired in a repeating sequence. Hence, the pressure at the outlet of each combustion tube varies cyclically. A nozzle of the present invention channels the expansion of the pulsed combustion gases from the multiple combustion tubes into a common exhaust stream, in such a manner as to enhance performance in two ways: (1) It reduces the cyclic variations of pressure at the outlets of the combustion tubes so as to keep the pressure approximately constant near the optimum level needed for filling the tubes, regardless of atmospheric pressure at the altitude of operation; and (2) It maximizes the transfer of momentum from the exhaust gas to the engine, thereby maximizing thrust. The figure depicts a typical engine equipped with a nozzle according to the invention. The nozzle includes an interface section comprising multiple intake ports that couple the outlets of the combustion tubes to a common plenum. Proceeding from its upstream to its downstream end, the interface section tapers to a larger cross-sectional area for flow. This taper fosters expansion of the exhaust gases flowing from the outlets of the combustion tubes and contributes to the desired equalization of exhaust combustion pressure. The cross-sectional area for flow in the common plenum is greater than, or at least equal to, the combined cross-sectional flow areas of the combustor tubes. In the common plenum, the exhaust streams from the individual combustion tubes mix to form a single compound subsonic exhaust stream. Downstream of the common plenum is the throat that tapers to a smaller flow cross section. In this throat, the exhaust gases become compressed to form a compound sonic gas stream. Downstream of the throat is an expansion section, which typically has a bell or a conical shape. (The expansion section can be truncated or even eliminated in the case of an air-breathing engine.) After entering the expansion section, the exhaust gases expand rapidly from compound sonic to compound supersonic speeds and are then vented to the environment. The basic invention admits of numerous variations. For example, the combustion tubes can be arranged around the central axin a symmetrical or asymmetrical pattern other than the one shown in the figure. For another example, the flow cross-sectional area(s) of one or more of the intake ports in the interface section, of the common plenum, the throat, and/or the expansion section can be varied, either symmetrically or asymmetrically, to adjust dynamics of the exhaust stream or to direct the thrust vector away from the central axis.

Bratkovich, Thomas E.↗
NASA NTRS1998

Mixing of Supersonic Jets in a RBCC Strutjet Propulsion System

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during take-off and low speed flight. A scale model of the Strutjet device was built and tested to investigate the mixing of the streams as a function of distance from the Strut exit plane in simulated sea level take-off conditions. The Planar Laser Induced Fluorescence (PLIF) diagnostic method has been employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air. The ratio of the pressure in the turbine exhaust to that in the rocket nozzle wall at the point where the two jets meet, is the independent variable in these experiments. Tests were accomplished at values of 1.0 (the original design point), 1.5 and 2.0 for this parameter at 8 locations downstream of the rocket nozzle exit. The results illustrate the development of the mixing zone from the exit plane of the strut to a distance of about 18 equivalent rocket nozzle exit diameters downstream (18"). These images show the turbine exhaust to be confined until a short distance downstream. The expansion into the ingested air is more pronounced at a pressure ratio of 1.0 and 1.5 and shows that mixing with this air would likely begin at a distance of 2" downstream of the nozzle exit plane. Of the pressure ratios tested in this research, 2.0 is the best value for delaying the mixing at the operating conditions considered.

Muller, S.↗