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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 307 records · Page 17

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NASA NTRS1975

Preliminary study of advanced turbofans for low energy consumption

This analysis determines the effect of higher overall engine pressure ratios (OPR's), bypass ratios (BPR's), and turbine rotor-inlet temperature on a Mach-0.85 transport having a range of 5556 km (3000 nmi) and carrying a payload of 18144 kg (40,000 lbs-200 passengers). Sideline noises (jet plus fan) of between 91 and 106 EPNdB (FAR36) are considered. Takeoff gross weight (TOGW), fuel consumption (kg/pass. km) and direct operating cost (DOC) are used at the figures of merit. Based on predicted 1985 levels of engine technology and a noise goal of 96 EPNdB, the higher-OPR engine results in an airplane that is 18 percent lighter in terms of TOGW, uses 22.3 percent less fuel, and has a 14.7 percent lower DOC than a comparable airplane powered by a current turbofan. Cooling the compressor bleed air and lowering the cruise Mach number appear attractive in terms of further improving the figures of merit.

Knip, G.↗

NASA NTRS1975

Experimental quiet engine program

Full-scale low-tip-speed fans, a full-scale high-tip-speed fan, scale model versions of fans, and two full-scale high-bypass-ratio turbofan engines, were designed, fabricated, tested, and evaluated. Turbine noise suppression was investigated. Preliminary design studies of flight propulsion system concepts were used in application studies to determine acoustic-economic tradeoffs. Salient results are as follows: tradeoff evaluation of fan tip speed and blade loading; systematic data on source noise characteristics and suppression effectiveness; documentation of high- and low-fan-speed aerodynamic and acoustic technology; aerodynamic and acoustic evaluation of acoustic treatment configurations, casing tip bleed, serrated and variable pitch rotor blades, leaned outlet guide vanes, slotted tip casings, rotor blade shape modifications, and inlet noise suppression; systematic evaluation of aerodynamic and acoustic effects; flyover noise projections of engine test data; turbine noise suppression technology development; and tradeoff evaluation of preliminary design high-fan-speed and low-fan-speed flight engines.

Cornell, W. G.↗

NASA NTRS1974

Engineering simulation evaluation of the two-segment noise abatement approach in the Douglas DC-8-61

The development and operational evaluation are discussed of a noise abatement two-segment approach which is safe and operationally acceptable for routine use in air carrier service. An engineering simulation evaluation was conducted; it was found that the entry speed and initial configuration are more important in the DC-8-61 than in the 727 because of the DC-8's aerodynamic cleanliness and drag programming constraints. It was also found that the altitude required for the DC-8 upper and lower transitions is less than the 727 transitions principally due to the pitching moment from the underslung DC-8 engines as compared to the 727 engines which produce no pitching moment with a change in thrust. The 727 approach was a reduced flaps (30 degrees) procedure with a 10-knot airspeed bleed in the lower transition; the DC-8 approach is a full flaps (50 degrees) procedure with constant airspeed established as soon as possible after upper segment transition.

Nylen, W. E.↗

NASA NTRS1975

An experimental and analytical investigation of axisymmetric diffusers

A finite difference computer program for turbulent compressible flow was used to establish the performance of several diffuser shapes for experimental testing. The diffusers were designed to have a linear change in Mach number, a linear change in pressure, or a curvature fitted by a quadratic equation. Testing was performed with M = 0.1 to 0.9 with and without boundary layer bleed. Above M = 0.6, data were obtained with a normal shock upstream of the diffuser entrance. Peak static pressure recovery occurred with a diffuser inlet M0.75. The quadratic diffuser yielded the highest total pressure recovery.

Povinelli, L. A.↗

NASA NTRS1975

Effect of wall suction on performance of a short annular diffuser at inlet Mach numbers up to 0.5

A short annular diffuser equipped with wall bleed (suction)capability was evaluated at inlet Mach numbers of 0.186 to 0.5. The diffuser had an area ratio of 4.0 and a length-to-inlet height ratio of 1.6. Test results show that the exit velocity profiles, typical of annular jet flow without suction, could be considerably flattened by application of wall suction. This improved performance was also reflected in diffuser effectiveness (static-pressure recovery) and total-pressure loss results. At the inlet Mach number of 0.5 diffuser static-pressure recovery is equal to or better than at lower inlet Mach numbers for comparable suction rates.

Juhasz, A. J.↗

NASA NTRS1975

Experimental investigation of a Mach 6 fixed-geometry inlet featuring a swept external-internal compression flow field

An investigation of a fixed-geometry, swept external-internal compression inlet was conducted at a Mach number of 6.0 and a test-section Reynolds number of 1.55 x 10 to the 7th power per meter. The test conditions was constant for all runs with stagnation pressure and temperature at 20 atmospheres and 500 K, respectively. Tests were made at angles of attack of -5 deg, 0 deg, 3 deg, and 5 deg. Measurements consisted of pitot- and static-pressure surveys in inlet throat, wall static pressures, and surface temperatures. Boundary-layer bleed was provided on the centerbody and on the cowl internal surface. The inlet performance was consistently high over the range of the angle of attack tested, with an overall average total pressure recovery of 78 percent and corresponding adiabatic kinetic-energy efficiency of 99 percent. The inlet throat flow distribution was uniform and the Mach number and pressure level were of the correct magnitude for efficient combustor design. The utilization of a swept compression field to meet the starting requirements of a fixed-geometry inlet produced neither flow instability nor a tendency to unstart.

Torrence, M. G.↗

NASA NTRS1974

Test techniques for obtaining off-nominal compressor data during engine tests

Several unique techniques and related devices are in use at the Lewis Research Center for off-design testing of fan and compressor sections in full-scale jet engines. The devices presented not only permit a wide range of experimental conditions but also minimize downtime for hardware changes. The techniques involve use of such devices as inlet pressure distortion jets, a hydrogen burner for inlet temperature distortions, fan back pressure jets to simulate a variable area nozzle, and either an inflow-outflow bleed system or a fuel spurt system to alter compressor discharge pressure.

Biesiadny, T. J.↗

NASA NTRS1975

The Mach disc in truncated plug nozzle flows

The first shock reflection within truncated plug nozzle propulsive jets is investigated experimentally using a free jet blow-down facility. The locations and sizes of these reflections in axisymmetric plug nozzles, with cylindrical shrouds and conical truncated plugs, are presented. Data are presented for a range of ambient to nozzle total pressure ratios. The effects of Mach disk location and size resulting from changing plug length and from varying amounts of base bleed are documented.

Giel, T. V., Jr.↗

NASA NTRS1975

An experimental and analytical investigation of axisymmetric diffusers

A finite difference computer program for turbulent compressible flow was used to establish the performance of several diffuser shapes for experimental testing. The diffusers were designed to have a linear change in Mach number, a linear change in pressure or a curvature fitted by a quadratic equation. Testing was performed with M = 0.1 to 0.9 with and without boundary layer bleed. Above M = 0.6, data were obtained with a normal shock upstream of the diffuser entrance. Peak static pressure recovery occurred with a diffuser inlet M-0.75. The quadratic diffuser yielded the highest total pressure recovery.

Povinelli, L. A.↗

NASA NTRS1976

Heat-Transfer Characteristics of Partially Film Cooled Plug Nozzle on a J-85 Afterburning Turbojet Engine

Plug nozzle film cooling data were obtained downstream of a slot located at 42 percent of the total plug length on a J-85 engine. Film cooling reduced the aft end wall temperature as much as 150 K, reduced total pressure loss in the upstream convection cooling passages by 50 percent, and reduced estimated compressor bleed flow requirement by 14 percent compared to an all convectively cooled nozzle. Shock waves along the plug surface strongly influenced temperature distributions on both convection and film cooled portions. The effect was most severe at nozzle pressure ratios below 10 where adverse pressure gradients were most severe.

Nosek, S. M.↗

NASA NTRS1972

Feasibility study of inlet shock stability system of YF-12

The feasibility of self actuating bleed valves as a shock stabilization system in the inlet of the YF-12 is considered for vortex valves, slide valves, and poppet valves. Analytical estimation of valve performance indicates that only the slide and poppet valves located in the inlet cowl can meet the desired steady state stabilizing flows, and of the two the poppet valve is substantially faster in response to dynamic disturbances. The poppet valve is, therefore, selected as the best shock stability system for the YF-12 inlet.

Blausey, G. C.↗

NASA NTRS1976

Dynamic blade row compression component model for stability studies

This paper describes a generalized dynamic model which has been developed for use in compression component aerodynamic stability studies. The model is a one-dimensional, pitch-line, blade row, lumped volume system. Arbitrary placement of blade free volumes upstream, within, and downstream of the compression component as well as the removal of bleed flow from the exit of any rotor or stator are model options. The model has been applied to a two-stage fan and an eight-stage compressor. The clean inlet pressure ratio/flow maps and the surge line have been reproduced, thereby validating the capability of the dynamic model to reproduce the steady-flow characteristics of the compression component. A method for determining the onset of an aerodynamic instability which is associated with surge is described. Sinusoidally time-varying inlet and exit boundary conditions have been applied to the eight stage compressor as examples of the manner in which this model may be used for stability studies.

Tesch, W. A.↗

NASA NTRS1976

Advanced porous transonic wind-tunnel nozzles

Advances in the design and control of the porous bleed flow distribution along the nozzle walls solves the problem of nonuniform flow at off-design conditions. This is accomplished in a two dimensional nozzle with porous parallel sidewalls backed with a single plenum chamber and employing a sliding compartment wall or backed with multiple plenum chambers within which the pressure can be controlled.

Sorensen, N. E.↗

NASA NTRS1976

A simplified analysis of propulsion installation losses for computerized aircraft design

A simplified method is presented for computing the installation losses of aircraft gas turbine propulsion systems. The method has been programmed for use in computer aided conceptual aircraft design studies that cover a broad range of Mach numbers and altitudes. The items computed are: inlet size, pressure recovery, additive drag, subsonic spillage drag, bleed and bypass drags, auxiliary air systems drag, boundary-layer diverter drag, nozzle boattail drag, and the interference drag on the region adjacent to multiple nozzle installations. The methods for computing each of these installation effects are described and computer codes for the calculation of these effects are furnished. The results of these methods are compared with selected data for the F-5A and other aircraft. The computer program can be used with uninstalled engine performance information which is currently supplied by a cycle analysis program. The program, including comments, is about 600 FORTRAN statements long, and uses both theoretical and empirical techniques.

Morris, S. J., Jr.↗

NASA NTRS1977

Circumferential distortion modeling of the TF30-P-3 compression system

Circumferential inlet pressure and temperature distortion testing of the TF30 P-3 turbofan engine was conducted. The compressor system at the test conditions run was modelled according to a multiple segment parallel compressor model. Aspects of engine operation and distortion configuration modelled include the effects of compressor bleeds, relative pressure-temperature distortion alignment and circumferential distortion extent. Model predictions for limiting distortion amplitudes and flow distributions within the compression system were compared with test results in order to evaluate predicted trends. Relatively good agreement was obtained. The model also identified the low pressure compressor as the stall-initiating component, which was in agreement with the data.

Mazzawy, R. S.↗

NASA NTRS1977

Directionally solidified eutectic alloy gamma-beta

A pseudobinary eutectic alloy composition was determined by a previously developed bleed-out technique. The directionally solidified eutectic alloy with a composition of Ni-37.4Fe-10.0Cr-9.6Al (in wt%) had tensile strengths decreasing from 1,090 MPa at room temperature to 54 MPa at 1,100 C. The low density, excellent microstructural stability, and oxidation resistance of the alloy during thermal cycling suggest that it might have applicability as a gas turbine vane alloy while its relatively low high temperature strength precludes its use as a blade alloy. A zirconium addition increased the 750 C strength, and a tungsten addition was ineffective. The gamma=beta eutectic alloys appeared to obey a normal freezing relation.

Tewari, S. N.↗

NASA NTRS1977

Development of the technology for the fabrication of reliable laminar from control panels

Materials were assessed and fabrication techniques were developed for use in the manufacture of wing surface materials compatible with the application of both aluminum alloys and nonmetallic composites. The concepts investigated included perforations and slots in the metallic test panels and microporosity and perforations in the composite test panels. Perforations were produced in the metallic test panels by the electron beam process and slots were developed by controlled gaps between the metal sheets. Microporosity was produced in the composite test panels by the resin bleed process, and perforations were produced by the fugitive fiber technique. Each of these concepts was fabricated into test panels, and air flow tests were conducted on the panels.

Meade, L. E.↗

NASA NTRS1977

Performance at Mach 3.52 of a Mach 2.5 all external-compression, two dimensional inlet

A quick-look test was conducted in the 10- by 10-foot supersonic wind tunnel to provide some qualitative information on operating a low Mach number design inlet in a highly oversped condition. The model which is a two-dimensional, three-ramp, all-external-compression inlet designed to match a J-85 turbojet engine at Mach 2.5 was run at Mach 3.52 and zero angle-of-attack and yaw. Results are presented showing the effect of external ramp angle combinations, diffuser ramp position, and amount of bleed.

Cubbison, R. W.↗