Total Pressure Recovery Coefficient Research Articles

Redesign of an 11 cm-diameter Micro Diffuser

In order to improve the performances of an 11 cm-diameter turbine engine, this article suggests to substitute a new-style micro diffuser redesigned based on a new concept for the traditional diffuser having poor performances. The new diffuser comprises integral blades and splitters, which are taken for a series of ducts in designing. This article investigates the effects of the cross-section area distribution along the flow path on the redesigned diffuser's performances. Having furnished with the new diffuser in place of the original vaned one, the 11 cm-diameter prototype engine is tested on the rig for its performances. CFD and experiments have shown that the improved diffuser with the unchanged original size has gained excellent performance parameters of pressure coefficient over 0.65 and total pressure recovery coefficient over 0.9. Equipped with the redesigned micro diffuser, the engine increases the thrust by 11% and decreases the specific fuel consumption by 9%.

Control of the corner separation in a compressor cascade by steady and unsteady plasma aerodynamic actuation

This paper reports experimental results on using steady and unsteady plasma aerodynamic actuation to control the corner separation, which forms over the suction surface and end wall corner of a compressor cascade blade passage. Total pressure recovery coefficient distribution was adopted to evaluate the corner separation. Corner separation causes significant total pressure loss even when the angle of attack is 0°. Both steady and unsteady plasma aerodynamic actuations suppress the corner separation effectively. The control effect obtained by the electrode pair at 25% chord length is as effective as that obtained by all four electrode pairs. Increasing the applied voltage improves the control effect while it augments the power requirement. Increasing the Reynolds number or the angle of attack makes the corner separation more difficult to control. The unsteady actuation is much more effective and requires less power due to the coupling between the unsteady actuation and the separated flow. Duty cycle and excitation frequency are key parameters in unsteady plasma flow control. There are thresholds in both the duty cycle and the excitation frequency, above which the control effect saturates. The maximum relative reduction in total pressure loss coefficient achieved is up to 28% at 70% blade span. The obvious difference between steady and unsteady actuation may be that wall jet governs the flow control effect of steady actuation, while much more vortex induced by unsteady actuation is the reason for better control effect.

Effects of boundary-layers bleeding on unstart/restart characteristics of hypersonic inlets

AbstractA series of mixed-compression hypersonic inlets at different bleeding rates were simulated at different freestream conditions in this paper. The unstart/restart characteristics of hypersonic inlets were analysed and the reasons why the unstart/restart phenomenon is in existence is presented. The unstart/restart characteristics of hypersonic inlets at different bleeding rates were given. The effects of boundary-layer bleeding on the performance parameter (mass-captured coefficient, total-pressure recovery coefficient), starting and restarting Mach number of hypersonic inlets were discussed. In conclusion, boundary-layer bleeding can improve the performance parameter of hypersonic inlets, and can reduce the starting and restarting Mach number, and can broad the operation range of the hypersonic inlet.

Effects of wall cooling on performance parameters of hypersonic inlets

The internal flowfield of a 2-D mixed-compression hypersonic inlet was simulated numerically at different wall temperatures. The effects of wall cooling on performance parameters of the hypersonic inlet without and with the backpressure, and especially on the maximum backpressure of a fixed-geometry hypersonic inlet were discussed. The inner physical mechanism of wall cooling improving the maximum backpressure of hypersonic inlets was analyzed. In contrast with no wall cooling, the static pressure ratio of hypersonic inlets without backpressure is reduced slightly, but the mass-captured coefficient, total-pressure recovery coefficient and the flow uniformity of hypersonic inlets at the isolator exit are improved by the action of wall cooling. The interaction between boundary layers and shocks is weakened due to wall cooling, which leads to that the boundary layers separations at the entrance of the isolator caused by the high backpressure occur later, and it can improve the maximum backpressure ratio of hypersonic inlets. With the wall temperatures decreasing, the maximum backpressure ratio and mass-captured coefficient are added, and the total-pressure recovery coefficient of hypersonic inlets is reduced.

Experimental study of different control methods for hypersonic air inlets

An experimental study of different control methods for hypersonic air inlets aimed at ensuring reliable starting of these apparatuses and improving their operating characteristics in the range of Mach numbers 2 to 8 is reported. Conditions for boundary-layer separation and possibilities for preventing this separation by using modified diffuser configurations and/or perforation bleedage are examined. An air-inlet model was tested for operation in an intermittent wind tunnel and in a blow-down wind tunnel respectively in the Mach-number ranges 2 to 6 and 5 to 8. Distributions of static and total air pressures on the walls of the model and in several cross sections were measured, together with air flow coefficients and total-pressure recovery coefficients. Perforation bleedage is shown to offer an efficient means to facilitate air-inlet starting. Perforation bleed has enabled a more than two-fold increase in the air flow coefficient on the model with sidewalls. A perforation-bleed panel installed closer to the air-inlet throat proved to be more efficient. The possibility of sudden starting of the air-inlet apparatus was checked in the intermittent wind tunnel; it was shown that, here, sudden starting could be realized. The data obtained in the intermittent wind tunnel proved to be consistent with data obtained in the blow-down wind tunnel with up to 150-ms blowdown time.

Forebody compressibility research of hypersonic vehicle

Three kinds of forebody model of hypersonic vehicles were studied with numerical simulation method. It shows that the two-order compressive ramp model is the best selection among the three for its good evaluative parameters value at the cowl of the inlet. This model can provide higher value of flux coefficient and total pressure recovery coefficient and lower average Mach number compared with those of the other two models. Simultaneously different compressive angles may have different effects. The configuration which the first-order of compressive angle is 4° and the second 5° is the optimum combination. Furthermore factors such as attack angle were concerned. Better result may be obtained with a range of attack angles. Based on the work above the integrated design for forebody/inlet of a hypersonic vehicle was performed. The numerical result shows that this integrated model provides good flow field quality for inlet and engine work.

Numerical and experimental studies of 3D hypersonic inlet

The results of numerical and experimental studies of a new configuration of 3D hypersonic inlet with the minimum throat area, which was called a convergent inlet, are presented in this paper. It is shown that the use of this inlet configuration allows one to reduce the drag and thermal protection of surfaces of a hypersonic engine within the entire range of flight velocities. The calculations were performed within the framework of inviscid gas model by the method of finite volumes. The flow and inlet characteristics, taking account of viscosity, were also calculated using the boundary layer equations. The experimental studies were performed within the Mach number range from 2 to 10.7 and Reynolds number based on the model inlet height of Re=1–5×106. The results included the flow parameters on the external compression surface and in the inlet duct, the Mach number in the throat, the air flow rate, the total pressure recovery coefficient, the inlet drag, and the boundary layer characteristics on compression surfaces were determined, including the skin friction coefficients. These results are also compared with the data for traditional 2D inlets.