Pressure Recovery Coefficient Research Articles

Experimental and numerical investigation on a supersonic inlet with large bleed window

AbstractThe design of a two-dimensional supersonic inlet with large bleed window at low Mach number was developed. Numerical simulation and wind tunnel experiments were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the single-degree-of-freedom variable geometry scheme adopted in this paper guarantees the steady work of the inlet over a wide speed range. The large bleed window caused by rotation of the compression ramp appears near the throat at low Mach number. Low-pressure airflow near the bleed window neutralises the original high-pressure airflow behind the shock train, which decreases the overall pressure of the downstream region of the internal contraction section. To match the lower pressure, the structure of the shock train changes from strong $\lambda$-type to weak $\lambda$-type, and finally to a normal shock wave as backpressure increases at Mach number 2.5. Herein, the total pressure recovery coefficient of the inlet near the critical condition improves by 8.5% as the backpressure ratio (Pe/P0) adds from 13 to 14.6 at Mach number 2.5. It proves that the scheme is effective on terminal shock wave control and inlet performance improvement. In addition, due to the background wave and the bleed window, two kinds of shock wave oscillation occur when the backpressure ratio is 13.1.

Direct-Connect Test of Solid Scramjet with Symmetrical Structure

The solid scramjet has become one of the most promising engine types. In this paper, we report the first direct-connect test of a solid scramjet with symmetrical structure, carried out using boron-based fuel-rich solid propellant as fuel. During the test, which simulated a flight environment at Mach 5.6 and 25 km, the performance of the solid scramjet was obtained by measuring the pressure, thrust, and mass flow. The results show that, due to the change in the combustion area of the propellant and the deposition of the throat in the gas generator during the test, the equivalence ratio gradually increased from 0.54 to 0.63. In a solid scramjet, it is possible to obtain a symmetrical distribution of the flow field within the combustor. Moreover, in a multi-cavity combustor, the combustion state expands from the cavity to the center of the flow channel. The performance of the solid scramjet increased during the test, reaching a combustion efficiency of about 42%, a total pressure recovery coefficient of 0.35, and a thrust gain specific impulse of about 418 s. The solid scramjet with symmetrical structure is feasible. The cavity configuration adopted in this paper can reduce the ignition delay time of fuel-rich gas and improve the combustion efficiency of gas-phase combustible components. The shock trains in the isolator are conducive to the recovery of the total pressure. The performance of the solid scramjet is limited by the low combustion efficiency of the particles.

Optimization of the aerodynamic characteristics of a NACA air intake based on DoE and numerical methods

The Design of Experiment (DoE) method combined with numerical simulations is employed to optimize the aerodynamic characteristics of a National Advisory Committee for Aeronautics (NACA) air intake. The drag force, total pressure recovery coefficient, mass flow rate at the outlet, and Mach number at the outlet are used to evaluate the performance of the air intake based on computational fluid dynamics (CFD) numerical methods. Three critical geometric parameters, namely the ramp angle a, ramp divergence angle θ, and turning corner radius R, are considered as variables in the DoE design. Results under the baseline conditions indicated that axial vortexes are formed near the side edge close to the ramp wall. DoE analysis indicated that above four performance parameters are largely influenced by the ramp angle a, while the ramp divergence angle θ has negligible effects. A small radius of the turning corner is beneficial to enhance the pressure recovery coefficient and the mass flow rate. The Mach number at the outlet declines and the velocity uniformity on a certain section increases with decreasing value of ramp angle a. Meanwhile, larger total pressure recovery coefficient, mass flow rate, and vortex tube length are acquired under a smaller value of a. Thus, the aerodynamic performance of this air intake can be improved by a set of optimized geometric parameters.

Обґрунтування моделі турбулентної в’язкості для дослідження характеристик співвісного гвинтовентилятора і вхідного пристрою ГТД

The issues of substantiation of the most rational, based on adequacy, model of turbulent viscosity for mathematical modeling of the flow near the propfan and in the inlet of the turbine-propeller engine are considered. It was found that at present there is no universal turbulence model for determining the parameters of the boundary layer, energy loss in the flow, and laminar-turbulent transition. Analysis of the results of previous studies showed that there is a need to select and justify a turbulent viscosity model for each type of research object. The task of modeling the flow near the propfan and in the inlet device of the power plant was performed using the ANSYS CFX software product, which allows using various standard mathematical models and tools for modeling turbulent flow. The object of research is an annular axial inlet device, in front of which there is a coaxial propfan with two rows of propellers: the first row has eight blades, the second - six. 7 types of models of turbulent viscosity, which most fully describe the phenomena in the flow around the propfan and the inlet device, have been investigated: k-ωmodel; SSТ (shear stress transport) SST Transitional №1 Fully turbulence; SST Transitional №2 Specified Intermittency; SST Transitional №3 Gamma model; SST Transitional №4 Gamma theta model; SST Transitional №5 Intermittency. The results of mathematical modeling of the flow near the propfan and in the inlet device at the corresponding operating mode of the turbopropfan engine using the selected models of turbulent viscosity, the total pressure value in front of and behind the inlet device was obtained to determine the total pressure recovery coefficient in it and the value of the propfan thrust. The value of the recovery factor of the total pressure in the inlet device and the propfan thrust are compared with the flight test data of the prototype. An analysis of the comparison of the values of the total pressure recovery factor in the inlet device and the propfan thrust showed that the use of the SST Transitional №4 Gamma theta model allows obtaining the value of the total pressure recovery factor in the inlet device and the propfan thrust that is closest to the flight test results.

Numerical Investigation into the Effects of Design Parameters on the Flow Characteristics in a Turbine Exhaust Diffuser

In this study, we numerically investigated the effects of design parameters, such as the strut geometry or diffusion angle, on the performance of an industrial turbine exhaust diffuser. Turbine exhaust diffusers are commonly used to change the kinetic energy of exhaust gases from the outlet of turbine stages into the static pressure. The turbine exhaust diffuser investigated in this work consisted of an annular diffuser with five identical struts equally spaced around the front circumference and a conical diffuser with a hub extension at the rear. Four design parameters were considered and several values for each parameter were tested in this study. The aerodynamic performances of the studied diffusers were evaluated according to their pressure recovery coefficients and rates of total pressure loss. Contours for the velocity, pressure, and entropy increase were plotted and compared for the various diffuser shapes. The numerical results showed that the strut thickness and the axially swept angle of the strut significantly influence the aerodynamic performance of the turbine exhaust diffuser, whereas the strut lean angle and the diffuser hade angle are less important.

Thermodynamic optimization of the indirect precooled engine cycle using the method of cascade utilization of cold sources

To solve the problem of engine performance degradation caused by excessive fuel consumption in the precooled engine cycle, a new optimization method based on the cascade utilization of cold source is proposed, in which the cold sources with different temperature ranges and different working fluids are reasonably matched and adequately utilized. Based on this method, two new optimization cycles are put forward, i.e., the air reheat precooling cycle (ARPC) and the hydrogen reheat precooling cycle (HRPC). To evaluate the performances of these two cycles, a unified model is established. According to the simulation results, the fuel consumption of the HRPC could be reduced by 0–21.04% compared with the traditional precooled engine. Besides, the specific impulse of the HRPC, which can be raised by 16.78% when the heat transfer effectiveness is set to 0.9, is higher than that of the ARPC, which can be increased by 12.64% at most. Moreover, considering the operating constraints, the reduction of the minimum temperature difference of the regenerator and the increase of the pressure recovery coefficient of intake are effective ways to expand the performance boundaries. The results in this paper are beneficial for the performance optimization of the precooled engine cycle.

Design and analysis of a supersonic axisymmetric inlet based on controllable bleed slots

Aiming at the problem of starting and inlet-engine matching of axisymmetric variable geometry inlet with complex structure, a novel design concept of axisymmetric inlet is proposed based on controllable bleed slots switch. The laws of the typical geometric parameters of multi-slots effecting on inlet start, supercritical and subcritical performances are numerically studied. Accordingly, a three-dimensional aerodynamic design scheme of the inlet operating at Ma1.5-4 and control laws of the bleed slots are gained, which is slightly altered inlet surface and structure is simplified. The aims of minimum starting Mach number to 2 and inlet-engine matching inlet are achieved by controlling switch of bleeding slots, and inlet operates well at high total pressure recovery and low exit Mach number. The results show that: at Ma=2, the inlet starts with 30% bleeding and 16.5% total pressure recovery coefficient improving; At Ma=3, mass flow of inlet-engine matches with 23% subsonic bleeding but cost of total pressure loss about 2.6%.

Investigation on Supersonic Flow Control Using Nanosecond Dielectric Barrier Discharge Plasma Actuators

In this paper, the effects of streamwise Nanosecond Dielectric Barrier Discharge (NS-DBD) actuators on Shock Wave/Boundary Layer Interaction (SWBLI) are investigated in a Mach 2.5 supersonic flow. In this regard, the numerical investigation of NS-DBD plasma actuator effects on unsteady supersonic flow passing a 14° shock wave generator is performed using simulation of Navier-Stokes equations for 3D-flow, unsteady, compressible, and k ‐ ω SST turbulent model. In order to evaluate plasma discharge capabilities, the effects of plasma discharge length on the flow behavior are studied by investigating the flow friction factor, the region of separation bubble formation, velocity, and temperature distribution fields in the SWBLI region. The numerical results showed that plasma discharge increased the temperature of the discharge region and boundary layer temperature in the vicinity of flow separation and consequently reduced the Mach number in the plasma discharge region. Plasma excitation to the separation bubbles shifted the separation region to the upstream around 6 mm, increased SWBLI height, and increased the angle of the separation shock wave. Besides, the investigations on the variations of pressure recovery coefficient illustrated that plasma discharge to the separation bubbles had no impressive effect and decreased pressure recovery coefficient. The numerical results showed that although the NS-DBD plasma actuator was not effective in reducing the separation area in SWBLI, they were capable of shifting the separation shock position upstream. This feature can be used to modify the structure of the shock wave in supersonic intakes in off-design conditions.

A low-loss vector exhaust guide vanes and its application

The exhaust device of gas-driven fan propulsion system (for a VTOL aircraft) adopts vector exhaust guide vane (VEGV) to achieve vector thrust in the range of 0-90°, which requires that the deflection angle of the VEGV reach about ±45∘ and the total pressure recovery coefficient above 0.985. Therefore, the exhaust device needs a wide-range and low-loss VEGV to ensure efficient operation of the propulsion system. And the key of this VEGV is to eliminate the separation of the suction side under large deflection angle. So, this paper designed a new type of VEGV that can meet the demands by reducing the inverse pressure gradient of the suction surface (one of the necessary conditions for two-dimensional separation). In order to realize this design process, the spectral method of small disturbance equation (SMSDE) for two-dimensional subsonic flow was developed. Then, two VEGVs were designed by the SMSDE and verified by numerical simulation with Reynolds number 106 and blade solidity 1.18. The results show that the two VEGVs eliminate the separation of suction side at inlet Mach number of 0.25 and blade stagger angle of 45°. And the VEGV with lateral blade guarantees this characteristic up to Mach 0.3. Besides, when the inlet Mach number is below 0.3, the VEGV with lateral blade can ensure that the total pressure recovery coefficient is greater than 0.985, the outlet area ratio is greater than 0.95 and the exhaust angle error is less than 3.5°. Finally, Experimental verification of the vector exhaust device using the VEGV with lateral blade was carried out. The results show that the new VEGV has a higher total pressure recovery coefficient and outlet area ratio than the traditional VIGV at large deflection angles.

Optimization Design and Experimental Study of S-shaped Inlet

Lip and diffuser are two important parts of S-shaped inlet, which are difficult to balance in the optimization design of S-shaped inlet. This paper presents an optimization method for S-shaped inlet. Firstly, the parametric models of the lip and the diffuser are established. Combining with automatic mesh generation technology, numerical simulation and genetic algorithm, an optimization platform of s-shaped inlet design based on Isight is established. Aiming at the design of inlet for an unmanned aerial vehicle (UAV), compared with the prototype, the total pressure recovery coefficient of the optimized inlet obtained by this optimization method is increased by 1.5%, and the steady-state circumferential distortion is reduced by 58.6%. The wind tunnel experiment is carried out to obtain the aerodynamic performance of the inlet, and the experimental results are in good agreement with the numerical simulation results.

Performance of 2-D inlets with curved compression surface formed by segments with controllable compression angle

Suppose that curved surface consists of tiny polylines. Two 2-D inlets with curved surfaces formed by equal compression angles of the micro-element segments and a slight increase in the compression angle of the micro-element segments are designed respectively. The numerical simulation method is used to compare the performance of two curved surface 2-D inlets with the reference three-wedge compression 2-D inlet. Select NASA classic test data, in order to determine the turbulence model and calculation method chosen by the numerical simulation Fluent software. Numerical simulation results show: the pressure ratio of the curved compression system at the place corresponding to the wedge compression second shock wave is only 0.55 of the wedge compression pressure ratio, which is much smaller than the critical pressure ratio of the boundary layer separation; at design point, the performance of three inlets is similar; at non-design points, the 2-D inlets with the curved surface have a higher flow coefficient and total pressure recovery coefficient. The configuration of the small segments compression angles is the key to the performance of the curved surface compression two-dimensional inlet.

Parameterization and optimization design of a two-dimensional axisymmetric hypersonic inlet

Optimization method, as a promising way to improve inlet aerodynamic performance, has received increasing attention. The present research is undertaken to design a two-dimensional axisymmetric hypersonic inlet using parametric optimization. The inlet configuration is parameterized and optimized in consideration of total pressure recovery and starting performance. A Pareto front is obtained by solving the multi-objective optimization problem. Then, the flow structures of the optimized inlets are analyzed and the starting performances are evaluated. Results show that the total pressure loss mainly occurs in the internal contraction section, especially near the inlet entrance, and therefore the total pressure recovery coefficient can be greatly improved by decreasing external compression. As a result, the guidance for designing high-performance inlets is concluded. Besides, it is found that as the internal contraction ratio increases, the inlet starting ability becomes worse, which attributes to the larger separation bubble at the inlet entrance. Finally, the total pressure recovery coefficient and the starting Mach number of the optimized inlets are obtained, which can be a reference for engineering design.

Performance evaluation for scramjet based on ground direct-connected test: A method investigation

Scramjet has become one of the most efficient hypersonic vehicles because of its broad application potentiality. The ground direct-connected test is a relatively easy and economical way to test the performance of the scramjet under “real” condition. In this paper, an efficient performance evaluation method based on ground direct-connected test was developed. It can determine the performance parameters of the scramjet combustor (internal resistance, combustion efficiency, and total pressure recovery coefficient) and that (specific impulse and thrust) of the overall scramjet during the working time under the assumptions of adiabatic inlet and adiabatic-isentropic nozzle. Compared with the previous evaluation methods, using this method not only greatly reduces the time and money cost, but also provides more accurate boundary conditions for unsteady numerical simulation, and reference values for free jet and flight test furtherly. In the end, a case analysis was conducted and the real-time performance parameters were obtained. The results show that the internal resistance, combustion efficiency, and total pressure recovery of combustor are about 280 N, 0.69, and 0.21, respectively. The specific impulse and thrust of the overall scramjet reach about 6000 Nskg−1 and 370 Nskg−1, correspondingly.

A Simple Theory and Performance Prediction for a Shrouded Wind Turbine with a Brimmed Diffuser

We developed a new wind turbine system that consists of a diffuser shroud with a broad-ring brim at the exit periphery and a wind turbine inside it. The shrouded wind turbine with a brimmed diffuser, which we called a “wind lens turbine” (WLT), has demonstrated power augmentation by a factor of about 2–5 compared with a bare wind turbine for a given turbine diameter and wind speed. The increase in power output depends on the diffuser shape and length and the brim height. However, a simple theory presented in this paper argues that only two performance coefficients are needed to predict the performance of WLT. The coefficients are the back pressure coefficient of the brim and the pressure recovery coefficient of the diffuser. We theoretically showed that the back pressure coefficient was particularly important for the performance of WLT. Finally, the simple theory was evaluated with experimental results. The results showed good agreement with each other.

Numerical analysis on the conceptual design of an air-breathing engine inlet working in mach number 3∼5.5

Based on the equal-intensity shock theory, this article designed a supersonic inlet working in Mach number 3.0∼5.5 with the background of an air-breathing engine. The inlet applied the four-shock train mixed compression configuration and inserted a sidewall compression at the beginning of the isolator. Through developing effective 3D RANS computations validated by current experiments, the performance of the designed inlet was identified. The designed inlet self-starts at freestream Mach number Ma∞ = 3.0 under which the total pressure recovery coefficient has dramatic increment, and the aerodynamic choking at the inlet throat no longer presents; the inlet keeps working at all studied flight states with zero angle of attack (AoA) and achieves shock-on-lip at the design point Ma∞ = 5.0. Both positive and negative AoAs can disrupt the equal-intensity shock allocations, which degrade the inlet performance. The inlet obtains maximum total pressure recovery coefficient at zero AoA. The maximum back pressure at Ma∞ = 3.0∼5.5 obtained by the inlet surpasses the requirements and keeps a certain margin. The inlet performance basically meets all the goals proposed by the engine.

Experimental and numerical investigation of a supersonic inlet with double S-bend diffuser

Design of a supersonic inlet with double S-bend diffuser was developed. Numerical simulation and wind tunnel experiment were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the variable geometry scheme adopted solves the contradiction between starting performance at low Mach number and aerodynamic performance at high Mach number. The inlet works normally and stably over a wide speed range. At design point, the total pressure recovery coefficient reaches 0.47. In addition, two different kinds of inlets with double S-bend diffuser and single S-bend diffuser were studied. Compared with the double S-bend diffuser, the total pressure recovery coefficient of the single S-bend diffuser is higher at low Mach number (Ma0 < 3) and lower at high Mach number (Ma0 > 3). With the increase of backpressure, shock train mainly moves upstream along the low-energy flow region in the diffuser. For the double S-bend diffuser, shock train will first move along the lower corner and then along the upper corner. For the single S-bend diffuser, it will only move along the upper corner. The strong secondary flow of the double S-bend is the main reason for the above phenomenon.

Experimental investigation of a diffuser for use in skydiving vertical wind tunnel

Diffusers are vital components of vertical wind tunnels. In this experimental study, the velocity distribution, pressure recovery coefficient, turbulence intensity, and flow separation of a 1:13 model of a diffuser with an octagonal inlet and square outlet for use in the skydiving vertical wind tunnel have been investigated. The experiments were carried out both in suction and blowing flow modes. Results showed no significant pressure loss and separation in the diffuser under study. Also, the pressure loss results were compared with the results obtained using semi-empirical correlations for a diffuser with a circular inlet and square outlet, which is geometrically the closest to the diffuser under consideration in the present work. It was observed that to correctly predict the pressure loss in the model diffuser using the semi-empirical correlations, the diffuser big-angle plane of 10.66° should be used in these correlations instead of the equivalent cone angle. Besides, if the effect of the non-uniformity of the velocity in the diffuser is not considered in the determination of the pressure loss, a significant error will incur in determining the pressure loss coefficient.

Design and Performance of a Compact Air-Breathing Jet Hybrid-Electric Engine Coupled With Solid Oxide Fuel Cells

A compact air-breathing jet hybrid-electric engine coupled with solid oxide fuel cells (SOFC) is proposed to develop the propulsion system with high power-weight ratios and specific thrust. The heat exchanger for preheating air is integrated with nozzles. Therefore, the exhaust in the nozzle expands during the heat exchange with compressed air. The nozzle inlet temperature is obviously improved. SOFCs can directly utilize the fuel of liquid natural gas after being heated. The performance parameters of the engine are acquired according to the built thermodynamic and mass models. The main conclusions are as follows. 1) The specific thrust of the engine is improved by 20.25% compared with that of the traditional jet engine. As pressure ratios rise, the specific thrust increases up to 1.7 kN/(kg·s−1). Meanwhile, the nozzle inlet temperature decreases. However, the temperature increases for the traditional combustion engine. 2) The power-weight ratio of the engine is superior to that of internal combustion engines and inferior to that of turbine engines when the power density of SOFC would be assumed to be that predicted for 2030. 3) The total pressure recovery coefficients of SOFCs, combustors, and preheaters have an obvious influence on the specific thrust of the engine, and the power-weight ratio of the engine is strongly affected by the power density of SOFCs.

An Improved Design of a Swirl Distortion Generator

In order to study the effect of swirl distortion on the performance and stability of the compressor, this paper optimizes the design of the existing swirl distortion generator, and designs three different structures of pair swirl distortion generator. Using CFD technology, a comparative analysis of the standard pair swirl distortion flow fields and the pair swirl distortion flow fields produced by three newly designed swirl distortion generators and the existing pair swirl distortion generator is performed. The results show that the optimally designed distortion generator reproduces the standard pair swirl distortion flow field to a higher degree than existing distortion generators. On the three measurement rings, the average swirl angle error is reduced by 5.35%, 19.88%, and 8.22%. The total pressure recovery coefficient is increased by 2.2%, which proves that the design method in this paper is more reasonable and achieves the purpose of optimal design.

Mixing Enhancement by Pulsed Energy Deposition in Jet/Shock Wave Interaction

The mechanism of the supersonic round jet mixing enhancement by a pulsed energy deposition is studied numerically by solving the three-dimensional unsteady Reynolds-averaged Navier–Stokes equations with the second-order scheme in space and time. A round jet at different Mach numbers (, 1.76, 3.00) is parallelly injected into a coflowing air stream of Mach number 2.5, interacting with the oblique shock generated by a compression ramp of 20°. The density ratio between the jet and external flow is 0.5. The laser pulse with at 20 kHz is modeled as an ellipsoid for the energy addition, and proposed to put above or inside the jet for jet mixing control. It is found that the energy deposition inside the jet is much more effective for the jet mixing amplification compared with that above the jet. The jet mixing enhancement is achieved through the large-scale vortices, initiated by the energy deposition zone interacting with the oblique shock wave. It is shown that the jet cross-sectional area is doubled, and the turbulent kinetic energy is increased, even reaching 10 times the original level at the downstream location. However, a slight decrease about 0.3% of the total pressure recovery coefficient is observed far downstream.