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At least 433 records · Page 24

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NASA NTRS1986

Plume characteristics of single-stream and dual-flow conventional and inverted-profile nozzles at equal thrust

The plume velocity and temperature decay rates of single-stream, conventional dual-flow and inverted-profile dual-flow nozzles are compared at equal values of ideal thrust over a wide range of flow conditions. The comparisons are made in terms of constant velocity and temperature contour maps. The results show that both dual-flow nozzle types have much greater plume velocity and temperature decay rates than those of equivalent thrust single-stream nozzles when the respective secondary flows were at ambient temperature. With hot secondary flows, the inverted-profile dual-flow plumes decayed significantly faster than those of single-stream nozzles; however, the decay rates for the conventional dual-flow streams were about the same as those for the single-stream nozzles. Consequently, with hot secondary flows, the inverted-profile dual-flow plumes decayed much faster than the conventional dual-flow plumes at equal thrust.

Vonglahn, U. H.↗

NASA NTRS1986

Multiaxis aircraft control power from thrust vectoring at high angles of attack

Extensive research programs conducted at the Langley Research Center have shown that thrust vectoring can be provided by multifunction (nonaxisymmetric) nozzles. Most of this research has been conducted on pitch vectoring at both static and forward flight conditions. Recent efforts have been aimed at evaluating yaw vectoring concepts at static (wind off) conditions. This paper summarizes results for three different twin-engine fighter configurations tested over a Mach number range of 0.15 to 2.47 at angles of attack up to 35 deg. The objective of these investigations was to determine the multiaxis control power characteristics provided by thrust vectoring. All three configurations employed two-dimensional convergent-divergent nozzles which provided pitch vectoring by differential deflection of the upper and lower nozzle divergent flaps. Three different means of yaw vectoring were tested: (1) a translating nozzle sidewall; (2) yaw flaps located in the nozzle sidewalls; and (3) canted nozzles. These investigations were conducted in the Langley 16-Foot Transonic Tunnel and the Lewis 10x10-Foot Supersonic Tunnel. Longitudinal and direction control power from thrust vectoring was greater than that provided by aerodynamic control effectors at low speed or at high angles of attack.

Capone, F. J.↗

NASA NTRS1985

A zonal computational procedure adapted to the optimization of two-dimensional thrust augmentor inlets

A viscous-inviscid interaction methodology based on a zonal description of the flowfield is developed as a mean of predicting the performance of two-dimensional thrust augmenting ejectors. An inviscid zone comprising the irrotational flow about the device is patched together with a viscous zone containing the turbulent mixing flow. The inviscid region is computed by a higher order panel method, while an integral method is used for the description of the viscous part. A non-linear, constrained optimization study is undertaken for the design of the inlet region. In this study, the viscous-inviscid analysis is complemented with a boundary layer calculation to account for flow separation from the walls of the inlet region. The thrust-based Reynolds number as well as the free stream velocity are shown to be important parameters in the design of a thrust augmentor inlet.

Lund, T. S.↗

NASA NTRS1986

Test program to provide confidence in liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

An experimental program has been planned at the NASA Lewis Research Center to build confidence in the feasibility of liquid oxygen cooling for hydrocarbon fueled rocket engines. Although liquid oxygen cooling has previously been incorporated in test hardware, more runtime is necessary to gain confidence in this concept. In the previous tests, small oxygen leaks developed at the throat of the thrust chamber and film cooled the hot-gas side of the chamber wall without resulting in catastrophic failure. However, more testing is necessary to demonstrate that a catastrophic failure would not occur if cracks developed further upstream between the injector and the throat, where the boundary layer has not been established. Since under normal conditions cracks are expected to form in the throat region of the thrust chamber, cracks must be initiated artificially in order to control their location. Several methods of crack initiation are discussed in this report. Four thrust chambers, three with cracks and one without, should be tested. The axial location of the cracks should be varied parametrically. Each chamber should be instrumented to determine the effects of the cracks, as well as the overall performance and durability of the chambers.

Armstrong, E. S.↗

NASA NTRS1985

Net thrust calculation sensitivity of an afterburning turbofan engine to variations in input parameters

The calculated value of net thrust of an aircraft powered by a General Electric F404-GE-400 afterburning turbofan engine was evaluated for its sensitivity to various input parameters. The effects of a 1.0-percent change in each input parameter on the calculated value of net thrust with two calculation methods are compared. This paper presents the results of these comparisons and also gives the estimated accuracy of the overall net thrust calculation as determined from the influence coefficients and estimated parameter measurement accuracies.

Hughes, D. L.↗

NASA NTRS1986

Plume characteristics of single-stream and dual-flow conventional and inverted-profile nozzles at equal thrust

Von Glahn, U. H.↗

NASA NTRS1986

Effect of port corner geometry on the internal performance of a rotating-vane-type thrust reverser

An investigation has been conducted in the static-test facility of the Langley 16-Foot Transonic Tunnel to determine the effects of reverser port geometry on the internal performance of a nonaxisymmetric rotating-vane-type thrust reverser. Thrust reverser vane positions representing a spoiled-trust (partially deployed) position and a full-reverse-thrust (fully deployed) position were tested with each port geometry variable. The effects of upstream port corner radius and wall angle on internal performance were determined. In addition, the effect of the length of a simulated cooling liner (blunt-base step) near the reverser port entrance was investigated; five different lengths were tested. All tests were conducted with no external flows, and nozzle pressure ratio was varied from 1.2 to 5.0.

Berrier, B. L.↗

NASA NTRS1987

Breadboard RL10-11B low thrust operating mode

Cryogenic space engines require a cooling process to condition engine hardware to operating temperature before start. This can be accomplished most efficiently by burning propellants that would otherwise be dumped overboard after cooling the engine. The resultant low thrust operating modes are called Tank Head Idle and Pumped Idle. During February 1984, Pratt & Whitney conducted a series of tests demonstrating operation of the RL10 rocket engines at low thrust levels using a previously untried hydrogen/oxygen heat exchanger. The initial testing of the RL10-11B Breadboard Low Thrust Engine is described. The testing demonstrated operation at both tank head idle and pumped idle modes.

Kmiec, Thomas D.↗

NASA NTRS1987

Full-scale thrust reverser testing in an altitude facility

A two-dimensional convergent-divergent exhaust nozzle designed and fabricated by Pratt and Whitney Aircraft was installed on a PW1128 turbofan engine and tested during thrust reverser operation in an altitude facility at NASA Lewis Research Center. A unique collection system was used to capture the thrust reverser exhaust gas and transport it to the primary exhaust collector. Tests were conducted at three flight conditions with varying amounts of thrust reverse at each condition. Some reverser exhaust gas spillage by the collection system was encountered but engine performance was unaffected at all flight conditions tested. Based on the results of this test program, the feasibility of altitude testing of advanced multi-function exhaust nozzle systems has been demonstrated.

Mehalic, Charles M.↗

NASA NTRS1987

Effect of varying internal geometry on the static performance of rectangular thrust-reverser ports

An investigation has been conducted to evaluate the effects of several geometric parameters on the internal performance of rectangular thrust-reverser ports for nonaxisymmetric nozzles. Internal geometry was varied with a test apparatus which simulated a forward-flight nozzle with a single, fully deployed reverser port. The test apparatus was designed to simulate thrust reversal (conceptually) either in the convergent section of the nozzle or in the constant-area duct just upstream of the nozzle. The main geometric parameters investigated were port angle, port corner radius, port location, and internal flow blocker angle. For all reverser port geometries, the port opening had an aspect ratio (throat width to throat height) of 6.1 and had a constant passage area from the geometric port throat to the exit. Reverser-port internal performance and thrust-vector angles computed from force-balance measurements are presented.

Re, Richard J.↗

NASA NTRS1987

Multiaxis control power from thrust vectoring for a supersonic fighter aircraft model at Mach 0.20 to 2.47

The aeropropulsive characteristics of an advanced twin-engine fighter aircraft designed for supersonic cruise have been studied in the Langley 16-Foot Tansonic Tunnel and the Lewis 10- by 10-Foot Supersonic Tunnel. The objective was to determine multiaxis control-power characteristics from thrust vectoring. A two-dimensional convergent-divergent nozzle was designed to provide yaw vector angles of 0, -10, and -20 deg combined with geometric pitch vector angles of 0 and 15 deg. Yaw thrust vectoring was provided by yaw flaps located in the nozzle sidewalls. Roll control was obtained from differential pitch vectoring. This investigation was conducted at Mach numbers from 0.20 to 2.47. Angle of attack was varied from 0 to about 19 deg, and nozzle pressure ratio was varied from about 1 (jet off) to 28, depending on Mach number. Increments in force or moment coefficient that result from pitch or yaw thrust vectoring remain essentially constant over the entire angle-of-attack range of all Mach numbers tested. There was no effect of pitch vectoring on the lateral aerodynamic forces and moments and only very small effects of yaw vectoring on the longitudinal aerodynamic forces and moments. This result indicates little cross-coupling of control forces and moments for combined pitch-yaw vectoring.

Capone, Francis J.↗

NASA NTRS1986

Proven, long-life hydrogen/oxygen thrust chambers for space station propulsion

A 25 lb sub f hydrogen/oxygen thruster has been developed and proven as a viable candidate to meet the needs of the Space Station Program. Likewise, a 50 lb sub f hydrogen/oxygen thrust chamber has been developed and has demonstrated reliable, long-life expectancy at anticipated Space Station operating conditions. Both these thrust chambers were based on design criteria developed in previous thruster programs. Extensive thermal analysis and models were used to design the thrusters to achieve total impulse goals of 2 million lb sub f sec. Test data from each thruster are compared to the analytical predictions for the performance and heat transfer characteristics. Also, the results of thrust chamber life verification tests are presented.

Richter, G. Paul↗

NASA NTRS1986

Vacuum chamber pressure effects on thrust measurements of low Reynolds number nozzles

Tests were conducted to investigate the effect of vacuum facility pressure on the performance of small thruster nozzles. Thrust measurements of two converging-diverging nozzles with an area ratio of 140 and an orifice plate flowing unheated nitrogen and hydrogen were taken over a wide range of vacuum facility pressures and nozzle throat Reynolds numbers. In the Reynolds number range of 2200 to 12,000 there was no discernable viscous effect on thrust below an ambient to total pressure ratio of 1000. In nearly all cases, flow separation occurred at a pressure ratio of about 1000. This was the upper limit for obtaining an accurate thrust measurement for a conical nozzle with an area ratio of 140.

Sovey, J. S.↗

NASA NTRS1986

Multiaxis aircraft control power from thrust vectoring at high angles of attack

Extensive research programs conducted at the Langley Research Center have shown that thrust vectoring can be provided by multifunction (nonaxisymmetric) nozzles. Most of this research has been conducted on pitch vectoring at both static and forward flight conditions. Recent efforts have been aimed at evaluating yaw vectoring concepts at static (wind off) conditions. This paper summarizes results for three different twin-engine fighter configurations tested over a Mach number range of 0.15 to 2.47 at angles of attack up to 35 deg. The objective of these investigations was to determine the multiaxis control power characteristics provided by thrust vectoring. All three configurations employed two-dimensional convergent-divergent nozzles which provided pitch vectoring by differential deflection of the upper and lower nozzle divergent flaps. Three different means of yaw vectoring were tested: (1) a translating nozzle sidewall; (2) yaw flaps located in the nozzle sidewalls; and (3) canted nozzles. These investigations were conducted in the Langley 16-Foot Transonic Tunnel and the Lewis 10 x 10-Foot Supersonic Tunnel. Longitudinal and direction control power from thrust vectoring was greater than that provided by aerodynamic control effectors at low speed or at high angles of attack.

Capone, Francis J.↗

NASA NTRS1987

Full-scale thrust reverser testing in an altitude facility

A two-dimensional convergent-divergent exhaust nozzle designed and fabricated by Pratt and Whitney Aircraft was installed on a PW1128 turbofan engine and tested during thrust reverser operation in an altitude facility at NASA Lewis Research Center. A unique collection system was used to capture the thrust reverser exhaust gas and transport it to the primary exhaust collector. Tests were conducted at three flight conditions with varying amounts of thrust reverse at each condition. Some reverser exhaust gas spillage by the collection system was encountered but engine performance was unaffected at all flight conditions tested. Based on the results of this test program, the feasibility of altitude testing of advanced multifunction exhaust nozzle systems has been demonstrated.

Mehalic, Charles M.↗

NASA NTRS1988

Static performance of an axisymmetric nozzle with post-exit vanes for multiaxis thrust vectoring

An investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to determine the flow-turning capability and the nozzle internal performance of an axisymmetric convergent-divergent nozzle with post-exit vanes installed for multiaxis thrust vectoring. The effects of vane curvature, vane location relative to the nozzle exit, number of vanes, and vane deflection angle were determined. A comparison of the post-exit-vane thrust-vectoring concept with other thrust-vectoring concepts is provided. All tests were conducted with no external flow, and nozzle pressure ratio was varied from 1.6 to 6.0.

Berrier, Bobby L.↗

NASA NTRS1987

Numerical grid generation and flow simulation in SSME thrust chamber

The development of liquid and solid rocket engines for future space projects demands a detailed optimization process for highly efficient performance and cost reasons. Also, testing of full size engines may not be feasible when the large size requires test facilities which are cost prohibitive or if vacuum operation cannot be acquired. For such situations only scaling from small test scale measurements or accurate analytical predictions will provide the performance prior to actually flying the mission. A rigorous approach for simulating the combustion processes in liquid rocket engines by employing a direct solution of Navier-Stokes equations within the entire volume of the thrust chambers is presented. This method is illustrated in the solution of reactive flow in the Space Shuttle Main Engine (SSME) thrust chamber. The objective is to review recent improvements in the mathematical model and to present the grid generation methodology suitable for rocket thrust chamber geometries.

Gross, K. W.↗

NASA NTRS1988

Wave processes in scramjet thrust generation

The investigation explores the steady, supersonic expansion and compression wave processes involved in generating thrust from a combustion jet in a two-dimensional scramjet. Two types of thrust generating wave are produced by the jet. The first is the injection compression waves associated with injection and combustion of the fuel, and the second is the expansion interaction waves associated with interaction between the exhaust nozzle expansion and the jet. Analysis shows that the strength of a wave is not modified by crossing the jet, which implies that, with a given flow configuration, the jets in a multiple jet combustion array produce thrust independently of one another. Shock tunnel experiments with hydrogen fuel and three simple scramjet configurations are used to identify the two types of waves, to confirm the effect on a wave of crossing the jet, and to demonstrate the influence of the basic, fuel-off flow configuration on the effects produced by the waves.

Stalker, R. J.↗