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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 811 records · Page 45
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Comparison of aerogrids and punched plates for smoothing flow from short annular diffusers

Scale model tests were conducted to evaluate the effectiveness of aerogrids and punched plates in producing flat velocity profiles downstream of short diffusers as would be used between the compressor and combustor of advanced aircraft engines. The diffuser had an area ratio of 4.17 and a length-to-inlet-height ratio of 2.07. The aerogrids tested were plates containing 1123 contoured venturis in parallel with geometric blockages of 83, 74, and 61 percent, respectively. The punched plates contained 1123 sharp-edged orifices with blockages of 58 and 30 percent. The results show that aerogrids, with higher effective blockage for the same pressure loss, are more effective flow-smoothing devices than the punched plates. Also, the overall pressure loss decreases and the exit velocity profile becomes flatter as either type of grid is moved closer to the diffuser exit plane.

Rumpf, R. L.↗
NASA NTRS1976

Wing surface-jet interaction characteristics of an upper-surface blown model with rectangular exhaust nozzles and a radius flap

The wing surface jet interaction characteristics of an upper surface blown transport configuration were investigated in the Langley V/STOL tunnel. Velocity profiles at the inboard engine center line were measured for several chordwise locations, and chordwise pressure distributions on the flap were obtained. The model represented a four engine arrangement having relatively high aspect ratio rectangular spread, exhaust nozzles and a simple trailing edge radius flap.

Bloom, A. M.↗
NASA NTRS1978

Evaluation of FIDC system

A fuel vapor injector/igniter system was evaluated for its effect on automobile engine performance, fuel economy, and exhaust emissions. Improved fuel economy and emissions, found during the single cylinder tests were not realized with a multicylinder engine. Multicylinder engine tests were conducted to compare the system with both a stock and modified stock configuration. A comparison of cylinder-to-cylinder equivalence ratio distribution was also obtained from the multicylinder engine tests. The multicylinder engine was installed in a vehicle was tested on a chassis dynamometer to compare the system with stock and modified stock configurations. The fuel vapor injector/igniter system (FIDC) configuration demonstrated approximately five percent improved fuel economy over the stock configuration, but the modified stock configuration demonstrated approximately twelve percent improved fuel economy. The hydrocarbon emissions were approximately two-hundred-thirty percent higher with the FIDC system than with the stock configuration. Both the FIDC system and the modified stock configuration adversely affected driveability. The FIDC system demonstrated a modest fuel savings, but with the penalty of increased emissions, and loss of driveability.

Hall, R. A.↗
NASA NTRS1978

Static tests of the J97 powered, external augmenter V/STOL wind tunnel model

Results of the static testing (zero forward speed) of the J97-powered, external augmentor, large scale, V/STOL model are discussed. With a ground clearance of 7.5 feet, believed to have put the model essentially out of ground effect, a gross thrust augmentation ratio of 1.60 at nozzle pressure ratio (NPR) = 3.0 was measured for the fuselage augmentor. A similar figure was apparent for the wing augmentor. An overall ratio of model thrust to bare engine thrust of 1.52 was determined at NPR = 3.0. The structural integrity of the model was well demonstrated and duct pressure losses were small.

Garland, D. B.↗
NASA NTRS1983

Baseline performance and emissions data for a single-cylinder, direct-injected diesel engine

Comprehensive fuel consumption, mean effective cylinder pressure, and emission test results for a supercharged, single-cylinder, direct-injected, four-stroke-cycle, diesel test engine are documented. Inlet air-to-exhaust pressure ratios were varied from 1.25 to 3.35 in order to establish the potential effects of turbocharging techniques on engine performance. Inlet air temperatures and pressures were adjusted from 34 to 107 C and from 193 to 414 kPa to determine the effects on engine performance and emissions. Engine output ranged from 300 to 2100 kPa (brake mean effective pressure) in the speed range of 1000 to 3000 rpm. Gaseous and particulate emission rates were measured. Real-time values of engine friction and pumping loop losses were measured independently and compared with motored engine values.

Dezelick, R. A.↗
NASA NTRS1984

Improving the efficiency of smaller transport aircraft

The efficiency of small transport aircraft can be improved through the adaptation of high altitude turbine engines, and that flights reaching altitudes of 40,000, 60,000, and 80,000 feet can show savings in both flight time and fuel consumption even for trips as short as 500 miles. Studies for a 40-passenger high altitude transport are presented. An increase in structural weight due to larger wing areas, larger engines, and larger engine frontal areas would make the ratio of gross weight to payload look less favorable, but the efficiency of the plane in passenger miles per gallon would increase with altitude. It is also suggested that supercritical airfoils be designed to achieve higher lift coefficients and speeds. A reduction of reduced drag through the use of horizontal or vertical wing tip extensions is also discussed.

Jones, R. T.↗
NASA NTRS1984

A shock wave capability for the improved Two-Dimensional Kinetics (TDK) computer program

The Two Dimensional Kinetics (TDK) computer program is a primary tool in applying the JANNAF liquid rocket engine performance prediction procedures. The purpose of this contract has been to improve the TDK computer program so that it can be applied to rocket engine designs of advanced type. In particular, future orbit transfer vehicles (OTV) will require rocket engines that operate at high expansion ratio, i.e., in excess of 200:1. Because only a limited length is available in the space shuttle bay, it is possible that OTV nozzles will be designed with both relatively short length and high expansion ratio. In this case, a shock wave may be present in the flow. The TDK computer program was modified to include the simulation of shock waves in the supersonic nozzle flow field. The shocks induced by the wall contour can produce strong perturbations of the flow, affecting downstream conditions which need to be considered for thrust chamber performance calculations.

Nickerson, G. R.↗
NASA NTRS1987

Experimental and theoretical study of propeller spinner/shank interference

A fundamental investigation into the aerodynamic interference associated with propeller spinner and shank regions has been conducted. The research program involved a theoretical assessment of solutions previously proposed, followed by a systematic experimental study to supplement the existing data base. As a result, a refined computational procedure has been established for prediction of interference effects in terms of either interference drag or propeller thrust and torque coefficients. These quantities have been examined with attention to engineering parameters such as spinner finess ratio, blade shank form, and number of blades. Also, cascade effects and spinner/shank juncture interference have been semi-empirically modeled using existing theories and placed into a compatible form with an existing propeller performance code.

Cornell, C. C.↗
NASA NTRS1989

2-D and 3-D mixing flow analyses of a scramjet-afterbody configuration

A cold simulant gas study of propulsion/airframe integration for a hypersonic vehicle powered by a scramjet engine is presented. The specific heat ratio of the hot exhaust gases are matched by utilizing a cold mixture of argon and Freon-12. Solutions are obtained for a hypersonic corner flow and a supersonic rectangular flow in order to provide the upstream boundary conditions. The computational test examples also provide a comparison of this flow with that of air as the expanding supersonic jet, where the specific heats are assumed to be constant. It is shown that the three-dimensional computational fluid capabilities developed for these types of flow may be utilized to augment the conventional wind tunnel studies of scramjet afterbody flows using cold simulant exhaust gases, which in turn can help in the design of a scramjet internal-external nozzle.

Baysal, Oktay↗
NASA NTRS1991

Planar imaging of the mixing characteristics of a multi-element triplet injector

Planar light scattering has been used to investigate the turbulent mixing in a multi-element triplet injector at mixture ratio conditions typical of liquid rocket engine gas generators. Experiments were conducted at atmospheric pressure using air seeded with smoke particles to simulate the oxidizer flow and unseeded air to simulate the fuel. Data sets of 200 images (containing 122,000 simultaneous measurements each) of the instantaneous distribution of the scattered light were acquired at several positions throughout the flow. These images have been processed digitally to produce local distributions of the average and rms fluctuation of the mixture fraction. In addition, the spatial autocorrelation of the images has been calculated and analyzed to determine a mean length scale related to the large scale structures in the flow.

Makel, D. B.↗
NASA NTRS1952

Dynamic response at altitude of a turbojet engine with variable area exhaust nozzle

The dynamic characteristics of turbojet engine with variable exhaust nozzle area were investigated over a range of altitudes and flight Mach numbers. These characteristics generalize to standard static level condition. Data resulting from approximate step disturbances in either manipulated variable suggested functional relationships from which functions can be derived. The constants of the transfer functions are listed for five combinations of engine speed and exhaust nozzle area. The minimum data needed to completely define the transfer functions were: experimentally determined dynamic characteristics, such as engine time constant and initial rise ratio, resulting from either manipulated variable; static characteristics determined from steady-state performance curves.

Delio, Gene J↗
NASA NTRS2007

Experimental Results of Performance Tests on a Four-Port Wave Rotor

A series of tests has been performed on a four-port wave rotor suitable for use as a topping stage on a gas turbine engine, to measure the overall pressure ratio obtainable as a function of temperature ratio, inlet mass flow, loop flow ratio, and rotor speed. The wave rotor employed an open high pressure loop that is the high pressure inlet flow was not the air exhausted from the high pressure outlet, but was obtained from a separate heated source, although the mass flow rates of the two flows were balanced. This permitted the choice of a range of loop-flow ratios (i.e., ratio of high pressure flow to low pressure flow), as well as the possibility of examining the effect of mass flow imbalance. Imbalance could occur as a result of leakage or deliberate bleeding for cooling air. Measurements of the pressure drop in the high pressure loop were also obtained. A pressure ratio of 1.17 was obtained at a temperature ratio of 2.0, with an inlet mass flow of 0.6 lb/s. Earlier tests had given a pressure ratio of less than 1.12. The improvement was due to improved sealing between the high pressure and low pressure loops, and a modification to the movable end-wall which is provided to allow for rotor expansion.

Wilson, John↗
NASA NTRS2010

Space Shuttle Main Engine (SSME) Operational Capability

Through the years of the Space Shuttle Main Engine (SSME) program the engine has evolved and operational capabilities have been demonstrated beyond the original Shuttle requirements. In an effort to enhance flight safety and demonstrate safety features and margins, engines have been analyzed and tested at many different operating points. Various studies through the years evaluating the SSME for different applications both as a boost stage and upper stage have also added insight into the overall operational characteristics of the engine and have further defined safety margins for the Shuttle application. This paper will summarize the operational characteristics of the SSME from the original design requirements to the expanded capabilities demonstrated through analysis, lab testing and especially "off-nominal" engine testing leading to an increased understanding of the engine operational characteristics and safety margins. Basic engine characteristics such as thrust, mixture ratio, propellant inlet conditions, system redundancy, etc. will be examined.

Benefield, Philip↗
NASA NTRS2012

Space Launch System Accelerated Booster Development Cycle

With the retirement of the Space Shuttle, NASA is seeking to reinvigorate the national space program and recapture the public s interest in human space exploration by developing missions to the Moon, near-earth asteroids, Lagrange points, Mars, and beyond. The would-be successor to the Space Shuttle, NASA s Constellation Program, planned to take humans back to the Moon by 2020, but due to budgetary constraints was cancelled in 2010 in search of a more "affordable, sustainable, and realistic" concept2. Following a number of studies, the much anticipated Space Launch System (SLS) was unveiled in September of 2011. The SLS core architecture consists of a cryogenic first stage with five Space Shuttle Main Engines (SSMEs), and a cryogenic second stage using a new J-2X engine3. The baseline configuration employs two 5-segment solid rocket boosters to achieve a 70 metric ton payload capability, but a new, more capable booster system will be required to attain the goal of 130 metric tons to orbit. To this end, NASA s Marshall Space Flight Center recently released a NASA Research Announcement (NRA) entitled "Space Launch System (SLS) Advanced Booster Engineering Demonstration and/or Risk Reduction." The increased emphasis on affordability is evident in the language used in the NRA, which is focused on risk reduction "leading to an affordable Advanced Booster that meets the evolved capabilities of SLS" and "enabling competition" to "enhance SLS affordability. The purpose of the work presented in this paper is to perform an independent assessment of the elements that make up an affordable and realistic path forward for the SLS booster system, utilizing advanced design methods and technology evaluation techniques. The goal is to identify elements that will enable a more sustainable development program by exploring the trade space of heavy lift booster systems and focusing on affordability, operability, and reliability at the system and subsystem levels5. For this study, affordability is defined as lifecycle cost, which includes design, development, test, and engineering (DDT&E), production and operational costs (P&O). For this study, the system objectives include reducing DDT&E schedule by a factor of three, showing 99.9% reliability, flying up to four times per year, serving both crew and cargo missions, and evolving to a lift capability of 130 metric tons.3 After identifying gaps in the current system s capabilities, this study seeks to identify non-traditional and innovative technologies and processes that may improve performance in these areas and assess their impacts on booster system development. The DDT&E phase may be improved by incorporating incremental development testing and integrated demonstrations to mitigate risk. To further reduce DDT&E, this study will also consider how aspects of the booster system may have commonality with other users, such as the Department of Defense, commercial applications, or international partners; by sharing some of the risk and investment, the overall development cost may be reduced. Consideration is not limited to solid and liquid rocket boosters. A set of functional performance characteristics, such as engine thrust, specific impulse (Isp), mixture ratio, and throttle range are identified and their impacts on the system are evaluated. This study also identifies how such characteristics affect overall life cycle cost, including DDT&E and fixed and variable P&O.

Arockiam, Nicole↗
NASA NTRS2018

Design of Distributed Mail-Slot Propulsion System on a Hybrid Wingbody Aircraft

This paper deals with designing a thrust distribution strategy when a Turboelectric Distributed Propulsion (TeDP) system of 16 embedded propulsors is installed on an aerodynamically optimized hybrid wing-body configuration. This HWB previously designed to satisfy conditions of trim, longitudinally static stability and specific cargo space is employed as the baseline configuration for the current study of seeking an optimal propulsion/power system. According to the nature of the entrance flow condition for each distributed propulsion passage in hybrid wing-body aircraft, the ingested boundary layer thickness differs and results in different propulsive reaction. An optimal distribution of thrust and power output is determined by how the system utilizes the propulsive characteristics of each passage. The design space and the number of design variables are selected and described accordingly. An actuator disk model is employed to model thrust generation and shaft power from the propulsor. To carry out the optimization of the propulsion/power system on a computationally expensive CFD model, a Kriging method in conjunction with a Genetic Algorithm (GA) is applied. Throughout the design process, the propulsion performances of the sampled propulsion/power system are analyzed and compared to those of a clean flow engine. The performance metrics includes mass flow rate, fan pressure ratio besides the thrust and shaft power. Minimization of total shaft power from the distributed engine is performed at multiple thrust levels. The benefit of boundary layer ingestion propulsion system is quantified via comparison of thrust equivalent, shaft power and mass flow equivalent clean flow engines with CFD based system design.

Hybrid Wingbody↗
NASA NTRS1997

Design Protocols and Analytical Strategies that Incorporate Structural Reliability Models

Ceramic matrix composites (CMC) and intermetallic materials (e.g., single crystal nickel aluminide) are high performance materials that exhibit attractive mechanical, thermal and chemical properties. These materials are critically important in advancing certain performance aspects of gas turbine engines. From an aerospace engineer's perspective the new generation of ceramic composites and intermetallics offers a significant potential for raising the thrust/weight ratio and reducing NO(x) emissions of gas turbine engines. These aspects have increased interest in utilizing these materials in the hot sections of turbine engines. However, as these materials evolve and their performance characteristics improve a persistent need exists for state-of-the-art analytical methods that predict the response of components fabricated from CMC and intermetallic material systems. This need provided the motivation for the technology developed under this research effort. Continuous ceramic fiber composites exhibit an increase in work of fracture, which allows for "graceful" rather than catastrophic failure. When loaded in the fiber direction, these composites retain substantial strength capacity beyond the initiation of transverse matrix cracking despite the fact that neither of its constituents would exhibit such behavior if tested alone. As additional load is applied beyond first matrix cracking, the matrix tends to break in a series of cracks bridged by the ceramic fibers. Any additional load is born increasingly by the fibers until the ultimate strength of the composite is reached. Thus modeling efforts supported under this research effort have focused on predicting this sort of behavior. For single crystal intermetallics the issues that motivated the technology development involved questions relating to material behavior and component design. Thus the research effort supported by this grant had to determine the statistical nature and source of fracture in a high strength, NiAl single crystal turbine blade material; map a simplistic failure strength envelope of the material; develop a statistically based reliability computer algorithm, verify the reliability model and computer algorithm, and model stator vanes for rig tests. Thus establishing design protocols that enable the engineer to analyze and predict the mechanical behavior of ceramic composites and intermetallics would mitigate the prototype (trial and error) approach currently used by the engineering community. The primary objective of the research effort supported by this short term grant is the continued creation of enabling technologies for the macroanalysis of components fabricated from ceramic composites and intermetallic material systems. The creation of enabling technologies aids in shortening the product development cycle of components fabricated from the new high technology materials.

Duffy, Stephen F.↗