Over-expanded Conditions Research Articles

Acoustics from a rectangular supersonic nozzle exhausting over a flat surface.

A jet exhausting over a plate at different distances away from the nozzle was investigated to simulate jets exhausting over airframe surfaces and the jet-ground interaction during take-off and landing operations. A supersonic rectangular nozzle of 2:1 aspect ratio and 1.5 design Mach number was tested with and without the plate. Far-field acoustics from the cold and heated jets at over-expanded, design, and under-expanded conditions were measured at the reflected, sideline, and shielded azimuthal directions. When the plate starts at the nozzle exit, the "scrubbing" and "scattering" noise from the surface-jet interaction was observed at the low-end frequencies of the reflected and shielded measurements and increased as the plate approached the jet. In the sideline, the plate attached to the nozzle exit diminished the broadband shock-associated noise while the shadowgraph results showed a connection to weakening of the shock-cell structures. In the cold jet, screech was mitigated with the plate attached at the nozzle exit. When the plate was moved away from the nozzle, screech tones were intensified at the under-expanded condition. Crackle levels were significantly intensified in the sideline within a range of plate positions. Noise levels in the shielded region were considerably lower due to the shielding effect.

Measurements of unsteady pressure fluctuations in the near-field of a solid rocket motor plume

Near-plume fluctuating pressures were measured during five static burns of a two-stage solid rocket motor. An array of 11 water-cooled dynamic pressure sensors was used for the near-field survey, and a condenser microphone was used to monitor the far-field acoustic fluctuations. During the initial high-thrust phase of the burn, the plume was nearly ideally expanded, while in the following low-thrust phase, it was highly over-expanded and showed the presence of clear shock patterns. This paper presents time histories and spectra measured for the two thrust conditions. Spectra from very close to the plume show high levels of low-frequency fluctuations which are known to produce significant vibro-acoustic response of the spacecraft structures. The far-field microphone signal was dominated by mixing noise with little evidence of contribution from shock-associated noise, even for the over-expanded condition. The work was performed in support of an effort to improve predictions of the acoustic environment of a manned spacecraft, such as NASA's Orion Crew Vehicle, during pad abort scenarios.

Acoustic field of a shielded rectangular supersonic jet

The impact of a flat surface installed parallel to a supersonic rectangular jet of AR = 2 on the far field noise is studied. Far-field cold jet results at design, over-expanded and under-expanded conditions are compared between the free-field jet, the nozzle with the surface matching with the exit lip (h = 0De), and the surface at different stand-off positions away from the jet axis, h = 1, 2, 3 De. Results are shown for all jet azimuthal angles. When the surface is installed at h = 0De, broadband shock-associated noise intensity was decreased and its peak frequency shifted. At NPR = 2.5, shock noise appears to be entirely mitigated. Also, a low frequency noise component is observe below St = 0.15, and assumed to be related to the trailing edge/jet plume interaction. At NPR = 3.0, strong screech tones are mitigated with the surface installed at h = 0De. In the shielded region, noise levels are significantly lower for all plate positions. Mixing noise and Mach wave radiation are affected by the plat stand-off location at the ϕ = 90o. Heated jet measurements up TR = 3.0, near-field and high-speed shadowgraph visualization are performed and reported.

Numerical investigation of flow separation behavior in an over-expanded annular conical aerospike nozzle

A three-part numerical investigation has been conducted in order to identify the flow separation behavior—the progression of the shock structure, the flow separation pattern with an increase in the nozzle pressure ratio (NPR), the prediction of the separation data on the nozzle wall, and the influence of the gas density effect on the flow separation behavior are included. The computational results reveal that the annular conical aerospike nozzle is dominated by shock/shock and shock/boundary layer interactions at all calculated NPRs, and the shock physics and associated flow separation behavior are quite complex. An abnormal flow separation behavior as well as a transition process from no flow separation at highly over-expanded conditions to a restricted shock separation and finally to a free shock separation even at the deign condition can be observed. The complex shock physics has further influence on the separation data on both the spike and cowl walls, and separation criteria suggested by literatures developed from separation data in conical or bell-type rocket nozzles fail at the prediction of flow separation behavior in the present asymmetric supersonic nozzle. Correlation of flow separation with the gas density is distinct for highly over-expanded conditions. Decreasing the gas density or reducing mass flow results in a smaller adverse pressure gradient across the separation shock or a weaker shock system, and this is strongly coupled with the flow separation behavior. The computational results agree well with the experimental data in both shock physics and static wall pressure distribution at the specific NPRs, indicating that the computational methodology here is advisable to accurately predict the flow physics.

Experimental investigation of flow through planar double divergent nozzles

Dual bell nozzle is one of the feasible and cost effective techniques for altitude adaptation. Planar double divergent nozzle with a rectangular cross section was designed for two different NPR׳s to simulate and investigate the flow regimes similar to those inside the dual bell nozzle. Measurements involved flow visualization using Schlieren technique and wall static pressure measurements. The flow transition between the two nozzles at the respective inflection points and the formation of recirculation region due to flow separation was analyzed in detail. Cold flow tests were performed on the double divergent nozzle in the over-expanded conditions to study the shock wave characteristics. The results obtained from the two independent double divergent nozzles were compared with those obtained from a single divergent nozzle of the same area ratio. From the experiments it was observed that inflection angle played a key role in defining the type of shock structures existing inside the double divergent nozzles.

A study on the thrust vector control using a bypass flow passage

Bypass actuators are both indispensable and critical components for thrust vector control of supersonic vehicles. Detailed flow features concerned with bypass flow play a key role in thrust vectoring systems. An analytical model was created in supersonic nozzle with bypass flow passage. Meanwhile, a computational fluid dynamics-based study was conducted for studying this phenomenon. To validate the analytical and numerical models, comparisons with experimental result were performed. Good agreement was observed in overall region. Analytical and numerical pressure distributions under different injection positions were compared. The distance between separation point and injection position becomes smaller as the secondary injection position moves upstream. This is because of the reducing of bypass flow rate. The effects of expanded ratio and mass flow rate were adequately researched and tested. Significant variation of system thrust ratio and specific impulse takes place in the over-expanded conditions. It is indicated that the secondary flow produces great effects for over-expanded conditions. The separation point of boundary layer moves upstream as bypass flow rate increases. System thrust ratio and specific impulse will reduce if bypass flow rate increases.

Exit Geometry Effect on Jet Mixing

An elliptical nozzle of 3:1 aspect ratio (AR) issuing a Mach 1.5 jet has been used to investigate its mixing promoting efficiency under over-expanded and under-expanded conditions at the nozzle exit. Azimuthal asymmetry of the elliptical shape is interpreted as a benefit to the mixing process, which is significantly enhanced in comparison to an equivalent circular jet under identical conditions. The elliptical jet decays faster than its circular counterpart throughout the jet field. It is found that continuous variation in the size of the mixing promoting vortices shed from the nozzle exit, as a result of its azimuthal asymmetry, is responsible for faster decay of elliptical jet. The visualization of elliptic jet corroborates the fact that the waves prevailing in the elliptical jet are significantly weaker than those in the circular jet.

Free compressible jet investigation

The nozzle pressure ratio (NPR) effect on a supersonic turbulent jet was investigated. A dedicated convergent/divergent nozzle together with a flow feeding system was designed and manufactured. A nozzle Mach exit of M j = 1.5 was selected in order to obtain a convective Mach number of M c = 0.6. The flow was investigated for over-expanded, correctly expanded and under-expanded jet conditions. Mach number, total temperature and flow velocity measurements were carried out in order to characterise the jet behaviour. The inlet conditions of the jet flow were monitored in order to calculate the nozzle exit speed of sound and evaluate the mean Mach number distribution starting from the flow velocity data. A detailed analysis of the Mach results obtained by a static Pitot probe and by a particle image velocimetry measurement system was carried out. The mean flow velocity was investigated, and the axial Mach decay and the spreading rate were associated with the flow structures and with the compressibility effects. Aerodynamics of the different jet conditions was evaluated, and the shock cells structures were detected and discussed correlating the jet structure to the flow fluctuation and local turbulence. The longitudinal and radial distribution of the total temperature was investigated, and the temperature profiles were analysed and discussed. The total temperature behaviour was correlated to the turbulent phenomena and to the NPR jet conditions. Self-similarity condition was encountered and discussed for the over-expanded jet. Compressibility effects on the local turbulence, on the turbulent kinetic energy and on the Reynolds tensor were discussed.

Millisecond-scale shock-train evolution in high-pressure ratio nozzles: Schlieren imaging and qualitative analysis of shock–boundary layer interaction

The classical picture of shock evolution in nozzles holds that under over-expanded flow conditions, a single, nominally normal shock exists within the nozzle. Focusing on the highly dynamic flow produced during blow-down of an experimental, high-nozzle pressure ratio, planar nozzle, this article presents visual evidence that shock-trains – here, a pair of parallel, nominally normal shocks – dominate the rapidly evolving flow field. Three principal results are presented in this study. First, high-speed schlieren images of the evolving nozzle flow are reported. Second, a simple qualitative model of shock–boundary layer/recirculation zone interaction is proposed and used to explain observed millisecond-scale shock-train structure. Third, limited wall pressure measurements and schlieren images are combined to propose a second qualitative model of shock-train–boundary layer/recirculation zone evolution on the longer blow-down process time-scale. The results provide insight into millisecond-scale compressible flow dynamics within high-nozzle pressure ratios .

Acoustic assessment of small-scale military-style nozzles with chevrons

With the emergence of more powerful fighter aircraft, supersonic jet noise reduction devices are being intensively researched. Small-scale measurements are a crucial step in evaluating the potential of noise reduction concepts at an early stage in the design process. With this in mind, the present study provides an acoustic assessment of small-scale military-style nozzles with chevrons. Comparisons are made between the present measurements and those made by NASA at moderate scale. Measurements made with baseline nozzles (without chevrons) show excellent agreement with NASA data establishing the accuracy of the scaling methodology. The effect of chevrons on supersonic jets is then investigated for cold jets, highlighting the crucial role of the jet operating conditions on the effects of chevrons on the jet flow and subsequent acoustic benefits. At low Reynolds numbers (small scale) the penetration of the chevrons in the jet flow is the most important chevron parameter in reducing the generated noise. A small-scale heat simulated jet is investigated in the over-expanded condition and shows no substantial noise reduction from the chevrons. This is contradictory to moderate-scale measurements. The discrepancy is attributed to a Reynolds number low enough to sustain an annular laminar boundary layer in the nozzle that separates in the over-expanded flow condition. Transition of the boundary layer to turbulent flow is induced with inner roughness of the nozzle and results in more noise reduction with the chevrons. The resulting effect is comparable to results from NASA, seemingly validating the hypothesis made that jets of too low Reynolds number cannot be used directly to accurately measure chevron noise reduction. These results are important in assessing the limitations of small-scale measurements in this particular jet noise reduction method. © 2012 Institute of Noise Control Engineering.

Color Schlieren imaging of high-pressure overexpanded planar nozzle flow using a simple, low-cost test apparatus

Complex flow features within rocket nozzles can exert significant influence on both the dynamics and safety of rockets during flight. Specifically, under over-expanded flow conditions, during, low altitude flight, random, often large side loads can appear within nozzles. While significant research has focused on this classical problem, due to the high nozzle pressure ratios (NPR) extant across rocket nozzles, most experimental work: (1) has focused on measuring wall pressure distributions under conditions when side loads appear, (2) has been carried out in large government or industrial test facilities, and (3) has only provided limited, though crucially important, visualization data. This short paper describes the construction and operation of a very simple, low cost test apparatus that allows imaging of flow features within planar nozzles, under the high NPR conditions characteristic of medium-to-large rockets. Representative color Schlieren images of flow shock structure obtained within the test apparatus are also presented and briefly described.

Discrete tone emission from high-pressure ratio supersonic jets fromconvergent-divergent nozzles

The results of a test program to investigate the discrete tone generation from free axisymmetric supersonic jets issuing from convergent-divergent conical nozzles of different sizes and expansion ratios in a range of stagnation pressure ratio never reported before are presented. In both highly overexpanded and underexpanded jet conditions, the second shock-cell length was found to be in good agreement with Pack's modified theoretical formula for the first shock-cell length of slightly incorrectly expanded supersonic jets, when the nozzle diameter was replaced by the jet diameter dj corresponding to the fully expanded jet Mach number M-. The use of dj in normalizing shock-associated noise data was first proposed by Tarn and Tanna. To a good approximation, the second shock-cell length displayed a linear dependence on the parameter p2 = M.J — rather than on (J. The apparent length of the first shock cell was also found to display linear dependence on P 2 but only in the overexpanded jet condition. The linear variation of the second shock-cell length with p2 was found to be invariant for all the nozzles tested, but in the case of the first shock cell, the proportionality constant varied with the nozzle expansion ratio. The discrete tone fundamental frequency was found not to depend on the nozzle exit diameter but to be inversely proportional to the nozzle throat diameter. With a constant ratio between the disturbance convection velocity and the fully expanded jet velocity, the combination of Powell's simplified feedback loop equation with the shock-cell length obtained from the modified theoretical formula predicted to a good practical accuracy the discrete tone fundamental frequency.

Thrust vector control utilizing asymmetric jet nozzles

R interest in thrust vector control with rigidly mounted thrusters has been revived. The results described in this Note are associated with the side-force generated by a supersonic jet with an exit plane which is oblique to the centerline of the exhaust flow. As shown in Fig. 1, depending on whether the flow is overor under-expanded, the direction of the exhaust flow is deflected either toward or away from the normal to the exit plane by an amount related to the pressure imbalance and the projected area of the exit plane in the lateral direction. Therefore, control over the direction and amount of side-force is available by varying the stagnation pressure and exit plane obliquity. Although the present results are associated with a twodimensional configuration, with the exit plane obliquity established by extending one wall of the nozzle, application to an axisymmetric configuration is conceivable. For example, consider a fixed axisymmetric thruster enclosed by a cylindrical housing which is split longitudinally into various sections (see Fig. 1). At the null position, the end of the cylinder coincides with the exit plane of the thruster. As various combinations of cylindrical sections are translated forward of the exit plane, the radially projected area of the exit plane may be oriented toward any azimuthal angle. The amount of projected area and its orientation may also be changed rapidly, which adds another control feature to the thrust vectoring. The focus of the work described in Ref. 2 involved the extent to which viscous effects affected the performance, particularly at over-expanded conditions. A two-dimensional supersonic jet of aspect ratio (height/width) equal to 1.5 was established over the Mach number range 1.0 to 3.0. One wall of the nozzle contour was then extended a specified fraction of the nozzle height to establish an oblique exit plane. Transparent and parallel lateral sidewalls bounded the nozzle and extensions thereof through which the flow angle could be observed. The side force generated was measured with a strain gage balance to an accuracy of ±2% of the reading, based on the scatter in repeated experiments and comparison to the measured flow angle. In order to assess the degree to which viscous effects impede accurate analytical prediction, the amount of asymmetry was limited to maintain the wall extension within the domain established by the Mach wave originating from the opposite nozzle wall at the exit plane. Figures 2-4 depict the forces measured on three nozzles whose exit plane is offset by an amount 1.0, 4/3, and 5/3 times the jet height, respectively. Comparison with predictions of two-dimensional inviscid flow theory is shown also. These results may be converted into force coefficients by dividing the results by the surface area of the wall extension and the exit dynamic pressure. The corresponding plot of the Fig. 3 data is presented in Fig. 5. These results indicate that, even in the simple case established here, nontrivial deviations between inviscid flow theory and experiment are observed, although the general trends predicted by analysis are correct. It is also clear that the vectoring control may extend well into the overexpanded region before the nozzle boundary layers begin to