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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 NTRS1980

Cold-air investigation of a 4 1/2 stage turbine with stage-loading factor of 4.66 and high specific work output. 2: Stage group performance

The stage group performance of a 4 1/2 stage turbine with an average stage loading factor of 4.66 and high specific work output was determined in cold air at design equivalent speed. The four stage turbine configuration produced design equivalent work output with an efficiency of 0.856; a barely discernible difference from the 0.855 obtained for the complete 4 1/2 stage turbine in a previous investigation. The turbine was designed and the procedure embodied the following design features: (1) controlled vortex flow, (2) tailored radial work distribution, and (3) control of the location of the boundary-layer transition point on the airfoil suction surface. The efficiency forecast for the 4 1/2 stage turbine was 0.886, and the value predicted using a reference method was 0.862. The stage group performance results were used to determine the individual stage efficiencies for the condition at which design 4 1/2 stage work output was obtained. The efficiencies of stages one and four were about 0.020 lower than the predicted value, that of stage two was 0.014 lower, and that of stage three was about equal to the predicted value. Thus all the stages operated reasonably close to their expected performance levels, and the overall (4 1/2 stage) performance was not degraded by any particularly inefficient component.

Whitney, W. J.↗
NASA NTRS1952

Effect of Various Blade Modifications in Performance of a 16-Stage Axial-flow Compressor. IV - Effect on Over-all Performance Characteristics of Decreasing Twelfth through Fifteenth Stage Stator-blade Angles 3 deg and Increasing Stator Angles in the Inlet Stages

The performance of a 16-stage axial-flow compressor, in which two modifications of unloaded inlet stages were combined with loaded exit stages, has been determined. In the first modification the exit stages were loaded by decreasing the twelfth through fifteenth stage stator angles 3 deg. as compared with the blade angles in the original compressor, and the inlet stages were unloaded by increasing the blade angles the following amounts: guide vanes and first-stage stator, 6 deg; second- and third-stage stators, 4 deg.; and fourth-stage stators, 3 deg. The over-all performance of this configuration was compared with that of the compressor with the original blade angles. The peak efficiency was increased at all speeds below design and the weight flow was higher at speeds below 80 percent of design, the same at 80 percent of design, and lower at speeds abovce 80 percent of design. The maximum reduction in weight flow occurred at design speed. The surge limit line was higher at speeds between 75 and 90 percent of design when presented on a pressure ratio against weight flow basis. The second configuration was the same as the first with the exception that the second-, third-, and fourth-stage stator blade angles were the same as in the compressor with the original blade angles. A comparison of the performance of this configuration with that of the compressor with the original blade angles showed the same general trends of changes in performance as the first configuration. Comparisons were made of compressor configurations to show the effects upon the performance of decreased loading in the inlet stages. Below 75 percent of design speed, decreased loading results in increased weight flow and peak efficiency; above 80 percent of design speed, decreased loading in the inlet stages results in decreased weight flow and small changes in peak efficiencies. Between 75 and 90 percent of design the changes in surge weight flow and pressure ratio were such that the surge limit line was raised with decreased loading in the inlet stages when presented as pressure ratio against weight flow.

Hatch, James E.↗
NASA NTRS1976

Single-stage experimental evaluation of low aspect ratio, highly loaded blading for compressors. Part 9: Stage F and stage G, volume 1

Two single-stage, 0.77 hub/tip ratio axial-flow compressors were tested to evaluate the effectiveness of low aspect ratio blading as a means of obtaining higher stage loadings. One compressor, designated Stage F, was comprised of circular arc blading with an aspect ratio of 0.9 for both the rotor and stator. This compressor was tested with uniform inlet flow, hub radial, tip radial, and 180 deg arc circumferential inlet distortion. The second compressor, designated Stage G, was comprised of multiple circular arc blading with an aspect ratio of 1.0 for both the rotor and stator. This compressor was tested with uniform inlet flow only. Design rotor tip speeds for Rotor F and Rotor G were 285 m/sec (934 ft/sec) and 327 m/sec (1,074 ft/sec) respectively. Both stages operated at high loading levels with adequate efficiency and operating range. The peak efficiencies and corresponding average stage diffusion factors for Stages F and G at design rotor speed were 86.4% and 84.1% and 0.59 and 0.55 respectively. The surge margin at peak efficiency for Stage F was 12.6% and the corresponding value for Stage G was 16.5%. Both stages experienced a loss in efficiency with increasing rotor speed; however, the multiple circular arc rotor delayed the characteristic loss in efficiency within increasing Mach number to higher Mach number.

Cheatham, J. G.↗
NASA NTRS1991

A comparison of single-stage and two-stage airbreathing launch vehicles at high staging Mach numbers

A predominantly airbreathing SSTO vehicle is compared with three different two stages to orbit (TSTO) configurations which stage at Mach numbers 10, 12, and 14. The technologies utilized on each of the vehicles are consistent with a successful NASP technology development program in the areas of structures, propulsion systems, and subsystems. To determine the effect of subsystem and structural technologies on the relative weights of the SSTO and TSTO configurations, weight sensitivity trades are presented.

Stanley, Douglas O.↗
NASA NTRS1956

Experimental Investigation of a Five-stage Axial-flow Research Compressor with Transonic Rotors in All Stages III : Interstage Data and Individual Stage Performance Characteristics

For use in computing the detailed performance parameters of individual blade rows of a multistage compressor, radial distributions of total pressure, total temperature, static pressure and air-flow angle are tabulated. The data cover a range of air flow from choke to the approximate surge limit for equivalent rotor speeds from 70 to 100 percent of design. Equivalent stage performance curves determined from the radial survey measurements indicate both the individual performance and the matching of the stages at the speeds investigated.

Sandercock, Donald M↗
NASA NTRS1991

Technology and staging effects on two-stage-to-orbit systems

Horizontal takeoff and landing two-stage systems with an airbreathing first stage and rocket second stage are evaluated for staging Mach numbers that range from 5 to 14. All systems are evaluated with advanced technologies being developed in the NASP Program and sized to the same mission requirements. With these advanced technologies, the two-stage systems are heavier than the single stage. The weights of the two-stage systems are closely related to staging. Using a rocket on the first stage to accelerate from the turboramjet limit of Mach 6 to Mach 10 signiificantly decreases dry weight as compared to the Mach 6-staged system. The optimum dry weight staging Mach number for the scramjet two-stage system is Mach 12. At a 40 percent weight growth (current technology level), the scramjet two-stage systems are half the weight and less sensitive to weight changes than the single stage, but still require substantial technology development in the areas of inlets, nozzles, ramjet propulsion, active cooling, and high-temperature structures.

Wilhite, Alan W.↗
NASA NTRS2015

Exprimental Results of the First Two Stages of an Advanced Transonic Core Compressor Under Isolated and Multi-Stage Conditions.

NASA's Environmentally Responsible Aviation (ERA) Program calls for investigation of the technology barriers associated with improved fuel efficiency for large gas turbine engines. Under ERA, the highly loaded core compressor technology program attempts to realize the fuel burn reduction goal by increasing overall pressure ratio of the compressor to increase thermal efficiency of the engine. Study engines with overall pressure ratio of 60 to 70 are now being investigated. This means that the high pressure compressor would have to almost double in pressure ratio while keeping a high level of efficiency. NASA and GE teamed to address this challenge by testing the first two stages of an advanced GE compressor designed to meet the requirements of a very high pressure ratio core compressor. Previous test experience of a compressor which included these front two stages indicated a performance deficit relative to design intent. Therefore, the current rig was designed to run in 1-stage and 2-stage configurations in two separate tests to assess whether the bow shock of the second rotor interacting with the upstream stage contributed to the unpredicted performance deficit, or if the culprit was due to interaction of rotor 1 and stator 1. Thus, the goal was to fully understand the stage 1 performance under isolated and multi-stage conditions, and additionally to provide a detailed aerodynamic data set for CFD validation. Full use was made of steady and unsteady measurement methods to understand fluid dynamics loss source mechanisms due to rotor shock interaction and endwall losses. This paper will present the description of the compressor test article and its measured performance and operability, for both the single stage and two stage configurations. We focus the paper on measurements at 97% corrected speed with design intent vane setting angles.

transonic compressors↗
NASA NTRS1993

Analysis of the staging maneuver and booster glideback guidance for a two-stage, winged, fully reusable launch vehicle

One of the promising launch concepts that could replace the current space shuttle launch system is a two-stage, winged, vertical-takeoff, fully reusable launch vehicle. During the boost phase of ascent, the booster provides propellant for the orbiter engines through a cross-feed system. When the vehicle reaches a Mach number of 3, the booster propellants are depleted and the booster is staged and glides unpowered to a horizontal landing at a launch site runway. Two major design issues for this class of vehicle are the staging maneuver and the booster glideback. For the staging maneuver analysis, a technique was developed that provides for a successful separation of the booster from the orbiter over a wide range of staging angles of attack. A longitudinal flight control system was developed for control of the booster during the staging maneuver. For the booster glide back analysis, a guidance algorithm was developed that successfully guides the booster from the completion of the staging maneuver to a launch site runway while encountering many off-nominal atmospheric, aerodynamic, and staging conditions.

Naftel, J. Christopher↗
NASA NTRS2015

Experimental Results of the First Two Stages of an Advanced Transonic Core Compressor Under Isolated and Multi-Stage Conditions

NASA's Environmentally Responsible Aviation (ERA) Program calls for investigation of the technology barriers associated with improved fuel efficiency of large gas turbine engines. Under ERA the task for a High Pressure Ratio Core Technology program calls for a higher overall pressure ratio of 60 to 70. This mean that the HPC would have to almost double in pressure ratio and keep its high level of efficiency. The challenge is how to match the corrected mass flow rate of the front two supersonic high reaction and high corrected tip speed stages with a total pressure ratio of 3.5. NASA and GE teamed to address this challenge by using the initial geometry of an advanced GE compressor design to meet the requirements of the first 2 stages of the very high pressure ratio core compressor. The rig was configured to run as a 2 stage machine, with Strut and IGV, Rotor 1 and Stator 1 run as independent tests which were then followed by adding the second stage. The goal is to fully understand the stage performances under isolated and multi-stage conditions and fully understand any differences and provide a detailed aerodynamic data set for CFD validation. Full use was made of steady and unsteady measurement methods to isolate fluid dynamics loss source mechanisms due to interaction and endwalls. The paper will present the description of the compressor test article, its predicted performance and operability, and the experimental results for both the single stage and two stage configurations. We focus the detailed measurements on 97 and 100 of design speed at 3 vane setting angles.

turbomachinery↗
NASA NTRS1954

Experimental Investigation of Stage Performance of J71 Three-stage Turbine

An experimental investigation of the stage performance of the J71 three-stage turbine at design speed and pressure ratio was conducted as part of a program involving research on the design and operational characteristics of high-work-output multistage axial-flow turbines. The investigation indicated that the over-all work output of the turbine based on temperature measurements at design speed and pressure ratio was 95 percent of the design value, with the greatest loss in work output and efficiency occurring in the first stage. Choking in the second-stage stator over most of the pressure ratios investigated prevented the attainment of design first-stage work output and also prevented any increase in work output of the first stage with increasing over-all pressure ratio at design speed. High losses in the first stage were a contributing factor to this choking condition.

Forrette, Obert E↗
NASA NTRS1951

Effect of Angle of Incidence of Second-Stage Vane Assembly on Third-Stage Compressor-Blade Vibration and Engine Performance

Strain-gages were used to measure blade vibrations causing failures in the third stage of a production 11-stage axial-flow compressor. After the serious third-stage vibration was detected, a series of investigations were conducted with second-stage vane assemblies of varying angles of incidence. Curves presented herein show the effect of varying the angle of incidence of second-stage vane assembly on third-stage rotor-blade vibration amplitude and engine performance. A minimum vibration amplitude was obtained without greatly affecting the engine performance with a second-stage vane assembly of 9deg. greater angle of incidence than the assembly normally furnished with the engine.

Meyer, Andre J., Jr.↗
NASA NTRS2015

Stage-by-Stage and Parallel Flow Path Compressor Modeling for a Variable Cycle Engine

This paper covers the development of stage-by-stage and parallel flow path compressor modeling approaches for a Variable Cycle Engine. The stage-by-stage compressor modeling approach is an extension of a technique for lumped volume dynamics and performance characteristic modeling. It was developed to improve the accuracy of axial compressor dynamics over lumped volume dynamics modeling. The stage-by-stage compressor model presented here is formulated into a parallel flow path model that includes both axial and rotational dynamics. This is done to enable the study of compressor and propulsion system dynamic performance under flow distortion conditions. The approaches utilized here are generic and should be applicable for the modeling of any axial flow compressor design.

Parallel Flow Path modeling↗
NASA NTRS2015

Stage-by-Stage and Parallel Flow Path Compressor Modeling for a Variable Cycle Engine, NASA Advanced Air Vehicles Program - Commercial Supersonic Technology Project - AeroServoElasticity

This paper covers the development of stage-by-stage and parallel flow path compressor modeling approaches for a Variable Cycle Engine. The stage-by-stage compressor modeling approach is an extension of a technique for lumped volume dynamics and performance characteristic modeling. It was developed to improve the accuracy of axial compressor dynamics over lumped volume dynamics modeling. The stage-by-stage compressor model presented here is formulated into a parallel flow path model that includes both axial and rotational dynamics. This is done to enable the study of compressor and propulsion system dynamic performance under flow distortion conditions. The approaches utilized here are generic and should be applicable for the modeling of any axial flow compressor design accurate time domain simulations. The objective of this work is as follows. Given the parameters describing the conditions of atmospheric disturbances, and utilizing the derived formulations, directly compute the transfer function poles and zeros describing these disturbances for acoustic velocity, temperature, pressure, and density. Time domain simulations of representative atmospheric turbulence can then be developed by utilizing these computed transfer functions together with the disturbance frequencies of interest.

Compressor Modeling Modeling↗
NASA NTRS1981

Cold-air investigation of first stage of 4-1/2-stage, fan drive turbine with average stage-loading factor of 4.66

The design procedure and the development of the blading geometry for the 4 and 1/2 stage turbine are discussed. Results obtained with the first stage, operated as a single stage turbine, are presented. A free vortex design meets the design requirements without incurring problems such as excessive turning, negative reaction, or high Mach number. Cold air tests of the single stage turbine showed that the turbine developed design work (stage loading factor of 5.26) at an efficiency of 0.86, which was the efficiency predicted by a reference method. The mass flow at this condition was 0.88, which occurred at design speed and a pressure ratio of 1.407, corresponding to a stage loading factor of 4.35. The efficiency at this condition was 0.003 higher than that predicted by the reference method.

Whitney, W. J.↗