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At least 127 records · Page 7

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NASA NTRS1961

Annular Internal-External-Expansion Rocket Nozzles for Large Booster Applications

For large-thrust booster applications, annular rocket nozzles employing both internal and external expansion are investigated. In these nozzles, free-stream air flows through the center as well as around the outside of the exiting jet. Flaps for deflecting the rocket exhaust are incorporated on the external-expansion surface for thrust-vector control. In order to define nozzle off-design performance, thrust vectoring effectiveness, and external stream effects, an experimental investigation was conducted on two annular nozzles with area ratios of 15 and 25 at Mach 0, 2, and 3 in the Lewis 10- by 10-foot wind tunnel. Air, pressurized to 600 pounds per square inch absolute, was used to simulate the exhaust flow. For a nozzle-pressure-ratio range of 40 to 1000, the ratio of actual to ideal thrust was essentially constant at 0.98 for both nozzles. Compared with conventional convergent-divergent configurations on hypothetical boost missions, the performance gains of the annular nozzle could yield significant orbital payload increases (possibly 8 to 17 percent). A single flap on the external-expansion surface of the area-ratio-25 annular nozzle produced a side force equal to 4 percent of the axial force with no measurable loss in axial thrust.

Connors, James F.↗

NASA NTRS1988

Techniques utilized in the simulated altitude testing of a 2D-CD vectoring and reversing nozzle

Simulated altitude testing of a two-dimensional, convergent-divergent, thrust vectoring and reversing exhaust nozzle was accomplished. An important objective of this test was to develop test hardware and techniques to properly operate a vectoring and reversing nozzle within the confines of an altitude test facility. This report presents detailed information on the major test support systems utilized, the operational performance of the systems and the problems encountered, and test equipment improvements recommended for future tests. The most challenging support systems included the multi-axis thrust measurement system, vectored and reverse exhaust gas collection systems, and infrared temperature measurement systems used to evaluate and monitor the nozzle. The feasibility of testing a vectoring and reversing nozzle of this type in an altitude chamber was successfully demonstrated. Supporting systems performed as required. During reverser operation, engine exhaust gases were successfully captured and turned downstream. However, a small amount of exhaust gas spilled out the collector ducts' inlet openings when the reverser was opened more than 60 percent. The spillage did not affect engine or nozzle performance. The three infrared systems which viewed the nozzle through the exhaust collection system worked remarkably well considering the harsh environment.

Block, H. Bruce↗

NASA NTRS1984

Supersonic jet shock noise reduction

Shock-cell noise is identified to be a potentially significant problem for advanced supersonic aircraft at takeoff. Therefore NASA conducted fundamental studies of the phenomena involved and model-scale experiments aimed at developing means of noise reduction. The results of a series of studies conducted to determine means by which supersonic jet shock noise can be reduced to acceptable levels for advanced supersonic cruise aircraft are reviewed. Theoretical studies were conducted on the shock associated noise of supersonic jets from convergent-divergent (C-D) nozzles. Laboratory studies were conducted on the influence of narrowband shock screech on broadband noise and on means of screech reduction. The usefulness of C-D nozzle passages was investigated at model scale for single-stream and dual-stream nozzles. The effect of off-design pressure ratio was determined under static and simulated flight conditions for jet temperatures up to 960 K. Annular and coannular flow passages with center plugs and multi-element suppressor nozzles were evaluated, and the effect of plug tip geometry was established. In addition to the far-field acoustic data, mean and turbulent velocity distributions were measured with a laser velocimeter, and shadowgraph images of the flow field were obtained.

Stone, J. R.↗

NASA NTRS1984

Supersonic jet shock noise reduction

Stone, J. R.↗

NASA NTRS1956

The Compressible Laminar Boundary Layer with Heat Transfer and Arbitrary Pressure Gradient

An approximate method for the calculation of the compressible laminar boundary layer with heat transfer and arbitrary pressure gradient, based on Thwaites' correlation concept, is presented. With the definition of dimensionless shear and heat-transfer parameters and an assumed correlation of these parameters in terms of a momentum parameter, a complete system of relations for calculating skin friction and heat transfer results. Knowledge of velocity or temperature profiles is not necessary in using this calculation method. When the method is applied to a convergent-divergent, axially symmetric rocket nozzle, it shows that high rates of heat transfer are obtained at the initial stagnation point and at the throat of the nozzle. Also indicated are negative displacement thicknesses in the convergent portion of the nozzle; these occur because of the high density within the lower portions of the cooled boundary layer. (author)

Cohen, Clarence B↗

NASA NTRS1979

Performance characteristics of nonaxisymmetric nozzles installed on the F-18 aircraft

The Langley Research Center has conducted an experimental program on a model of the F-18 airplane to determine the performance of nonaxisymmetric nozzles relative to the aircraft's baseline axisymmetric nozzle. The performance of a single expansion ramp (ADEN) and two-dimensional convergent-divergent (2-D C-D) nozzle were compared to the baseline axisymmetric nozzles. The effects of vectoring and reversing were also studied. Performance of a modified YF-17 airplane with the ADEN nozzle was also estimated. The results of this investigation indicate that nonaxisymmetric nozzles can be installed on a twin engine fighter airplane with equal or better performance than axisymmetric nozzles. The nonaxisymmetric nozzles also offer potential for innovative and improved aircraft maneuver through thrust vectoring and reversing. The YF-17/ADEN flown as a technology demonstrator would have reduced performance compared to an unmodified YF-17. However, on an equal aircraft weight basis, performance would essentially be equivalent. This study also showed that the YF-17 can serve as a testbed to validate nonaxisymmetric nozzle technology.

Capone, F. J.↗

NASA NTRS1979

Effect of several geometric parameters on the static internal performance of three nonaxisymmetric nozzle concepts

Effects of several geometric parameters on the internal performance of nonaxisymmetric convergent-divergent, single-ramp expansion, and wedge nozzles were investigated at nozzle pressure ratios up to approximately 10. In addition, two different thrust-vectoring schemes were investigated with the wedge nozzle. The results indicated that as with conventional round nozzles, peak nonaxisymmetric nozzle, internal performance occurred near the nozzle pressure ratio required for fully expanded exhaust flow. Nozzle sidewall length or area generally had little effect on the internal performance of the nozzles investigated.

Berrier, B. L.↗

NASA NTRS1983

Advanced technology exhaust nozzle development

The requirement for greater tactical aircraft operational capabilities has led to increasing research emphasis on the refinement of engine/airframe integration methods and exhaust nozzle flexibility. A major prospective advancement in the development of these capabilities takes the form of multifunctional exhaust nozzle systems with thrust reversal and thrust vectoring features, whose operation will be shared by both airframe and powerplant control systems. Attention is presently given to the two-dimensional convergent-divergent and single expansion ramp nozzle designs, with emphasis on the variable geometry mechanical systems by which they assume cruising flight, vectoring, and thrust reversal operations. The nozzles have been wind tunnel model-tested for the cases of the F-18 fighter and a supersonic cruise configuration concept.

Dusa, D. J.↗

NASA NTRS1959

Overexpanded Performance of Conical Nozzles with Area Ratios of 6 and 9 With and Without Supersonic External Flow

An investigation of the thrust characteristics and internal pressure distributions of two convergent-divergent 15 deg. half-angle exhaust nozzles having area ratios of 6 and 9 was made in the NASA Lewis 10- by 10-foot supersonic wind tunnel. The tests were conducted at free-stream Mach numbers of 0, 2.0, 2.5, 3.0, and 3.5 over a range of nozzle pressure ratios from 3 to 105. Attempts were made to induce separation of the overexpanded nozzle flow using secondary airflow and a wedge. Nozzle flow expansion under all free-stream conditions followed one-dimensional theory until separation from the nozzle wall occurred. In quiescent air the nozzle flow expanded to a pressure approximately one-half the base pressure before separation. When the nozzles were tested with supersonic external flow at the same effective pressure ratios, the nozzle flow separated with negligible expansion below the base pressure. The effect of a supersonic stream on internal nozzle flow separation characteristics was well defined only at a free-stream Mach number of 2.0. Thrust data at supersonic free-stream conditions indicate that only a small percentage of the ideal nozzle thrust will be available at nozzle pressure ratios below design. However, the overexpanded primary nozzle thrust loss was decreased by injecting large quantities of secondary air near the nozzle exit. In most cases no net gain in thrust resulted from secondary-air injection when the nozzle thrust was compared with the ideal thrust of both the primary and secondary airflows.

Musial, Norman T.↗

NASA NTRS1990

Exhaust environment measurements of a turbofan engine equipped with an afterburner and 2D nozzle

A test to measure the acoustic noise and static pressure environment on a structure exposed to engine exhaust flow was conducted at the NASA Lewis Research Center Propulsion Systems Laboratory by using an F100 derivative engine with a two-dimensional convergent-divergent (2D/CD) non-flight-weight demonstrator nozzle. A highly instrumented, water cooled flat panel was placed behind the 2D/CD nozzle, and tests were conducted at simulated Mach/altitude flight conditions with the engine at military or maximum-afterburner power setting. The panel instrumentation consisted of acoustic pressure microphones, thermocouples, and static-pressure pickups. Measurements indicated that the exhaust environment may excite structural resonances up to 10,000 HZ and that overall sound pressure levels decrease with increasing altitude.

Brase, L. O.↗

NASA NTRS1984

Experimental investigation of shock-cell noise reduction for dual-stream nozzles in simulated flight

Six scale-model nozzles were tested in an anechoic facility to evauate the effectiveness of convergent-divergent (C-D) terminations in reducing shock-cell noise of unsuppressed and mechanically suppressed coannular plug nozzles. One hundred fifty-three acoustic test points with inverted velocity profiles were conducted under static and simulated flight conditions. Diagnostic flow visualization with a shadowgraph and velocity measurements with a laser velocimeter were performed on selected plumes. Shock-cells were identified on the plug and downstream of the plug of the unsuppressed convergent coannular nozzle with truncated plug. Broadband peak frequencies predicted with the two shock-cell structures were correlated with the observed spectra using the measured shock-cell spacings. Relative to a convergent circular nozzle, the perceived noise level (PNL) data at an observer angle of 60 deg relative to inlet, indicated a reduction of (1) 6.5 dB and 9.2 dB with unsuppressed C-D coannular nozzle with truncated plug and (2) 7.7 dB and 8.3 dB with suppressed C-D coannular nozzle under static and simulated flight conditions, espectively. The unsuppressed C-D coannular nozzle with truncated plug, operating at the C-D design condition, had shock-cells downstream of the plug with no shock-cells on the plug. The downstream shock-cells were eliminated by replacing the truncated plug with a smooth extension to obtain an additional 2.4 dB and 3 dB front quadrant PNL reduction, under static and simulated flight conditions, respectively. Other results are discussed.

Janardan, B. A.↗

NASA NTRS1978

A review of thrust-vectoring schemes for fighter applications

This paper presents a review of thrust vectoring schemes for advanced fighter applications. Results are presented from wind tunnel and system integration studies on thrust vectoring nozzle concepts. Vectoring data are presented from wind tunnel tests of axisymmetric C-D (convergent-divergent) and nonaxisymmetric wedge, C-D, single ramp and USB (upper-surface blowing) nozzle concepts. Results from recent airframe/nozzle integration studies on the impact of thrust vectoring on weight, cooling and performance characteristics are discussed. This review indicates that the aircraft designer has, at his disposal, a wide range of thrust vectoring schemes which offer potential for added or improved aircraft capability.

Berrier, B. L.↗

NASA NTRS1957

Survey of the Acoustic near Field of Three Nozzles at a Pressure Ratio of 30

The sound pressures radiating from the exhaust streams of two convergent-divergent and one convergent nozzle were measured. Exit diameters were 1.206 in. for the expanded nozzle and 0.625 in. for the convergent nozzle. The results are presented in a series of contour maps of overall and fine 1/3-octave-band sound pressures. The location of the source of the noise in each 1/3-octave band in the frequency range of 30 to 16,000 cps and the total power radiated were determined and compared with those of subsonic jets.

FLOW, JET-MIXING↗

NASA NTRS1960

Experimental Evaluation of Rocket Exhaust Diffusers for Altitude Simulation

An experimental investigation of exhaust diffusers has been conducted to evaluate various methods of minimizing the overall pressure ratio (from chamber to ambient pressure) required to establish and maintain full expansion of the nozzle flow (altitude simulation). Exhaust-diffuser configurations investigated were (1) cylindrical diffusers, (2) diffusers with contraction, and (3) diffusers including a right-angle turn. Cylindrical diffusers were evaluated with primary nozzles of various area ratios and types, as well as two clustered configurations; the other diffusers were evaluated with individual nozzles of constant area ratio and varied type. Air was the working fluid, except for two check points obtained with JP-4 fuel and liquid-oxygen rocket engines and cylindrical diffusers. The minimum length-diameter ratio of cylindrical diffusers was about 6 for minimum pressure-ratio requirements. With cylindrical diffusers of adequate length, the pressure-ratio requirements were primarily a function of the ratio of diffuser to nozzle-throat areas and were essentially independent of primary-nozzle type (including two clustered configurations) or area ratio. The two check points obtained with rocket engines indicated the pressure-ratio requirements at given ratios of diffuser to nozzle-throat areas were lowered, as compared with the requirements with air, as a result of the reduced ratio of specific heats. The minimum length-diameter ratio of the contraction throat of convergent-divergent diffusers was also about 6 for minimum pressure-ratio requirements. With adequate contraction-throat length, the pressure-ratio requirements of such diffusers were appreciably below those of comparable cylindrical diffusers when used with conical and cutoff-isentropic nozzles, but not when used with a bell nozzle. Minimum pressure-ratio requirements of a diffuser including a simple long-radius right-angle turn at maximum diffuser area, obtained with the center of radius of the turn a minimum of 2 diffuser diameters downstream of the nozzle exit, were not appreciably above those of a comparable optimum cylindrical diffuser. A diffuser including a long-radius right-angle turn at a contraction minimum area had somewhat lower pressure-ratio requirements than the aforementioned simple turn.

Sivo, Joseph N.↗

NASA NTRS1989

Supersonic jet noise and the high speed civil transport

An evaluation is made of the comparative advantages of prospective SST engine noise-suppression systems, with a view to their effectiveness in meeting the federally-mandated community noise standards of FAR 36 Stage III. A noise-suppression system must be capable of removing at least 4 EPNdB of noise percent thrust loss at takeoff. While none of the suppressors presently discussed is capable of meeting this goal, the inverted velocity profile/annular convergent-divergent plug/acoustically-treated ejector suppressor combination of configurational elements appears to represent the most efficient noise-control apparatus. Noncircular cross-section nozzle geometries also furnish a general noise reduction advantage over circular ones.

Seiner, John M.↗

NASA NTRS1988

Results from NASA Langley experimental studies of multiaxis thrust vectoring nozzles

Multiaxis thrust-vectoring nozzles can furnish substantial combat performance gains for military aircraft; in pursuit of these gains, NASA-Langley has conducted an extensive experimental research program encompassing both static and freestream performance evaluations for numerous multiaxis thrust-vectoring nozzle concepts using air-powered simulation models. An evaluation is presently made of a selection of these experimental results. Two-dimensional and axisymmetric convergent-divergent configurations and hybrid configurations, with and without postexit vanes, have been studied.

Berrier, Bobby L.↗

NASA NTRS1985

Free-jet investigation of mechanically suppressed, high radius ratio coannular plug model nozzles

The experimental and analytical acoustic results of a scale-model investigation or unsuppressed and mechanically suppressed high-radius ratio coannular plug nozzles with inverted velocity and temperature profiles are summarized. Nine coannular nozzle configurations along with a reference conical nozzle were evaluated in the Anechoic Free-Jet Facility for a total of 212 acoustic test points. Most of the tests were conducted at variable cycle engine conditions applicable to advanced high speed aircraft. The tested nozzles included coannular plug nozzles with both convergent and convergent-divergent (C-D) terminations in order to evaluate C-D effectiveness in the reduction of shock-cell noise and 20 and 40 shallow-chute mechanical suppressors in the outer stream in order to evaluate their effectiveness in the reduction of jet noise. In addition to the acoustic tests, mean and turbulent velocity measurements were made on selected plumes of the 20 shallow-chute configuration using a laser velocimeter. At a mixed jet velocity of 700 m/sec, the 20 shallow-chute suppressor configuration yielded peak aft quadrant suppression of 11.5 and 9 PNdB and forward quadrant suppression of 7 and 6 PNdB relative to a baseline conical nozzles during static and simulated flight, respectively. The C-D terminations were observed to reduce shock-cell noise. An engineering spectral prediction method was formulated for mechanically suppressed coannular plug nozzles.

Janardan, B. A.↗

NASA NTRS1981

A theoretical and experimental investigation of shock-associated noise in supersonic jets

This paper examines the fundamental mechanism of broadband shock noise generation in an improperly expanded supersonic jet. The study includes circular convergent-divergent as well as circular convergent nozzles. The main source of shock noise is determined to be the transient interaction between the shock front and the convected vorticity within the jet plume. The discussion of the noise-generation mechanism is based on detailed numerical analysis, theoretical aerodynamic and acoustic modeling, refined measurements of the jet mean flow, shock-cell structure, turbulence, and noise. These results provide a broad-based generalization for the Harper-Bourne and Fisher analysis and prediction method.

Pao, S. P.↗