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NASA NTRS2006

Flight Test Results for Uniquely Tailored Propulsion-Airframe Aeroacoustic Chevrons: Shockcell Noise

Azimuthally varying chevrons (AVC) which have been uniquely tailored to account for the asymmetric propulsion-airframe aeroacoustic interactions have recently shown significant reductions in jet-related community noise at low-speed take-off conditions in scale model tests of coaxial nozzles with high bypass ratio. There were indications that such AVCs may also provide shockcell noise reductions at high cruise speeds. This paper describes the flight test results when one such AVC concept, namely, the T-fan chevrons with enhanced mixing near the pylon, was tested at full-scale on a modern large twin-jet aircraft (777-300ER) with focus on shockcell noise at mid-cruise conditions. Shockcell noise is part of the interior cabin noise at cruise conditions and its reduction is useful from the viewpoint of passenger comfort. Noise reduction at the source, in the exhaust jet, especially, at low frequencies, is beneficial from the perspective of reduced fuselage sidewall acoustic lining. Results are shown in terms of unsteady pressure spectra both on the exterior surface of the fuselage at several axial stations and also microphone arrays placed inside the fuselage aft of the engine. The benefits of T-fan chevrons, with and without conventional chevrons on the core nozzle, are shown for several engine operating conditions at cruise involving supersonic fan stream and subsonic or supersonic core stream. The T-fan AVC alone provides up to 5 dB low-frequency noise reduction on the fuselage exterior skin and up to 2 dB reduction inside the cabin. Addition of core chevrons appears to increase the higher frequency noise. This flight test result with the previous model test observation that the T-fan AVCs have hardly any cruise thrust coefficient loss (< 0.05%) make them viable candidates for reducing interior cabin noise in high bypass ratio engines.

Mengle, Vinod G.↗

NASA NTRS1999

Summary of HEAT 1 Aeroacoustics Installation Effects

A critical part of the NASA High-Speed Research (HSR) program is the demonstration of satisfactory suppression of the jet noise present at low airspeeds. One scheme for reducing jet exhaust noise generated by a future High-Speed Civil Transport (HSCT) is the use of a mixer/ ejector system which would entrain large quantities of ambient air into the exhaust flow from the powerplant in order to cool and slow the jet exhaust before it leaves the tailpipe. Of the variety of factors which can affect the noise suppression characteristics of the mixer/ejector system, the influence of the wing flow field and high-lift devices is not well understood. The effectiveness of the noise suppression device must be evaluated in the presence of the wing/high-lift system before definitive assessments can be made concerning HSCT noise. Of nearly equal importance is the evaluation of the performance of the high-lift system(s) in the presence of realistic propulsion units which feature high ambient flow entrainment rates and jet thrust coefficients. These noise suppressors must provide the required acoustic attenuation while not overly degrading the thrust efficiency of the propulsion system or the lift enhancement of the high-lift devices on the wing. The overall objective of the NASA High-lift Engine Aeroacoustics Technology program is to demonstrate satisfactory interaction between the jet noise suppressor and the high-lift system at airspeeds and angles of attack consistent with takeoff, climb, approach, and landing. In support of this program, an isolated aeroacoustic test of a 13.5%-scale, candidate mixer/ejector nozzle was performed in the Ames' Research Center 40- by 80-Foot Wind Tunnel. The purpose of the test was to measure the baseline aeroacoustic performance characteristics of this nozzle in isolation from the aerodynamic flowfield induced by an HSCT airframe. The test documented the acoustic signature of the nozzles with treated and hardwall ejector surfaces and with changes in the ratio of ejector-duct-to-jet-area over a wide range range of nozzle pressure ratios and freestream Mach numbers. The test also measured the thrust performance, ambient-flow aspiration ratio, and internal and external static pressures on the nozzles. The isolated aeroacoustic performance data has been compared with results obtained with this nozzle installed on a 13.5% Boeing Reference H HSCT configuration, semi-span model. The semi-span, aeroacoustics integration test documented the first-order effects of the airframe flowfield on the acoustic performance of the nozzle and the effect of the nozzle secondary inlet flows on the aerodynamic performance of the wing high-lift systems. This investigation is critical to understanding the mutual installation effects of mixer/ejector nozzles and wing high-lift systems.

Smith, Brian E.↗

NASA NTRS2016

Tapping the Brake for Entry, Descent, and Landing

A matrix of simulations of hypersonic flow over blunt entry vehicles with steady and pulsing retropropulsion jets is presented. Retropropulsion in the supersonic domain is primarily designed to reduce vehicle velocity directly with thrust. Retropropulsion in the hypersonic domain may enable significant pressure recovery through unsteady, oblique shocks while providing a buffer of reactant gases with relatively low total temperature. Improved pressure recovery, a function of Mach number squared and oblique shock angle, could potentially serve to increase aerodynamic drag in this domain. Pulsing jets are studied to include an additional degree of freedom to search for resonances in an already unsteady flow domain with an objective to maximize the time-averaged drag coefficient. In this paradigm, small jets with minimal footprints of the nozzle exit on the vehicle forebody may be capable of delivering the requisite perturbations to the flow. Simulations are executed assuming inviscid, symmetric flow of a perfect gas to enable a rapid assessment of the parameter space (nozzle geometry, plenum conditions, jet pulse frequency). The pulsed-jet configuration produces moderately larger drag than the constant jet configuration but smaller drag than the jet-off case in this preliminary examination of a single design point. The fundamentals of a new algorithm for this challenging application with time dependent, interacting discontinuities using the feature detection capabilities of Walsh functions are introduced.

Gnoffo, Peter A.↗

NASA NTRS1972

Hydrogen film/conductive cooling

Small scale nozzle tests using heated nitrogen were run to obtain effectiveness and wall heat transfer data with hydrogen film cooling. Effectiveness data are compared with an entrainment model developed from planar, unaccelerated flow data. Results indicate significant effects due to flow turning and acceleration. With injection velocity effects accounted for explicitly, heat transfer correlation coefficients were found to be the same with and without film cooling when properties are evaluated at an appropriate reference temperature for the local gas composition defined by the coolant effectiveness. A design study for an O2/H2 application with 300 psia (207 N/sq cm) chamber pressure and 1500 lbs (6670 N) thrust indicates an adiabatic wall design requires 4 to 5 percent of the total flow as hydrogen film cooling. Internal regenerative cooling designs were found to offer no reduction in coolant requirements.

Ewen, R. L.↗

NASA NTRS1999

High Speed Civil Transport (HSCT) Isolated Nacelle Transonic Boattail Drag Study and Results Using Computational Fluid Dynamics (CFD)

Nozzle boattail drag is significant for the High Speed Civil Transport (HSCT) and can be as high as 25% of the overall propulsion system thrust at transonic conditions. Thus, nozzle boattail drag has the potential to create a thrust-drag pinch and can reduce HSCT aircraft aerodynamic efficiencies at transonic operating conditions. In order to accurately predict HSCT performance, it is imperative that nozzle boattail drag be accurately predicted. Previous methods to predict HSCT nozzle boattail drag were suspect in the transonic regime. In addition, previous prediction methods were unable to account for complex nozzle geometry and were not flexible enough for engine cycle trade studies. A computational fluid dynamics (CFD) effort was conducted by NASA and McDonnell Douglas to evaluate the magnitude and characteristics of HSCT nozzle boattail drag at transonic conditions. A team of engineers used various CFD codes and provided consistent, accurate boattail drag coefficient predictions for a family of HSCT nozzle configurations. The CFD results were incorporated into a nozzle drag database that encompassed the entire HSCT flight regime and provided the basis for an accurate and flexible prediction methodology.

Midea, Anthony C.↗

NASA NTRS2005

High Speed Civil Transport (HSCT) Isolated Nacelle Transonic Boattail Drag Study and Results Using Computational Fluid Dynamics (CFD)

Nozzle boattail drag is significant for the High Speed Civil Transport (HSCT) and can be as high as 25 percent of the overall propulsion system thrust at transonic conditions. Thus, nozzle boattail drag has the potential to create a thrust drag pinch and can reduce HSCT aircraft aerodynamic efficiencies at transonic operating conditions. In order to accurately predict HSCT performance, it is imperative that nozzle boattail drag be accurately predicted. Previous methods to predict HSCT nozzle boattail drag were suspect in the transonic regime. In addition, previous prediction methods were unable to account for complex nozzle geometry and were not flexible enough for engine cycle trade studies. A computational fluid dynamics (CFD) effort was conducted by NASA and McDonnell Douglas to evaluate the magnitude and characteristics of HSCT nozzle boattail drag at transonic conditions. A team of engineers used various CFD codes and provided consistent, accurate boattail drag coefficient predictions for a family of HSCT nozzle configurations. The CFD results were incorporated into a nozzle drag database that encompassed the entire HSCT flight regime and provided the basis for an accurate and flexible prediction methodology.

Midea, Anthony C.↗

NASA NTRS1993

Reliability assessment of thrust chamber cooling concepts using probabilistic analysis techniques

The reliability of OFHC (Oxygen Free High Conductivity) copper and NARloy-Z thrust chambers is assessed by applying probabilistic structural analysis techniques to incorporate design parameter variability and uncertainty. Thrust chambers specifically evaluated are the cylindrical test fixtures employed in a plug-nozzle configuration at the NASA Lewis Research Center. Direct sampling Monte Carlo simulations based on a simplified life prediction methodology established probability densities of firing cycles to structural failure. Simulated cyclic lives demonstrated modest agreement to experiment. Similarly, regions of high structural failure probability were determined using a limit state approach employing calculated cumulative distribution functions for effective stress response and an assumed material strength distribution. A probability of failure of 0.012 was calculated at the center of the coolant channel hot-gas-side wall for an OFHC milled channel. Structural response was found to be sensitive to the uncertainties in the thrust chamber thermal environment and the material's thermal expansion coefficient.

Rapp, Douglas C.↗

NASA NTRS1993

Reliability assessment of thrust chamber cooling concepts using probabilistic analysis techniques

Rapp, Douglas C.↗

NASA NTRSDate not supplied

Testing of Mars Retropropulsion Concepts in the Langley Unitary Plan Wind Tunnel

Future Mars human lander missions will require using retrorockets beginning at supersonic conditions, something which has never been done before at Mars. Significant aerosciences challenges exist due to interactions between the retrorocket exhaust plumes, the freestream flow, and the vehicle that alter the aerodynamic behavior. Historical retropropulsion ground and flight data do not provide a basis for calibrating the computational uncertainties for aerodynamic interference forces and moments on proposed Mars descent vehicles, either because of insufficient data or dissimilar vehicles geometries and/or conditions. A new retropropulsion wind tunnel test was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to address some of the deficiencies of previous tests and to provide a basis for computational model comparisons on relevant geometries. The test included improvements in model design and measured data. Two sub-scale model geometries derived from full-scale vehicle concepts were tested: a blunt low lift-to-drag vehicle and a more slender geometry. The first model includes variations in nozzle parameters that were estimated to have first-order effects on the measurements: location, pointing direction, size, and exit-to-throat area ratio. Both models were instrumented with discrete steady and unsteady pressure sensors on the aerodynamic and interior flow path surfaces. The heats shields of both models also were covered with pressure sensitive paint that effectively measures a steady pressure distribution in an otherwise unsteady flow field. This pressure distribution was integrated over the surface geometry in order to estimate the steady aerodynamic force and moment coefficient contribution from the painted surfaces; this approach was used as a replacement for using a flow-through balance. This paper will cover the test objectives, model design, test matrix, and available test results for each model configuration at different freestream Mach numbers, model attitude, and non-dimensional thrust magnitude. A subsequent publication will cover uncertainty quantification of the test data and comparisons to computational fluid dynamics model results.

Powered Descent↗

NASA NTRS1993

Generic hypersonic vehicle performance model

An integrated computational model of a generic hypersonic vehicle was developed for the purpose of determining the vehicle's performance characteristics, which include the lift, drag, thrust, and moment acting on the vehicle at specified altitude, flight condition, and vehicular configuration. The lift, drag, thrust, and moment are developed for the body fixed coordinate system. These forces and moments arise from both aerodynamic and propulsive sources. SCRAMjet engine performance characteristics, such as fuel flow rate, can also be determined. The vehicle is assumed to be a lifting body with a single aerodynamic control surface. The body shape and control surface location are arbitrary and must be defined. The aerodynamics are calculated using either 2-dimensional Newtonian or modified Newtonian theory and approximate high-Mach-number Prandtl-Meyer expansion theory. Skin-friction drag was also accounted for. The skin-friction drag coefficient is a function of the freestream Mach number. The data for the skin-friction drag coefficient values were taken from NASA Technical Memorandum 102610. The modeling of the vehicle's SCRAMjet engine is based on quasi 1-dimensional gas dynamics for the engine diffuser, nozzle, and the combustor with heat addition. The engine has three variable inputs for control: the engine inlet diffuser area ratio, the total temperature rise through the combustor due to combustion of the fuel, and the engine internal expansion nozzle area ratio. The pressure distribution over the vehicle's lower aft body surface, which acts as an external nozzle, is calculated using a combination of quasi 1-dimensional gas dynamic theory and Newtonian or modified Newtonian theory. The exhaust plume shape is determined by matching the pressure inside the plume, calculated from the gas dynamic equations, with the freestream pressure, calculated from Newtonian or Modified Newtonian theory. In this manner, the pressure distribution along the vehicle after body expansion surface is then determined. The aerodynamic modeling, the engine modeling, and the exhaust plume analysis are described in more detail. A description of the computer code used to perform the above calculations is given and an input/output example is then given. The computer code is available on a Macintosh floppy disk.

Chavez, Frank R.↗

NASA NTRS1992

Flow analysis and design optimization methods for nozzle-afterbody of a hypersonic vehicle

This report summarizes the methods developed for the aerodynamic analysis and the shape optimization of the nozzle-afterbody section of a hypersonic vehicle. Initially, exhaust gases were assumed to be air. Internal-external flows around a single scramjet module were analyzed by solving the 3D Navier-Stokes equations. Then, exhaust gases were simulated by a cold mixture of Freon and Ar. Two different models were used to compute these multispecies flows as they mixed with the hypersonic airflow. Surface and off-surface properties were successfully compared with the experimental data. The Aerodynamic Design Optimization with Sensitivity analysis was then developed. Pre- and postoptimization sensitivity coefficients were derived and used in this quasi-analytical method. These coefficients were also used to predict inexpensively the flow field around a changed shape when the flow field of an unchanged shape was given. Starting with totally arbitrary initial afterbody shapes, independent computations were converged to the same optimum shape, which rendered the maximum axial thrust.

Baysal, O.↗

NASA NTRS1991

Flow analysis and design optimization methods for nozzle afterbody of a hypersonic vehicle

This report summarizes the methods developed for the aerodynamic analysis and the shape optimization of the nozzle-afterbody section of a hypersonic vehicle. Initially, exhaust gases were assumed to be air. Internal-external flows around a single scramjet module were analyzed by solving the three dimensional Navier-Stokes equations. Then, exhaust gases were simulated by a cold mixture of Freon and Argon. Two different models were used to compute these multispecies flows as they mixed with the hypersonic airflow. Surface and off-surface properties were successfully compared with the experimental data. In the second phase of this project, the Aerodynamic Design Optimization with Sensitivity analysis (ADOS) was developed. Pre and post optimization sensitivity coefficients were derived and used in this quasi-analytical method. These coefficients were also used to predict inexpensively the flow field around a changed shape when the flow field of an unchanged shape was given. Starting with totally arbitrary initial afterbody shapes, independent computations were converged to the same optimum shape, which rendered the maximum axial thrust.

Baysal, Oktay↗

NASA NTRS1992

Dynamic interactions between hypersonic vehicle aerodynamics and propulsion system performance

Described here is the development of a flexible simulation model for scramjet hypersonic propulsion systems. The primary goal is determination of sensitivity of the thrust vector and other system parameters to angle of attack changes of the vehicle. Such information is crucial in design and analysis of control system performance for hypersonic vehicles. The code is also intended to be a key element in carrying out dynamic interaction studies involving the influence of vehicle vibrations on propulsion system/control system coupling and flight stability. Simple models are employed to represent the various processes comprising the propulsion system. A method of characteristics (MOC) approach is used to solve the forebody and external nozzle flow fields. This results in a very fast computational algorithm capable of carrying out the vast number of simulation computations needed in guidance, stability, and control studies. The three-dimensional fore- and aft body (nozzle) geometry is characterized by the centerline profiles as represented by a series of coordinate points and body cross-section curvature. The engine module geometry is represented by an adjustable vertical grid to accommodate variations of the field parameters throughout the inlet and combustor. The scramjet inlet is modeled as a two-dimensional supersonic flow containing adjustable sidewall wedges and multiple fuel injection struts. The inlet geometry including the sidewall wedge angles, the number of injection struts, their sweepback relative to the vehicle reference line, and strut cross-section are user selectable. Combustion is currently represented by a Rayleigh line calculation including corrections for variable gas properties; improved models are being developed for this important element of the propulsion flow field. The program generates (1) variation of thrust magnitude and direction with angle of attack, (2) pitching moment and line of action of the thrust vector, (3) pressure and temperature distributions throughout the system, and (4) performance parameters such as thrust coefficient, specific impulse, mass flow rates, and equivalence ratio. Preliminary results are in good agreement with available performance data for systems resembling the NASP vehicle configuration.

Flandro, G. A.↗

NASA NTRS2008

Influence of Reduced Mass Flow Rate and Chamber Backpressure on Swirl Injector Fluid Mechanics

Industry interest in variable-thrust liquid rocket engines places a demand on engine injector technology to operate over a wide range of liquid mass flow rates and chamber backpressures. One injection technology of current interest for variable thrust applications is an injector design with swirled fluids. Current swirl injector design methodologies do not take into account how swirl injector design parameters respond to elevated chamber backpressures at less than design mass flow rates. The current work was created to improve state-of-the-art swirl injector design methods in this area. The specific objective was to study the effects of elevated chamber backpressure and off-design mass flow rates on swirl injector fluid mechanics. Using a backpressure chamber with optical access, water was flowed through a swirl injector at various combinations of chamber backpressure and mass flow rates. The film thickness profile down the swirl injector nozzle section was measured through a transparent nozzle section of the injector. High speed video showed measurable increases in the film thickness profile with application of chamber backpressure and mass flow rates less than design. At prescribed combinations of chamber backpressure and injected mass flow rate, a discrete change in the film thickness profile was observed. Measured injector discharge coefficient values showed different trends with increasing chamber backpressure at low mass flow rates as opposed to near-design mass flow rates. Downstream spray angles showed classic changes in morphology as the mass flow rate was decreased below the design value. Increasing chamber backpressure decreased the spray angle at any injection mass flow rate. Experimental measurements and discussion of these results are reported in this paper.

Kenny, R Jeremy↗

NASA NTRS2003

Derivation of the Data Reduction Equations for the Calibration of the Six-component Thrust Stand in the CE-22 Advanced Nozzle Test Facility

This paper documents the derivation of the data reduction equations for the calibration of the six-component thrust stand located in the CE-22 Advanced Nozzle Test Facility. The purpose of the calibration is to determine the first-order interactions between the axial, lateral, and vertical load cells (second-order interactions are assumed to be negligible). In an ideal system, the measurements made by the thrust stand along the three coordinate axes should be independent. For example, when a test article applies an axial force on the thrust stand, the axial load cells should measure the full magnitude of the force, while the off-axis load cells (lateral and vertical) should read zero. Likewise, if a lateral force is applied, the lateral load cells should measure the entire force, while the axial and vertical load cells should read zero. However, in real-world systems, there may be interactions between the load cells. Through proper design of the thrust stand, these interactions can be minimized, but are hard to eliminate entirely. Therefore, the purpose of the thrust stand calibration is to account for these interactions, so that necessary corrections can be made during testing. These corrections can be expressed in the form of an interaction matrix, and this paper shows the derivation of the equations used to obtain the coefficients in this matrix.

Wong, Kin C.↗

NASA NTRS1999

Spreading Characteristics of Compressible Jets from Nozzles of Various Geometries

The spreading characteristics of jets from several asymmetric nozzles, and a set of rectangular orifices are compared, covering a jet Mach number range of 0.3-2.0. The effect of 'tabs' for a rectangular and a round nozzle is also included in the comparison. Compared to a round jet, the jets from the asymmetric nozzles spread only slightly more at subsonic conditions whereas at supersonic conditions, when 'screech' occurs, they spread much more. The dynamics of the azimuthal vortical structures of the jet, organized and intensified under the screeching condition, are thought to be responsible for the observed effect at supersonic conditions. Curiously, the jet from a 'lobed' nozzle spreads much less at supersonic condition compared to all other cases; this is due to the absence of screech with this nozzle. Screech stages inducing flapping, rather than varicose or helical, flow oscillation cause a more pronounced jet spreading. At subsonic conditions, only a slight increase in jet spreading with the asymmetric nozzles contrasts previous observations by others. The present results show that the spreading of most asymmetric jets is not much different from that of a round jet. This inference is further supported by data from the rectangular orifices. In fact, jets from the orifices with small aspect ratio (AR) exhibit virtually no increase in the spreading. A noticeable increase commences only when AR is larger than about 10. Thus, 'shear layer perimeter stretching', achieved with a larger AR for a given cross-sectional area of the orifice, by itself, proves to be a relatively inefficient mechanism for increasing jet spreading. In contrast, the presence of streamwise vortices or 'natural excitation' can cause a significant increase - effects that might explain the observations in the previous investigations. Thus far, the biggest increase in jet spreading is observed with the tabs. This is true in the subsonic regime, as well as in the supersonic regime, in spite of the fact that screech is eliminated by the tabs. The characteristic spreading of the tabbed jets is explained by the induced motion of the tab-generated streamwise vortex pairs. The tabs, however, incur thrust loss; the flow blockage and loss in thrust coefficient, vis-a-vis the spreading increase, are evaluated for various configurations.

Zaman, K. B. M. Q.↗

NASA NTRS1956

Ram-jet Performance

The ram jet is basically one of the most dimple types of aircraft engine. It consists only of an inlet diffuser, a combustion system, and an exit nozzle. A typical ram-jet configuration is shown in figure 128. The engine operates on the Brayton cycle, and ideal cycle efficiency depends only on the ratio of engine to ambient pressure. The increased, engine pressures are obtained by ram action alone, and for this reason the ram jet has zero thrust at zero speed. Therefore, ram-jet-powered aircraft must be boosted to flight speeds close to a Mach number of 1.0 before appreciable thrust is generated by the engine. Since pressure increases are obtained by ram action alone, combustor-inlet pressures and temperatures are controlled by the flight speed, the ambient atmospheric condition, and by the efficiency of the inlet diffuser. These pressures and temperatures, as functions of flight speed and altitude, are shown in figure 129 for the NACA standard atmosphere and for practical values of diffuser efficiency. It can be seen that very wide ranges of combustor-inlet temperatures and pressures may be encountered over the ranges of flight velocity and altitude at which ram jets may be operated. Combustor-inlet temperatures from 500 degrees to 1500 degrees R and inlet pressures from 5 to 100 pounds per square inch absolute represent the approximate ranges of interest in current combustor development work. Since the ram jet has no moving parts in the combustor outlet, higher exhaust-gas temperatures than those used in current turbojets are permissible. Therefore, fuel-air ratios equivalent to maximum rates of air specific impulse or heat release can be used, and, for hydrocarbon fuels, this weight ratio is about 0.070. Lower fuel-air ratios down to about 0.015 may also be required to permit efficient cruise operation. This fuel-air-ratio range of 0.015 to 0.070 used in ram jets can be compared with the fuel-air ratios up to 0.025 encountered in current turbojets. Ram-jet combustor-inlet velocities range from 150 to 400 feet per second. These high linear velocities combined with the relatively low pressure ratios obtainable in ram jets require that the pressure drop through the combustor be kept low to avoid excessive losses in cycle efficiency. It has been estimated that, for a long-range ram-jet engine, an increase in pressure loss of one dynamic head would require a compensating 1-percent increase in combustion efficiency. Therefore, combustor pressure-loss coefficients (pressure drop/impact pressure) of the order of 1 to 4 are found in most current engines. The operating conditions described impose major problems in the design of stable and efficient ram-jet combustion systems. This chapter presents a survey of ram-jet combustor research and, where possible, points out criteria that may be useful in the design of ram-jet combustion systems.

Cervenko, A. J.↗