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Static tests of nozzle efficiency and pumping characteristics of conical turbojet plug nozzle with translating primary and secondary shrouds at Mach 0 to 2.0
Six coannular plug nozzle configurations having inverted velocity and temperature profiles, and a baseline convergent conical nozzle were tested for simulated flight acoustic evaluation in General Electric's Anechoic Free-Jet Acoustic Facility. The nozzles were tested over a range of test conditions that are typical of a Variable Cycle Engine for application to advanced high speed aircraft. The outer stream radius ratio for most of the configurations was 0.853, and the inner-stream-outer-stream area ratio was tested in the range of 0.54. Other variables investigated were the influence of bypass struts, a simple noncontoured convergent-divergent outer stream nozzle for forward quadrant shock noise control, and the effects of varying outer stream radius and inner-stream-to-outer-stream velocity ratios on the flight noise signatures of the nozzles. It was found that in simulated flight, the high-radius-ratio coannular plug nozzles maintain their jet noise and shock noise reduction features previously observed in static testing. The presence of nozzle bypass structs will not significantly effect the acoustic noise reduction features of a General Electric-type nozzle design. A unique coannular plug nozzle flight acoustic spectral prediction method was identified and found to predict the measured results quite well. Special laser velocimeter and acoustic measurements were performed which have given new insight into the jet and shock noise reduction mechanisms of coannular plug nozzles with regard to identifying further beneficial research efforts.
Outdoor static and 40 x 80 FT wind tunnel tests of the J79-15 engine/nacelle system with the conic nozzle and 32-chute exhaust suppressor were conducted to acquire the data necessary to evaluate the simulated in-flight signature of an engine-size 32-chute exhaust nozzle suppressor using the 40 x 80 ft wind tunnel and to study possible engine core noise contamination of the jet signature. The tests are described and and a sampling of the data acquired is presented. Included are aero performance summaries, as-measured and composite 1/3 OBSPL spectra for the 70 ft sideline high and low mics from the outdoor static tests, sideline traverse spectra and internal noise measurements from both the outdoor static and the 40 x 80 ft wind tunnel tests.
Boundary layer and heat transfer measurements for turbulent boundary layer laminarizing in conical nozzle flow with wall cooling
Planar laser-induced fluorescence visualisation is used to investigate nonuniformities in the flow of a hypersonic conical nozzle. Possible causes for the nonuniformity are outlined and investigated, and the problem is shown to be due to a small step at the nozzle throat. Entrainment of cold boundary layer gas is postulated as the cause of the signal nonuniformity.
A computational jet noise study of a short conical plug-nozzle (CPN) is presented. The CPN has an exit diameter of 45 mm and the geometrical configuration closely approximates that of an ideal contoured plug-nozzle having shockless flow at pressure ratio xi(sub d) = 3.62. The gasdynamics of the jet flows have been predicted using the CFD code, NPARC with k-epsilon turbulence model; these data are then used for noise computations based on the modified GE/MGB code. The study covers a range of pressure ratio, 2.0 less than or equal to xi less than or equal to 5.0. The agreement of the computational results with the available experimental data is favorable. The results indicate consistent noise reduction effectiveness of the CPN as compared to equivalent convergent, convergent-divergent and ideal contoured plug nozzles at all pressure ratios. At design pressure ratio, codes predict noise levels within 4.0 dB of the measurements; and at off-design pressure ratios, in general, within 5.0 dB except at very high frequencies when deviations up to 10 dB are noted. The shock formation mechanism in the CPN jet is noted to be basically different from those in the convergent and CD nozzle jets.
Distributions of nitric oxide vibrational temperature, rotational temperature and velocity have been measured in the hypersonic freestream at the exit of a conical nozzle, using planar laser-induced fluorescence. Particular attention has been devoted to reducing the major sources of systematic error that can affect fluorescence tempera- ture measurements, including beam attenuation, transition saturation effects, laser mode fluctuations and transition choice. Visualization experiments have been performed to improve the uniformity of the nozzle flow. Comparisons of measured quantities with a simple one-dimensional computation are made, showing good agreement between measurements and theory given the uncertainty of the nozzle reservoir conditions and the vibrational relaxation rate.
Ground-based model scale aeroacoustic data is frequently used to predict the results of flight tests while saving time and money. The value of a model scale test is therefore dependent on how well the data can be transformed to the full scale conditions. In the spring of 2000, a model scale test was conducted to prove the value of chevron nozzles as a noise reduction device for turbojet applications. The chevron nozzle reduced noise by 2 EPNdB at an engine pressure ratio of 2.3 compared to that of the standard conic nozzle. This result led to a full scale flyover test in the spring of 2001 to verify these results. The flyover test confirmed the 2 EPNdB reduction predicted by the model scale test one year earlier. However, further analysis of the data revealed that the spectra and directivity, both on an OASPL and PNL basis, do not agree in either shape or absolute level. This paper explores these differences in an effort to improve the data transformation from model scale to full scale.
Experimental investigations were conducted to evaluate the effects of a contoured nozzle and modified cathode shape on ammonia arcjet engine performance. The contoured nozzle performance data were compared to the performance data of an arcjet which had a 38-deg included-angle, conical nozzle. Thrust improvements of up to 10 percent were demonstrated which corresponded to 3 percent improvements in specific impulse and 10 percent improvements in thrust efficiency. Performance characterizations for the modified cathode tip were conducted with the contoured nozzle arcjet. A uniform 15 percent decrease in arc voltage was demonstrated over a mass flow range of 0.175 to 0.350 g/s. A 4 percent improvement in thrust efficiency was noted at 22.0 kW.
Bell Nozzle Kernel Analysis Program computes and analyzes the supersonic flowfield in the kernel, or initial expansion region, of a bell or conical nozzle. It analyzes both plane and axisymmetric geometrices for specified gas properties, nozzle throat geometry and input line.
Computer program calculates the inviscid one-dimensional equilibrium, frozen, and nonequilibrium nozzle expansion of gaseous propellant exhaust mixtures containing the elements - carbon, hydrogen, oxygen, nitrogen, fluorine and chlorine. The program performs calculations for conical nozzles only.
An improved acoustic impulse technique was developed and was used to study the transmission characteristics of duct/nozzle systems. To accomplish the above objective, various problems associated with the existing spark-discharge impulse technique were first studied. These included (1) the nonlinear behavior of high intensity pulses, (2) the contamination of the signal with flow noise, (3) low signal-to-noise ratio at high exhaust velocities, and (4) the inability to control or shape the signal generated by the source, specially when multiple spark points were used as the source. The first step to resolve these problems was the replacement of the spark-discharge source with electroacoustic driver(s). These included (1) synthesizing on acoustic impulse with acoustic driver(s) to control and shape the output signal, (2) time domain signal averaging to remove flow noise from the contaminated signal, (3) signal editing to remove unwanted portions of the time history, (4) spectral averaging, and (5) numerical smoothing. The acoustic power measurement technique was improved by taking multiple induct measurements and by a modal decomposition process to account for the contribution of higher order modes in the power computation. The improved acoustic impulse technique was then validated by comparing the results derived by an impedance tube method. The mechanism of acoustic power loss, that occurs when sound is transmitted through nozzle terminations, was investigated. Finally, the refined impulse technique was applied to obtain more accurate results for the acoustic transmission characteristics of a conical nozzle and a multi-lobe multi-tube supressor nozzle.
Aspects are examined of an investigation by the Marshall Space Flight Center into the conceptual feasibility of increasing the thrust performance of the Space Shuttle Main Engine (SSME) by using a conical nozzle extension fitted with an ablative insert in order to achieve a low-cost, near-term gain in payload. The ablating insert would provide a controlled increase in nozzle expansion ratio during launch and early climbout (first 30-60 seconds) so as to reduce thrust loss from nozzle over-expansion in the lower atmosphere. Summaries are given of JPL studies in the area of: defining the near-wall flow environment in the extended nozzle insert region; selecting potential insert materials; conceptualizing an extension/insert geometrical configuration; and identifying future experimental efforts necessary to verify the feasibility of the concepts.
Aerodynamic performance and jet noise characteristics of a one sixth scale model of the variable cycle engine testbed exhaust system were obtained in a series of static tests over a range of simulated engine operating conditions. Model acoustic data were acquired. Data were compared to predictions of coannular model nozzle performance. The model, tested with an without a hardwall ejector, had a total flow area equivalent to a 0.127 meter (5 inch) diameter conical nozzle with a 0.65 fan to primary nozzle area ratio and a 0.82 fan nozzle radius ratio. Fan stream temperatures and velocities were varied from 422 K to 1089 K (760 R to 1960 R) and 434 to 755 meters per second (1423 to 2477 feet per second). Primary stream properties were varied from 589 to 1089 K (1060 R to 1960 R) and 353 to 600 meters per second (1158 to 1968 feet per second). Exhaust plume velocity surveys were conducted at one operating condition with and without the ejector installed. Thirty aerodynamic performance data points were obtained with an unheated air supply. Fan nozzle pressure ratio was varied from 1.8 to 3.2 at a constant primary pressure ratio of 1.6; primary pressure ratio was varied from 1.4 to 2.4 while holding fan pressure ratio constant at 2.4. Operation with the ejector increased nozzle thrust coefficient 0.2 to 0.4 percent.
Computer program calculates the inviscid one-dimensional equilibrium, frozen, and nonequilibrium nozzle expansion of propellant exhaust mixtures containing carbon, hydrogen, oxygen, nitrogen, fluorine, chlorine and either aluminum, beryllium, boron or lithium. This program performs calculations for conical nozzles only.
Computational simulations and analysis of flow characterization tests in a high enthalpy arc-jet facility at NASA Ames Research Center are reported. These tests were conducted in the 10-MW TP3 facility 15-inch conical nozzle, and pitot pressure and heat flux probes are used to survey test flow downstream of the nozzle exit. Two 1.59-cm (0.625 in) diameter hemisphere probes were used: one with a stagnation-point pressure port, the other with a Gardon heat flux gage. Calibration data were also obtained using 10.16-cm iso-q and 20.32-cm flat-faced slug calorimeters. Experimental surveys of arc-jet test flow with pitot pressure and heat flux probes were obtained at eight arc-heater conditions, covering wide ranges in arc current (316–1756 A) and mass flow rate (32–500 g/s, without any cold-gas injection at the arc-heater plenum) and providing assessment of the flow uniformity in the facility. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested. These simulations take into account non-uniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields. Comparisons of computations with the experimental measurements are presented, showing reasonably good agreement. The probe survey data and accompanying analysis show that the test flow in the TP3 15-inch nozzle is highly non-uniform at most conditions, and the extent of flow non-uniformity depends on arc current and mass flow rate.
Computational simulations and analysis of flow characterization tests in a high enthalpy arc-jet facility at NASA Ames Research Center are reported. These tests were conducted in the 10-MW TP3 facility 15-inch conical nozzle, and pitot pressure and heat flux probes are used to survey test flow downstream of the nozzle exit. Two 1.59-cm (0.625 in) diameter hemisphere probes were used: one with a stagnation-point pressure port, the other with a Gardon heat flux gage. Calibration data were also obtained using 20.32-cm flat-faced slug calorimeters. Experimental surveys of arc-jet test flow with pitot pressure and heat flux probes were obtained at eight arc-heater conditions, covering wide ranges in arc current (316–1756 A) and mass flow rate (32–500 g/s, without any cold-gas injection at the arc-heater plenum) and providing assessment of the flow uniformity in the facility. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested. These simulations take into account non-uniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields. Comparisons of computations with the experimental measurements are presented, showing reasonably good agreement. The probe survey data and accompanying analysis show that the test flow in the TP3 15-inch nozzle is highly non-uniform at most conditions, and the extent of flow non-uniformity depends on arc current and mass flow rate.
Computer program calculates the inviscid axisymmetric nozzle expansion of propellant systems having both gaseous and condensed exhaust products. The program uses velocity and thermal lags and will perform calculations for contoured and conical nozzles.