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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

An improved computer model for prediction of axial gas turbine performance losses

The calculation model performs a rapid preliminary pitchline optimization of axial gas turbine annular flowpath geometry, as well as an initial estimate of blade profile shapes, given only a minimum of thermodynamic cycle requirements. No geometric parameters need be specified. The following preliminary design data are determined: (1) the optimum flowpath geometry, within mechanical stress limits; (2) initial estimates of cascade blade shapes; and (3) predictions of expected turbine performance. The model uses an inverse calculation technique whereby blade profiles are generated by designing channels to yield a specified velocity distribution on the two walls. Velocity distributions are then used to calculate the cascade loss parameters. Calculated blade shapes are used primarily to determine whether the assumed velocity loadings are physically realistic. Model verification is accomplished by comparison of predicted turbine geometry and performance with an array of seven NASA single-stage axial gas turbine configurations.

Jenkins, R. M.↗
NASA NTRS1980

Off-design performance loss model for radial turbines with pivoting, variable-area stators

An off-design performance loss model was developed for variable stator (pivoted vane), radial turbines through analytical modeling and experimental data analysis. Stator loss is determined by a viscous loss model; stator vane end-clearance leakage effects are determined by a clearance flow model. Rotor loss coefficient were obtained by analyzing the experimental data from a turbine rotor previously tested with six stators having throat areas from 20 to 144 percent of design area and were correlated with stator-to-rotor throat area ratio. An incidence loss model was selected to obtain best agreement with experimental results. Predicted turbine performance is compared with experimental results for the design rotor as well as with results for extended and cutback versions of the rotor. Sample calculations were made to show the effects of stator vane end-clearance leakage.

Meitner, P. L.↗
NASA NTRS1985

Chemical kinetic performance losses for a hydrogen laser thermal thruster

Projected requirements for efficient, economical, orbit-raising propulsion systems have generated investigations into several potentially high specific impulse, moderate thrust, advanced systems. One of these systems, laser thermal propulsion, utilizes a high temperature plasma as the enthalpy source. The plasma is sustained by a focused laser beam which maintains the plasma temperature at levels near 20,000 K. Since such temperature levels lead to total dissociation and high ionization, the plasma thruster system potentially has a high specific impulse decrement due to recombination losses. The nozzle flow is expected to be sufficiently nonequilibrium to warrant concern over the achievable specific impluse. This investigation was an attempt at evaluation of those losses. The One-Dimensional Kinetics (ODK) option of the Two-Dimensional Kinetics (TDK) Computer Program was used with a chemical kinetics rate set obtained from available literature to determine the chemical kinetic energy losses for typical plasma thruster conditions. The rates were varied about the nominal accepted values to band the possible losses. Kinetic losses were shown to be highly significant for a laser thermal thruster using hydrogen. A 30 percent reduction in specific impulse is possible simply due to the inability to completely extract the molecular recombination energy.

Mccay, T. D.↗
NASA NTRS1980

Loss model for off-design performance analysis of radial turbines with pivoting-vane, variable-area stators

An off-design performance loss model is developed for variable-area (pivoted vane) radial turbines. The variation in stator loss with stator area is determined by a viscous loss model while the variation in rotor loss due to stator area variation (for no stator end-clearance gap) is determined through analytical matching of experimental data. An incidence loss model is also based on matching of the experimental data. A stator vane end-clearance leakage model is developed and sample calculations are made to show the predicted effects of stator vane end-clearance leakage on performance.

Meitner, P. L.↗
NASA NTRS1980

Loss model for off-design performance analysis of radial turbines with pivoting-vane, variable-area stators

An off-design performance loss model for a radial turbine with pivoting, variable-area stators is developed through a combination of analytical modeling and experimental data analysis. A viscous loss model is used for the variation in stator loss with setting angle, and stator vane end-clearance leakage effects are predicted by a clearance flow model. The variation of rotor loss coefficient with stator setting angle is obtained by means of an analytical matching of experimental data for a rotor that was tested with six stators, having throat areas from 20 to 144% of the design area. An incidence loss model is selected to obtain best agreement with experimental data. The stator vane end-clearance leakage model predicts increasing mass flow and decreasing efficiency as a result of end-clearances, with changes becoming significantly larger with decreasing stator area.

Meitner, P. L.↗
NASA NTRS2013

The Aerodynamic Performance of an Over-the-Rotor Liner With Circumferential Grooves on a High Bypass Ratio Turbofan Rotor

While liners have been utilized throughout turbofan ducts to attenuate fan noise, additional attenuation is obtainable by placing an acoustic liner over-the-rotor. Previous experiments have shown significant fan performance losses when acoustic liners are installed over-the-rotor. The fan blades induce an oscillating flow in the acoustic liners which results in a performance loss near the blade tip. An over-the-rotor liner was designed with circumferential grooves between the fan blade tips and the acoustic liner to reduce the oscillating flow in the acoustic liner. An experiment was conducted in the W-8 Single-Stage Axial Compressor Facility at NASA Glenn Research Center on a 1.5 pressure ratio fan to evaluate the impact of this over-the-rotor treatment design on fan aerodynamic performance. The addition of a circumferentially grooved over-the-rotor design between the fan blades and the acoustic liner reduced the performance loss, in terms of fan adiabatic efficiency, to less than 1 percent which is within the repeatability of this experiment.

Bozak, Richard F.↗
NASA NTRS2013

The Aerodynamic Performance of an Over-The-Rotor Liner with Circumferential Grooves on a High Bypass Ratio Turbofan Rotor

While liners have been utilized throughout turbofan ducts to attenuate fan noise, additional attenuation is obtainable by placing an acoustic liner over-the-rotor. Previous experiments have shown significant fan performance losses when acoustic liners are installed over-the-rotor. The fan blades induce an oscillating flow in the acoustic liners which results in a performance loss near the blade tip. An over-the-rotor liner was designed with circumferential grooves between the fan blade tips and the acoustic liner to reduce the oscillating flow in the acoustic liner. An experiment was conducted in the W-8 Single-Stage Axial Compressor Facility at NASA Glenn Research Center on a 1.5 pressure ratio fan to evaluate the impact of this over-the-rotor treatment design on fan aerodynamic performance. The addition of a circumferentially grooved over-the-rotor design between the fan blades and the acoustic liner reduced the performance loss, in terms of fan adiabatic efficiency, to less than 1% which is within the repeatability of this experiment.

Bozak, Rick↗
NASA NTRS1947

Effect of Air-Flow Distribution and Total-Pressure Loss on Performance of One-Sixth Segment of Turbojet Combuster

An investigation has been conducted on a one-sixth segment of an annular turbojet combustor to determine the effects of modification in air-flow distribution and total-pressure loss on the performance of the segment. The performance features investigated during this series of determinations were the altitude operational limits and the temperature-rise efficiency. Altitude operational limits of the combustor segment, for the 19XB engine using the original combustor-basket design were approximately 38,000 feet at 17,000 rpm and 26,000 feet at 10,000 rpm. The altitude operational limits were approximately 50,000 feet at 17,000 rpm and 38,000 feet at 10,000 rpm for a combustor-basket design in which the air-passage area in the basket was redistributed so as to admit gradually no more than 20 percent of the air along the first half of the basket. In this case the total pressure loss through the combustor segment was not appreciably changed from the total-pressure loss for the original combustor basket design. Altitude operational limits of the combustor segment for the 19XB engine were above 52,000 feet at 17,000 rpm and were approximately 23,000 feet at 10,000 rpm for a combustor-basket design in which the distribution of the air-passage area in the basket was that of the original design but where the total-pressure loss was increased to 19 times the inlet reference kinetic pressure at an inlet-to-outlet density ratio of 2.4. The total-pressure loss for the original design was 14 times the inlet kinetic reference pressure at an inlet-to-outlet density ratio of 2.4.

Hill, Francis U.↗
NASA NTRS1980

Determination of efficiencies, loss mechanisms, and performance degradation factors in chopper controlled dc vehical motors. Section 1: Test program results and recommendations

The conventional series motor model is discussed as well as procedures for obtaining, by test, the parameters necessary for calculating performance and losses. The calculated results for operation from ripple free DC are compared with observed test results, indicating approximately 5% or less error. Experimental data indicating the influence of brush shift and chopper frequency are also presented. Both factors have a significant effect on the speed and torque relationships. The losses and loss mechanisms present in a DC series motor are examined and an attempt is made to evaluate the added losses due to harmonic currents and fluxes. Findings with respect to these losses is summarized.

Hamilton, H. B.↗
NASA NTRS1978

Short-term performance deterioration in JT9D-7A(SP) engine 695743

The level of performance deterioration that occurred during early service was studied and also the contribution of each engine module to the overall level of engine performance deterioration. To assist in the distribution of the performance losses to the module level, testing with expanded experimental instrumentation was accomplished to permit direct measurement of module performance. An analytical teardown of the engine was accomplished to permit assignment of module performance losses to individual mechanical damage mechanisms including blade and seal wear, roughness, and thermal distortion of turbine parts. The results of the testing show that the engine deteriorated 1.5 percent in thrust specific fuel consumption from its performance when new. Initial testing, which included water washing, showed that 0.2 percent in performance deterioration was due to surface contamination (dirt) and 0.1 percent was due to drift of the engine vane control system, leaving 1.2 percent in basic TSFC deterioration. This 1.2 percent TSFC loss was distributed among the engine modules with 0.6 percent caused by clearance changes, 0.4 percent loss due to thermal distortion, and 0.2 percent loss due to increased roughness of the fan and low-pressure compressor airfoils.

Bouchard, R. J.↗
NASA NTRS2019

A Performance Analysis of Folding Conformal Propeller Blade Designs

NASA’s X-57 Maxwell flight demonstrator has a high-lift system that includes 12 fixed- pitch high-lift propellers located upstream of the wing leading edge for lift augmentation at low speeds. These high-lift propellers are only required at low speeds, and to reduce drag, the propeller blades are folded conformally along the nacelles at other operating conditions. The method of designing the high-lift blades permits several variations of blade cross-section placement along the nacelle surface and a comparative performance analysis was needed to determine if any particular design showed significant benefits. We analyzed the performance of three conformal high-lift propeller designs and compared them to that of a non-conformal baseline propeller to establish both the benefit of stowable blades and the value of each variation. In this study, we first performed a drag analysis of each design in the stowed configuration at the X-57 cruise speed and altitude to determine the drag benefits of each conforming method. Then, among blade designs we compared the thrust, power, and lift for a given input shaft speed to establish any performance losses from the baseline. This analysis shows that the conformal blade designs do not have any appreciable performance losses compared to the baseline blades. Moreover, although the drag in the cruise condition is significantly less than for the non-folding baseline, the drag benefits of each conforming blade approach are similar and the value of each approach largely depends on the ease of integration into the nacelle. This paper presents the results of these studies and discusses the benefits and drawbacks of implementing the conformal blade designs. Specifically, we demonstrate that folding, conformal propeller blades contribute significantly less to cruise drag when compared to windmilling, with an increase relative to a. We also show a less than 1% difference in performance formal, folding propellers and the non-conforming baseline propeller.

Litherland, Brandon L.↗
NASA NTRS1979

Causes of high pressure compressor deterioration in service

The high mechanical reliability and low deterioration rate of the JT9D high-pressure compressor results in long utilization without exceeding engine operational limits. The increasing cost and decreasing supply of fuel have focused attention on the fuel burned implications of such high time use without refurbishment. The paper presents the results of JT9D high pressure compressor studies. The mechanical deterioration of the JT9D high-pressure compressor gaspath parts versus increasing service usage, documented from inspection of service parts, is presented and discussed including changes in airfoil roughness, blade length, airfoil contour and outer air seal trench characteristics. An estimate of the performance loss versus usage is related to each type of damage. The combined estimated high-pressure compressor performance loss for all mechanisms determined from part inspection is compared to historical engine test data to establish the validity of the predicted loss levels. The effect of cold section refurbishment on engine fuel consumption recovery and the results of an optimization study to determine the appropriate interval for high pressure compressor refurbishment are also reported.

Richardson, J. H.↗
NASA NTRS2005

Model-Scale Aerodynamic Performance Testing of Proposed Modifications to the NASA Langley Low Speed Aeroacoustic Wind Tunnel

Tests were performed on a 1/20th-scale model of the Low Speed Aeroacoustic Wind Tunnel to determine the performance effects of insertion of acoustic baffles in the tunnel inlet, replacement of the existing collector with a new collector design in the open jet test section, and addition of flow splitters to the acoustic baffle section downstream of the test section. As expected, the inlet baffles caused a reduction in facility performance. About half of the performance loss was recovered by addition the flow splitters to the downstream baffles. All collectors tested reduced facility performance. However, test chamber recirculation flow was reduced by the new collector designs and shielding of some of the microphones was reduced owing to the smaller size of the new collector. Overall performance loss in the facility is expected to be a 5 percent top flow speed reduction, but the facility will meet OSHA limits for external noise levels and recirculation in the test section will be reduced.

Booth, Earl R., Jr.↗
NASA NTRS1948

Analysis of Effects of Inlet Pressure Losses on Performance of Axial-Flow Type Turbojet Engine

The experimentally determined performance characteristics of an axial-flow turbojet engine have been used to estimate the effects of inlet total-pressure losses on net thrust and specific fuel consumption at a constant engine speed. At low altitudes and flight Mach numbers, inlet pressure losses cause an increase in engine discharge temperature and it is possible that the maximum allowable turbine temperature maybe exceeded. An inlet absolute total-pressure loss of 10 percent will result in a thrust loss of 14 percent and a 15-percent increase in specific fuel consumption based on net thrust. At high altitudes and flight Mach numbers, choking conditions exist in the exhaust nozzle and the inlet pressure losses do not affect the discharge temperatures. Under these conditions, a 10-percent loss in inlet absolute total pressure produces a 22-percent loss in net thrust and a 16-percent increase in specific fuel consumption. If the exhaust-nozzle-outlet area is adjusted to compensate for the effect of inlet losses on discharge temperature in the nonchoking cases (low altitude and Mach numbers), the thrust and fuel consumption will be changed in a manner similar to the results obtained in the choking cases.

Sanders, Newell D↗
NASA NTRS2019

Surface Flow Visualization of the High-Lift Common Research Model

A 10% scale version of the High-Lift Common Research Model (CRM-HL) was tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22) in support of the NASA Advanced Air Transport Technology (AATT) Project. The CRM-HL experiment included various configurations such as conventional and simple-hinged flaps, with and without engine nacelle/pylon, with and without nacelle chine, different Active Flow Control (AFC) methods (sweeping jets, alternating pulsed jets, and preconditioned boundary layer blowing), and their various parameters. This particular study is focused on the surface flow visualization of the conventional CRM-HL model at landing configuration. The conventional CRM-HL model with the single-slotted Fowler flap system serves as a baseline for the AFC-enabled simplified high-lift configuration as well as a high-lift technology development platform due to its publicly open geometry. Surface flow visualizations were performed using fluorescent minitufts, which were found to be nonintrusive to the aerodynamic performance. Tuft flow visualizations are supplemented with the relevant pressure and force measurements in order to understand the flow characteristics developed on the conventional CRM- HL model. In addition, three dimensional, unsteady, compressible Computational Fluid Dynamic (CFD) simulations were performed for selective cases. The surface streamlines and transverse velocity fluctuations obtained by the CFD simulations are qualitatively compared to the tuft direction and tuft unsteadiness, respectively. Force measurements of the CRM-HL model show performance degradation at higher angles of attack. Surface flow visualizations revealed the performance loss due to the nacelle/pylon wake that grows with angle of attack and eventually promotes flow separation over the inboard wing. This performance loss was successfully recovered by placing a chine on the engine nacelle.

Koklu, Mehti↗
NASA NTRS2005

Experimental Investigation of Ice Accretion Effects on a Swept Wing

An experimental investigation was conducted to study the effects of 2-, 5-, 10-, and 22.5-min ice accretions on the aerodynamic performance of a swept finite wing. The ice shapes tested included castings of ice accretions obtained from icing tests at the NASA Glenn Icing Research Tunnel (IRT) and simulated ice shapes obtained with the LEWICE 2.0 ice accretion code. The conditions used for the icing tests were selected to provide five glaze ice shapes with complete and incomplete scallop features and a small rime ice shape. The LEWICE ice shapes were defined for the same conditions as those used in the icing tests. All aerodynamic performance tests were conducted in the 7- x 10-ft Low-Speed Wind Tunnel Facility at Wichita State University. Six component force and moment measurements, aileron hinge moments, and surface pressures were obtained for a Reynolds number of 1.8 million based on mean aerodynamic chord and aileron deflections in the range of -15o to 20o. Tests were performed with the clean wing, six IRT ice shape castings, seven smooth LEWICE ice shapes, and seven rough LEWICE ice shapes. Roughness for the LEWICE ice shapes was simulated with 36-size grit. The experiments conducted showed that the glaze ice castings reduced the maximum lift coefficient of the clean wing by 11.5% to 93.6%, while the 5-min rime ice casting increased maximum lift by 3.4%. Minimum iced wing drag was 133% to 3533% greater with respect to the clean case. The drag of the iced wing near the clean wing stall angle of attack was 17% to 104% higher than that of the clean case. In general, the aileron remained effective in changing the lift of the clean and iced wings for all angles of attack and aileron deflections tested. Aileron hinge moments for the iced wing cases remained within the maximum and minimum limits defined by the clean wing hinge moments. Tests conducted with the LEWICE ice shapes showed that in general the trends in aerodynamic performance degradation of the wing with the simulated ice shapes were similar to those obtained with the IRT ice shape castings. However, in most cases, the ice castings resulted in greater aerodynamic performance losses than those obtained with the LEWICE ice shapes. For the majority of the LEWICE ice shapes, the addition of 36-size grit roughness to the smooth ice shapes increased aerodynamic performance losses.

Ice shape↗