Total Pressure Recovery Research Articles

Leading edge redesign of dual-peak type variable inlet guide vane and its effect on aerodynamic performance

Recently, a new type airfoil for variable inlet guide vane (VIGV), featuring “dual-peak” surface velocity pattern at high incidence, is proposed and shows wide low-loss operation range. To further improve its performance, this paper researches the influence of leading edge (LE) thickness and shape on the loss level and surface velocity features of the “dual-peak” type airfoil. Firstly, a polynomial-based continuous-curvature leading edge design method was briefly introduced and used in the LE redesign of sample airfoils. Then, steady simulations based on Reynolds-Averaged Navier-Stokes method (RANS), carried out by commercial software CFX after grid independent study, were used to determine the aerodynamic performance, surface velocity distribution and boundary-layer behaviors of all research airfoils. Simulation results indicate that there exists an optimized range of LE relative thickness that can achieve lower airfoil loss level at high incidence condition. For Case 1 ([Formula: see text]) and Case 2 ([Formula: see text]), the optimized LE relative thickness range is [Formula: see text] and [Formula: see text]. The LE shape optimization can further reduce the maximum incidence condition loss coefficient with proportion up to 18% for airfoils with optimal LE thickness. Analysis of flow mechanism indicates that the optimized LE thickness and shape can reduce the suction spike height and subsequent adverse pressure gradient, therefore, decrease the LE separation scale and result in a lower loss coefficient. As an application, a dual peak VIGV with circular LE, presented in previous paper as the optimized VIGV, is redesigned in the LE portion according to the research findings and achieved 0.6 percent improvement in passage-averaged total pressure recovery coefficient [Formula: see text] at extreme high stagger angle point and the low-loss operation range extends with about 5°, which confirms the effectiveness of the research findings in three-dimensional environment.

Flow Field Study of Mixed Compression Supersonic Air Intake with Cowl Ventilation

The present investigation is focused on the effect of cowl porosity on the performance of supersonic mixed compression air intake. Four different cases (namely 4.4 %, 5 %, 5.5 % and 7.2% of total cowl area) of cowl porosity at three contraction ratios of air intake have been studied. The pattern of the cowl porosity (Square shape) is chosen symmetrically along the span in the longitudinal direction from the cowl tip. Commercially available software Ansys is used in the computational studies to solve the RANS equations with the k-ω STD turbulence model. Various performance parameters of supersonic air intake are obtained and discussed. Excess amount of flow spillage appears near the cowl tip, which is responsible for the standing strong bow shock wave just before the throat for the uncontrolled case (Clean Model). The minimum energy losses and starting behavior of supersonic air intake are captured at 7.2 % cowl porosity for the contraction ratio of 1.25, which reveals the overall improvement in the flow physics and performance parameters. An increase of 32.73 % in the total pressure recovery is observed for 7.2 % cowl porosity at design contraction ratio of 1.25. All the simulations are performed at three contraction ratios of 1.22, 1.25 & 1.31.

Effect of air jet with injection pressure on the performance of mixed compression air intake

A computational study on the performance of mixed compression supersonic air intake has been carried out with and without air jet at various operating conditions. Commercially available software ANSYS was used and K-ω SST turbulence model was selected to capture the turbulent flow inside the air intake. All the simulations were simulated at a design Mach number of 2.2. Two Air jet of 1 mm diameter each and perpendicular to the local ramp surface have been placed in longitudinal direction at 0.47 times the total length of the air intake. Effect of variation of injection pressure on the flow field of air intake has been studied. Injection pressure has been varied with respect to the free stream pressure. Four different cases of injection pressure have been investigated. Three different positions (1.far away before the air jet, 2. immediately after the air jet and 3. far away behind the air jet) of normal shock were simulated to study the effect of air jet by varying the back pressure of the supersonic air intake. Significant reduction in the flow separation due the normal shock wave was noticed for all the cases of injection pressure, which further helps in improving the performance of the supersonic air intake. Important performance parameters such as flow distortion, mass flow ratio and total pressure recovery were calculated to measure the efficacy of supersonic air intake with air jet at various operating conditions.

Structural and thermal analysis on oblique detonation influenced by different forebody compressions in hydrogen-air mixtures

Various oblique detonation waves (ODW) structures have been investigated with the application background of ramjet engine, but effects of forebody compression have attracted little attention before. Two forebody compression models, the ESS (equal-strength-shock) compression and the SI (shock-isentropic) compression, have been introduced to study their effects on the ODW structure and thermal performance. Numerical results demonstrate that the compression method influences the ODW structure obviously, including the initiation location and the OSW–ODW transition type. Compared with the ESS compression, the SI compression leads to the downstream movement of initiation location, and the transition prefers to be abrupt rather not smooth. On the thermal features, the total pressure recovery is used to analyse the difference as the key parameter on the potential of engine impulse. The ESS compression corresponds a low total pressure loss when considering only the combustion process, while the SI compression performs better when considering the inlet–combustion whole process. Analysis on total pressure recovery gives the average total pressure recovery with flight Ma, altitude, and wedge angle. It is pointed out the SI compression has two shortcomings, including a long initiation length and a low total pressure recovery of combustion process, which should be considered in the future application.

Design and Numerical Investigations on a Dual-Duct Variable Geometry RBCC Inlet

Abstract A widely applicable and variable geometry 2-D rocket based combined cycle (RBCC) inlet characterized by the dual-duct design is conceptually put forward. The inlet operates as dual-duct status in the low Mach range (0~4), and transits to single-flowpath status in the following high Mach range (4~7). It accomplishes operational status transition through an 8.0-degree ramp rotation and a 4.0-degree cowl rotation at Mach 4. Through numerical simulations on typical flight Mach numbers, the observed starting Mach number is 2.2, which provides a sufficient operational window for a smooth ejector-to-ramjet mode transition. The RBCC inlet achieves comprehensive high mass capture coefficients in the overall wide flight range, especially in the low speed regimes. Suitable Mach numbers satisfying various combustion requirements in different modes together with high total pressure recovery coefficients are also obtained since the physical throat areas, compression angles, and the corresponding contraction ratios can be adjusted by a large margin through very limited rotations. The variable geometry scheme is not only feasible for practical realizations, but is also simple to arrange the dynamic sealing issues in a low-temperature environment in the RBCC engine.

Experiment and numerical investigation of flow control on a supersonic inlet diffuser

Design of a two-dimensional supersonic inlet 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. Inlet works normally and stably over a wide speed range. Compared with the diffuser without suction, the performance of the diffuser with suction is higher. Total pressure recovery coefficient σe improves by 2%-3% while outlet Mach number decreases by about 0.03 under different inflow Mach numbers. During the process of shock train moving upstream, the propagation gradually transforms from up and down mode to left and right mode. Flow separation caused by the shock train is mainly concentrated on the sidewall in the upstream of the diffuser. Sidewall suction case is more effective to eliminates the low-energy flow and stabilize the terminal shock wave. As backpressure increases, shock train changes from weak λ-type to strong λ-type and finally to X-type near the suction slots.

Optimal aerodynamic design of hypersonic inlets by using streamline-tracing techniques

Rectangular-to-Ellipse Shape Transition (REST) inlets are a class of inward turning inlets designed for hypersonic flight. The aerodynamic design of REST inlets involves very complex flows and shock-wave patterns. These inlets are used in highly integrated propulsive systems. Often the design of these inlets may require many geometrical constraints at different cross-section. In present work a design approach for hypersonic inward-turning inlets, adapted for REST inlets, is coupled with a multi-objective optimization procedure. The automated procedure iterates on the parametric representation and on the numerical solution of a base flow from which the REST inlet is generated by using streamline tracing and shape transition algorithms. The typical design problem of optimizing the total pressure recovery and mass flow capture of the inlet is solved by the proposed procedure. The accuracy of the optimal solutions found is discussed and the performances of the designed REST inlets are investigated by means of fully 3-D Euler and 3-D RANS analyses.

Erratum to: Optimization of the Number, Distribution of Work by Stages, and Solidity of Profile Vane Cascades during Compressor Design

The method of choosing the optimal number and distribution of work by stages is considered (based on providing the same total pressure recovery coefficient in stages and blade rows). A method is proposed for choosing the solidities of the profile cascades. The method effectiveness is confirmed by an example of choosing the optimal number of stages (three) and the distribution of work in a low-pressure compressor with a given of pressure ratio and peripheral speed.

Study on integrated control for supersonic inlet and turbofan engine model

Considering the supersonic inlet model with normal shock position feedback, the integrated control method of inlet and turbofan engine is studied. The integrated model includes the supersonic inlet model and the component level model of engine. Combining the relationship between the normal shock position and the total pressure recovery coefficient, the supersonic inlet and engine model is constructed. On the basis of this model, the normal shock position closed-loop control simulation is carried out, which shows that the normal shock position matching point could be stabilized near the optimal value while restraining the inlet stream disturbance. Furthermore, based on the H∞ control algorithm, an inlet and engine integrated control is designed to control the installation thrust and turbine pressure ratio with fuel, nozzle throat area, and normal shock position as control variables. The simulation results show that the response time of the integrated control is faster than the independent control. The integrated control has stronger ability to restrain the atmospheric disturbance, which could ensure the stable and reliable operation of the propulsion system.

Numerical investigation on flow and mixing characteristics inside a converging-diverging mixing duct of rocket-based combined-cycle engine in ejector mode

The supersonic mixing layers with complex background waves are formed by the rocket jet and the entrained air in the mixing section of rocket-based combined-cycle (RBCC) engine during ejector mode. To achieve mixing enhancement of primary and secondary flows in confined space, a converging-diverging mixing duct was designed in this study. For a typical flight regime of RBCC ejector mode, the effects of the mixer geometry including the contraction ratio, the throat position of mixer and the converging angle on the flow structures and mixing characteristics in the converging-diverging mixing duct under no backpressure condition were numerically investigated. A detailed discussion on the mechanism of mixing enhancement in this mixer configuration was also presented. The results indicate that appropriately increasing the contraction ratio can significantly strengthen the interactions of background shock waves with supersonic mixing layer in the converging-diverging mixing duct, which effectively promotes the rapid and sufficient mixing between primary flow and secondary flow. Besides, shortening the length from the mixer inlet to the throat position can enhance the growth rate of mixing layer and the mixing efficiency of two flows in the upstream region of converging-diverging mixing duct to some extent. However, in the downstream far-field region, the variations of throat position have little impact on the scalar mixing process in supersonic mixing layer. Moreover, it is found that increasing the converging angle has no obvious effect on intensifying the mixing performance of the converging-diverging mixing duct. On the contrary, it may lead to a dramatic decrease in total pressure recovery. Further analysis reveals that the oblique shocks generated at the compression corner and the separation shocks formed in the divergent section interact with the supersonic mixing layer, which contributes to mixing enhancement in the converging-diverging mixing duct.

The design and validation of over/under turbine-based combined cycle inlet

The design constraints including capture area and mode transition point of turbine-based combined cycle (TBCC) inlet are determined by the performance analysis of combined cycle propulsion system. According to the mass flow and thrust requirement of the combined cycle engine, the capture area of 1.0 m2 and mode transition between Ma 2.2 and 2.5 are recommended. Based on these two quantities an over/under TBCC inlet which can operate from takeoff to Ma 4.0 is designed. The variable geometry scheme of TBCC inlet is shown at different flight Mach number. Three dimensional numerical simulations of full-scale model and wind tunnel test of 8-percent scale model are explored to verify this inlet design method. The results from the numerical simulations and experimental tests indicate that the TBCC inlet could work properly during this Mach number range. The mass flow captured by the inlet and total pressure recovery at turbine/ramjet aerodynamic interface planes (AIP) meet the requirement of the TBCC propulsion system.

Investigations on the complex flows induced by dual-swept/dual-ramp wedges in supersonic flows

The caret inlet with a dual-swept/dual-ramp configuration has excellent stealth performance and aerodynamic capability. Most previous investigations on this configuration have focused on experiments and numerical simulations but there are relatively few theoretical investigations. In this study, the flow field characteristics of dual-swept/dual-ramp configuration are investigated analytically and numerically. An analytical approach that combines the shock dynamics with a “spatial dimension reduction” was used to analyze the characteristics of the wave structures and state parameters of the flow field. The effects of the sweep angles and inflow Mach number on the flow field characteristics are investigated. The results indicate that the problem of shock/shock interaction in two intersecting wedges of large back-swept angle is a problem of weak shock interaction. Therefore, the theory of weak shock interaction is used to investigate the flow field characteristics, including the uniformity of the flow field and the total pressure recovery performance.

Performance Enhancement of a Supersonic Air Intake by Applying a Heat Source

The main objective of this paper is to study the effects of applying a heat source ahead of a supersonic mixed compression air intake. This intake has been designed for the freestream Mach number 2. Four parameters of the heat source, including its location, cross-sectional shape, cross-sectional area, and heat generation rate, are investigated numerically to find a suitable heat source that enhances the performance of the intake. The research is conducted at the freestream Mach numbers of 1.8, 2, and 2.2. Total pressure recovery, mass flow ratio, flow distortion, and drag coefficient are considered as the intake performance parameters. Variation of the critical backpressure is further investigated. The results show that a bow shock is formed around the heat source that affects the intake shock system and can have desirable effects on most of the performance parameters and the critical backpressure at different freestream Mach numbers. It is also observed that the heat source location is the most effective parameter of the heat source, and should be selected meticulously.

Optimization of the Number, Distribution of Work by Stages, and Solidity of Profile Vane Cascades during Compressor Design

The method of choosing the optimal number and distribution of work by stages is considered (based on providing the same total pressure recovery coefficient in stages and blade rows). A method is proposed for choosing the solidities of the profile cascades. The method effectiveness is confirmed by an example of choosing the optimal number of stages (three) and the distribution of work in a low-pressure compressor with a given of pressure ratio and peripheral speed.

Flowfield characteristics in sidewall compression inlets

In this work, the flow fields in three-dimensional (3D) sidewall compression inlets are investigated theoretically and numerically. Combining spatial dimension reduction theory with the shock dynamics theory, a detailed analysis of the flowfield in the compression part of an inlet is proposed. Using this analytical approach, wave configurations are determined and flowfield parameters are solved; numerical simulations are conducted to validate these theoretical results. The effects of the attack angle and yaw angle on the pressure ratio, wave structures, and total pressure recovery performance are discussed in detail. Findings offer an effective way to study shock/shock interaction (SSI) in sidewall inlets, which is essential to the designing of such inlets.

Control of a Supersonic Inlet in Off-Design Conditions with Plasma Actuators and Bleed

Supersonic inlets are a key component of present and future air-breathing propulsion systems for high-speed flight. The inlet design is challenging because of several phenomena that must be taken under control: shock waves, boundary layer separation and unsteadiness. Furthermore, the intensity of these phenomena is strongly influenced by the working conditions and so active control systems can be particularly useful in off-design conditions. In this work, a mixed compression supersonic inlet with a double wedge ramp is considered. The flow field was numerically investigated at different values of Mach number. The simulations show that large separations appear at the higher Mach numbers on both the upper and lower walls of the duct. In order to improve the performances of the inlet two different control strategies were investigated: plasma actuators and bleed. Different locations of the plasma actuator are considered in order to also apply this technology to configurations with a diverter which prevents boundary layer ingestion. The potential of the proposed solutions is investigated in terms of total pressure recovery, flow distortion and power consumption.

A CFD Based Parametric Analysis of S-Shaped Inlet for a Novel Blended Wing Body Aircraft

The Advisory Council for Aeronautics Research in Europe (ACARE) has set an ambitious array of objectives to be accomplished by 2050 for civil aviation. It is often claimed that complying with those targets will not require evolution but, rather, revolution. If the growth in aviation has to be sustained in the future, then we must come up with radical aircraft and engine configurations which can meet the demands of future aviation. The AHEAD project (co-funded by the European Commission) investigated a novel multi-fuel blended wing body aircraft with a unique propulsion system to address the challenges of the future. The engine for this aircraft uses an embedded hybrid engine exploiting the boundary layer ingestion technique to increase the propulsive efficiency. Two major consequences of BLI are vital in this regard. Namely, loss of total pressure recovery and increased total pressure distortion at the Aerodynamic Interface Plane (AIP) or the engine fan-face. Hence, the inlet performance is measured by the total Pressure Recovery Factor (PRF) and Distortion Coefficient (DC60). The current research work aims to design an embedded inlet on a Blended Wing Body (BWB) aircraft that produces maximum value of PRF and minimum DC60. The aim of this research is to investigate the S-shaped inlet to understand the effect of various geometrical parameters on its performance. The Knowledge Based Engineering platform ParaPy is used to parametrize the S-shaped inlet and generate a variety of inlet geometries and volume meshes. These different variants were analysed using the Ansys® CFD code.

Experimental and Numerical Assessment on S-shaped Diffuser performance with different Turbulence Intensity

The turbulence structure in air intake S-shaped diffuser is proven to be influencing parameter on the diffuser performance. In the present research work, experimental and numerical investigations have been undertaken to explore the effect of inlet turbulence intensity level on the performance of air intake S-shaped diffuser. Detailed measurements including pressure and velocity at the inlet and outlet planes and static pressure on the top and bottom walls were taken at three pre-selected Reynolds numbers of 4.8×104, 6.4×104 and 7.5×104. The predicted corresponding turbulence intensities at the experimented three Re, were 4.16%, 3.15% and 2.8%. ANSYS-FLUENT 15 software with Standard k-ε turbulence model has been used for numerical simulations. Numerical results of static pressure recovery, total pressure loss coefficient and wall static pressure recovery have been compared with the experimental results. The experimental results indicate that increasing Reynolds number at the inlet of S-shaped diffuser resulting in slight decrease in measured turbulence intensity. Also, as Re increased from 4.8x104 to 7.5x104, the static pressure recovery increased by 9% and the total pressure loss coefficient was reduced by 4.9%.

Oblique Shock Control with Steady Flexible Panels

Flexible panels deforming under pressure loads have been suggested as a passive form of adaptive oblique shock control. This study investigates oblique shock–boundary-layer interactions on a steady flexible panel in a Mach 2.0 flow. Experiments were performed in the Imperial College supersonic wind tunnel, where shock generators were used to produce an oblique shock followed by a corner expansion. A parametric study was conducted, exploring different shock impingement positions and shock–expansion distances. The steady aerostructural response is studied using schlieren photography, static pressure distributions, photogrammetry measurements, and surface oil flow visualization. Two-dimensional numerical simulations were performed to assess the effects of the flexible panel on downstream total pressure recovery. These were validated against experimental wall pressure distributions and measurements from a Pitot rake. Results show reductions in both separation length (of up to 40%) and stagnation pressure losses (of up to 10%) if the flexible plate is used. These improvements occur for a range of shock positions spanning approximately 50% of the panel length and for all the shock–expansion distances considered. A model that captures the flow physics responsible for these trends is proposed. The results highlight the potential of flexible panels for practical oblique shock control.

Aerodynamic optimization to tandem stators by using sweep and dihedral

In the modern aerodynamic design of turbomachinery blades, the geometries of blades often need to be reshaped to achieve better aerodynamic performance by introducing extra parametric design variables. A higher variable dimension will lead to a larger sampling range as well as a sparser sample distribution, which challenges the effectiveness and stability of optimization schemes based on surrogate model by making the model prediction quality even poorer. In this paper, a multi-objective optimization based on Gaussian process model was carried out for a high dimensional design space. Based on the previous two-dimensional optimization, tandem stators of a modern compressor were optimized by the design of sweep and dihedral. The purpose of the study is to improve the aerodynamic performance of the compressor tandem stators as well as to provide an effective optimization scheme for high dimensional multi-objective optimization problems. The design of sweep and dihedral for reshaping the tandem stators consists of a total of 18 design variables. An improvement in total pressure recovery coefficient of at least 0.7% at positive incidence and at least 0.3% at negative incidence was obtained, much larger than that in the previous two-dimensional optimization. The optimization process shows that, by using Gaussian process as the surrogate model and a special sampling strategy, this optimization scheme is effective and efficient to handle this high dimensional space. The aerodynamic influences of design parameters of tandem blades were analyzed in detail and the superiority of sweep and dihedral in reducing aerodynamic loss was confirmed.