Supersonic Intake Research Articles

Highly-resolved large-eddy simulations of combustion stabilization in a scramjet engine model with cavity flameholder

Scramjet engines are high-speed airbreathing propulsion systems that do not require rotating elements to compress the air inlet stream. It is compressed dynamically through a supersonic intake system that is integrated in the forebody, thus leading to the required temperature and pressure levels for combustion to proceed within the combustor length. In such engines, the combustion chamber is crossed by a supersonic flow, which limits the time available to inject fuel, to mix it with oxidizer, to ignite the resulting mixture, and to reach complete combustion. Residence times can be increased thanks to cavities, which have the potential to stabilize combustion without excessive total pressure loss and are therefore used as flameholders in supersonic combustors. In the present study, we perform high-fidelity large-eddy simulations (LES) of a hydrogen jet in a supersonic crossflow (JISCF) of vitiated air, which is located upstream of a wall-mounted squared cavity. The performance of such high-fidelity LES does not only require the use of high-precision numerical schemes and reliable subgrid-scale models relevant to the so-called direct numerical simulation (DNS) limit, it is also strongly dependent on the mesh quality. Therefore, the present study places special emphasis on computational grid assessment through the introduction of a detailed numerical procedure, which aims at analysing mesh reliability. The corresponding procedure combines several verification subsets including (i) the inspection of distributions of the dimensions of the computational cells present at the wall location, (ii) the analysis of normalized velocity profiles and viscosity ratio in boundary layers, and (iii) the check of fields of some mesh quality indexes and associated distributions. For the geometry under consideration, it appears that the level of resolution imposed by a correct description of boundary layers leads to a mesh quality that is close to the one associated to DNS requirements. Combustion stabilization is then studied for two distinct values of the inlet vitiated airstream temperature. Two stabilization modes are recovered from the numerical simulations: cavity-stabilized and jet-wake stabilized regimes.

Repetitive Energy Deposition at a Supersonic Intake in Subcritical and Buzz Modes

This study experimentally investigated the effects of repetitive laser energy deposition using a supersonic intake model with a central conical compression surface and a cross-sectional duct in a Mach 1.92 indraft wind tunnel: especially in the subcritical and buzz modes. A single-pulse energy deposition was observed to suppress the flow separation at the compression surface by “sweeping” a shock wave system with the thermal bubble generated by the energy deposition. The duration of the sweeping effect was approximately in the subcritical mode and in the buzz mode. Furthermore, repetitive deposition of laser pulse energies was observed to moderate instabilities in both modes, and it increased the pressure recovery by as much as 8%; also, the occurrence of buzz was delayed, thereby widening the stable, subcritical regime. In both modes, there was a threshold value for the laser pulse repetition frequency , which corresponded to the duration of the sweeping effect (e.g., in subcritical mode). Below this frequency, the increase in the pressure was proportional to the repetition frequency; whereas above , the effect per single pulse was reduced.

Performance Evaluation of External and Mixed Compression Supersonic Air Intakes: Parametric Study

This research computationally evaluates the performance variations of two common types of supersonic air intakes, external and mixed compression, due to changes in the geometries of the intakes as well as some flow conditions. A computational fluid dynamics (CFD) solver was developed by the authors and used to simulate the compressible, two-dimensional (2D), and turbulent flow inside and around the supersonic air intakes. The main objective of this paper is to evaluate the role of geometric parameters, including spike angle, exit area, overall intake length, and flow parameters, such as the flight Mach number and back pressure ratio (BPR), on the performance of the intakes. Total pressure recovery (TPR), mass flow ratio (MFR), flow distortion (FD), and drag coefficient (CD) were considered as the performance factors of the supersonic intakes. The results indicated that increments of BPR had desirable effects on the performance of the intakes at supercritical conditions, and increments of the free-stream Mach number deteriorated the performance of both intakes. FD was the most sensitive parameter to the geometric variations, while TPR and MFR were almost unchanged. In addition, the superiority of the mixed compression intake was shown in terms of lower drag coefficient and FD in comparison with the external compression intake.

Performance assessment of supersonic and hypersonic intake systems with nano-particle injection

A novel approach to improve the performance of supersonic and hypersonic intake systems with nano-particle injection has been studied. A parametric study using Mach number (M∞), Stokes number (Stk), particle Eckert number (Ecp), and thermal transport number (αt) was conducted across a quasi-1D converging-diverging (C-D) supersonic intake at idealized and single-shock compression cases. Gains in stagnation pressure recovery were achieved for both the cases. Gains were observed in the idealized compression case when: Ecp>0.25 and αt>0.5 for a Mach number of 2.5; and Ecp>0.5 and αt>0.7 for a Mach number of 5. The results also showed that a combination of cooling, momentum exchange, and particle size is required to enhance intake performance. A rectangular mixed-compression intake at Mach 3 was simulated using an unsteady compressible gas-particle CFD solver in OpenFOAM. CFD simulations with nano-particle injection predicted a 16% gain in the exit pressure recovery.

Perforated Wall in Controlling the Separation Bubble Due to Shock Wave –Boundary Layer Interaction

Abstract The shock wave boundary layer interaction (SWBLI) induced separation bubble formation (SB) and its control has been investigated numerically in the mixed compression type of intake in the scramjet engine. The external compression has occurred due to the three successive oblique shocks formed from the three successive ramps of the forebody with the semi-wedge angle of 7.6, 7.0, and 9.4 respectively. The intake is designed in such a way that all three shocks converge and impinge on the leading edge of the cowl lip for the operating Mach number of 5.0. The numerical simulation is carried out by solving steady, compressible 2-D RANS equations using transitional SST k-ω turbulence model to capture the influence of SWBLI in the performance of supersonic intake. The formation of SB and its control by establishing the perforated wall in its proximity are investigated for three different cases based on the perforation with respect to SSB. Findings of the numerical simulation have concluded that the size of the SB decreases to an acceptable level while establishing the perforation in the entire fore-body wall in the isolator region. The feedback loop established between the upstream and downstream of SB could be a possible reason for reducing its size.

Design of Inward-Turning External Compression Supersonic Inlet for Supersonic Transport Aircraft

Inward-turning external compression intake is one of the hybrid intakes that employs both external and internal compression intakes principle. This intake is commonly developed for hypersonic flight due to its efficiency and utilizing fewer shockwaves that generate heat. Since this intake employ less shockwaves, this design can be applied for low supersonic (Mach 1.4 - 2.5) intakes to reduce noise generated from the shockwaves while maintaining the efficiency. Other than developing the design method, a tool is written in MATLAB language to generate the intake shape automatically based on the desired design requirement. To investigate the intake design tool code and the performance of the generated intake shape, some CFD simulation were performed. The intake design tool code can be validated by comparing the shockwave location and the air properties in every intake's stations. The performance parameters that being observed are the intake efficiency, flow distortion level at the engine face, and the noise level generated by the shockwaves. The design tool written in MATLAB is working as intended. Two dimensional axisymmetric CFD simulations validation has been done and the design meets the minimum requirement. As for the 3D inlet geometry, with a little modification on diffuser and equipping vent to release the buildup pressure, the inlet has been successfully met the military standard on inlet performance (MIL-E-5007D). This design method also has feature to fit every possible throat cross sectional shapes and has been proven to work as designed.Keywords— Inward-turning, Supersonic, Engine Intakes, Low- noise, Design Method

Transonic flow hysteresis in a twin intake model

ABSTRACTTurbulent airflow in a supersonic intake model of rectangular cross-section is studied numerically. Instability of shock waves formed in the intake and ahead of the entrance is examined. Solutions of the Reynolds-averaged Navier–Stokes equations are obtained with a finite-volume solver of second-order accuracy. The expulsion and the swallowing of the shocks with a variation of free-stream parameters are studied at both subsonic and supersonic conditions prescribed at the outlet. Hysteresis in the dependence of 2D flow on the free-stream velocity and angle-of-attack is documented. An influence of 3D effects on the flow is examined.

Comparative Flow Field Analysis of Boundary Layer Diverter Intake and Diverterless Supersonic Intake Configuration

In this paper comparative flow field analysis of two intake configuration i.e. Boundary Layer Diverter Intake and Diverterless Supersonic Intake is carried out based on dimensionless parameters under various flow conditions. Numerical analysis of aircraft intake is a complex phenomenon which involves both external and internal flow analysis. In this research, both external and internal flow characteristics of intake duct are analyzed in detail. A comprehensive mesh scheme is devised and implemented to accurately capture the flow behavior in external surrounding of intake duct and flow passing through the intake duct. The analysis is carried out at different flow conditions to analyze the flow behavior in subsonic and supersonic regimes. Engine design mass flow rate is used for accurate intake analysis and results are validated with available literature. Boundary layer diversion and pressure recovery are examined for each intake configuration and comparative analysis based on pressure recovery is carried out subsequently. The analysis reveals that at subsonic and transonic regimes, Boundary Layer Diverter intake is much more effective than Diverter less Supersonic Intake, however, in supersonic regime Diverter less Supersonic Intake is found be to more effective. The research can further help in modifying/ improving the design of an existing intake configuration for enhanced intake efficiency.

Проектирование гиперзвукового входного устройства на основе течения в угловом теле

Проектирование гиперзвукового входного устройства на основе течения в угловом теле

Study of Flow Field inside a Rectangular Air Intake with Pointed Cowl for Different Back Pressures at Mach 2.2

Computations and Experiments were performed to get an understanding of the flow field around a rectangular supersonic air intake with pointed cowl [90°] at different back pressures for Mach 2.2. The effect of Cowl shape on the ramp surface pressure distribution is discussed and compared with existing V-Notch [90°] intake model at free exit condition. It was found that using pointed cowl [90°] intake model, a better pressure recovery was achieved compared to the V-Notch [90°] intake model at Mach 2.2. Both Pointed and V-Notch intake models showed good starting characteristics. For change in back pressure, the occurence of normal shock, flow separation zone and flow reversal were observed. All experiments are performed only for the Pointed cowl [90°] intake model. All the 3-D computations were performed by using software available at B.I.T, Mesra, Ranchi.

Alterations of Cowl Lip for the Improvement of Supersonic-Intake Performance

This paper discusses the performance enhancement of supersonic air intake model through the implementation of blunted leading edge to the cowl lip section of the model. A supersonic air intake model with sharp cowl leading edge is initially considered to numerically investigate its performance. Mach 3, supersonic intake flow through the base model has been simulated using commercial CFD package Ansys Fluent-15. Comparison of numerical predictions and experimental measurements is presented to demonstrate the correctness and accuracy of numerical frame work followed in the present study. Higher order spatial accuracy of the solver along with suitably refined mesh helped in accurate capturing of the flow field. Modification to the cowl lip is proposed as an effective method to improve the performance of the supersonic air intake. Two different blunted cowl leading edge geometries were investigated to identify the possible enhancement in performance parameters. Improvement in mass capture and combustion stability attained through the use of forward shifted blunt cowl leading edge is presented. It is also revealed through the present study that the blunt cowl leading edge can reduce the intensity of shock wave boundary layer interaction occurring at the isolator entry section. Deviation in total pressure recovery and flow distortion observed with different supersonic air intake models are also discussed with reasons for the same. This study demonstrates the scope of overall improvement in scramjet engine performance through the use of suitably positioned blunt cowl leading edge.

Investigation of Flow Field Around the Pointed Cowl Air Intake at Mach 2.0

Abstract Experiments and computational studies were carried out to get an understanding of the flow field around a rectangular supersonic intake with pointed cowl shape. Experiments include quantitative pressure measurements and flow visualization studies by using schlieren techniques. The effects of the presence of various cowl shapes on ramp surface have been obtained computationally at Mach 2.0. The experiments were carried out only for the pointed cowl. Schlieren Photographs were taken. Three-Dimensional simulations were made by using FLUENT at supersonic speed. The details of the experiments and computations are discussed.

Design and Optimization of Supersonic Intake

The ideal supersonic intake was designed using the Theta-Beta-Mach relations and a CAD modelling software. After the design was done the optimization part was carried out. Initially the easy method of cowl deflection angle variation was used to get the optimum results at a particular angle and then the pressure recovery was maximized using the intake bleed technique. The bleed technique was optimized by fixing the best location of bleed according to the shock reflections in the intake and then maximizing the results analyzing the problem for various bleed diameter values. The mesh resolution for all the analyses was taken constant along with the turbulence model and solution methods. The analyses were carried out completely considering the free flow exit condition in the intake and the back pressure condition was not taken into account due to time constraints. Therefore, the pressure recovery or the efficiency of the intake is judged solely on the minimum velocity achieved in the intake. This value was used to maximize the intake too.

Forced Shock Oscillation Control in Supersonic Intake Using Fluid–Structure Interaction

This paper proposes a fluid–structure interaction-based method to restrict the undesirable shock oscillation process due to the downstream pressure perturbation and limit its adverse impact on a two-dimensional supersonic intake. A computational framework is established to illustrate the coupled solutions between the forced shock oscillation and the nonlinear structural vibration. Complex intake aerodynamics are captured with an implicit finite volume method. Finite element code is employed to determine the dynamic response of a flat flexible wall locally installed in a rigid cowl wall. Data transfer between the two physical fields is accomplished with a partitioned coupling scheme. The large-amplitude shock oscillation and poor variation in aerodynamic performance are found in the uncontrolled transient flow solutions of the rigid wall intake model. Conversely, a shock oscillation amplitude reduction of about 50% is achieved by the local flexible wall, and it shows an improvement to the aerodynamic performance of the supersonic intake as well. This paper demonstrates that the dynamic cavity generated by fluid–structure interaction contributes to outstanding performance of this new method. Additionally, the impact of the dynamic cavity is interpreted by two fundamental mechanisms: the local enlarged flowpath with a convergent–divergent shape and energy transfer within the fluid–structure interaction.

Performance Prediction of Internal Compression Supersonic Air Intake at Range of Mach Numbers (1.1-1.5)

In this research a numerical investigation on a supersonic air intake was done. The aim of this work is to investigate a variable geometry of cross-section area for supersonic air intake at range (1.1-1.5) Mach number, to get a maximum pressure recovery. In this work, the flow starts with a normal shock attached to the intake cowl lip. The flow is assumed compressible, inviscid, two-dimensional flow, unsteady, and axisymmetric. The equations (Continuity, Momentum, and Energy) were solved based on a finite volume method. The governing equations were solved iteratively using time marching technique. This part is analyzed for several Mach numbers, where the flow properties are determined from inlet of air intake to the diffuser exit. Results show that, the implementation of time marching scheme has succeeded in the prediction of the choked flow region, which is important in the study of the performance of convergent-divergent diffuser. Also the results indicated the absolute velocity increases along the convergent part and then start to decrease along divergent part independently on the values of free-stream Mach numbers.

Three-dimensional numerical simulation of the characteristics of a ramjet power plant with a continuous-detonation combustor in supersonic flight

Multi-variant three-dimensional numerical simulations demonstrate the feasibility of the continuous- detonation process in an annular combustor of a ramjet power plant operating on hydrogen as fuel and air as oxidant in conditions of flight at a Mach number of M 0 = 5.0 and an altitude of 20 km. Conceptual schemes of an axisymmetric power plant, 400 mm in external diameter and 1.3 to 1.5 m in length, with a supersonic intake, divergent annular combustor, and outlet nozzle with a frusto-conical central body are proposed. Calculations of the characteristics of the internal and external flows, with consideration given to the finite rate of turbulent-molecular mixing of the fuel mixture components with each other and with the combustion products, as well as the finite rate of chemical reactions and the viscous interaction of the flow with the bounding surfaces, have shown that, in these flight conditions, the engine of such a power plant has the following performance characteristics: the thrust, 10.7 kN; specific thrust, 0.89 (kN s)/kg; specific impulse, 1210 s; and specific fuel⋅consumption 0.303 kg/(N h). In this case, the combustor can operate with one detonation wave traveling in the annular channel at an average velocity of 1695 m/s, which corresponds to a detonation wave rotation frequency of 1350 Hz. It is shown that, an operating combustor has regions with subsonic flow of detonation products, but the flow is supersonic throughout its outlet section.

Buzz Cycle Description in an Axisymmetric Mixed-Compression Air Intake

Buzz phenomenon is shock oscillation ahead of the supersonic air intake when its mass flow rate is decreasing at off-design condition. The buzz onset and the buzz cycle of an axisymmetric mixed-compression supersonic intake have been experimentally investigated through pressure recording and shadowgraph flow visualization. The intake was designed for a freestream Mach number of 2.0; however, tests were conducted for , 2.0, and 2.2. All tests were performed at 0 deg angle of attack. Results show that there is a strong relation between the acoustic characteristics of the intake and the buzz fluctuations. This relation causes a new pattern for the buzz oscillations, large-amplitude oscillations with large frequency, that has features of both little buzz (Ferri-type instability) and big buzz (Dailey-type instability). Flow separation caused by the shock-wave/boundary-layer interaction and acoustic compression waves are the most important driving mechanisms in the buzz cycle. Both amplitude and frequency of the buzz oscillations vary as the backpressure is varied. As the freestream Mach number is increased, the dominant frequency of the buzz oscillations is decreased.

A Numerical Analysis of Supersonic Intake Buzz in an Axisymmetric Ramjet Engine

A numerical analysis was conducted to investigate the inlet buzz and combustion oscillation in an axisymmetric ramjet engine with wedge-type flame holders. The physical model of concern includes the entire engine flow path, extending from the leading edge of the inlet center-body through the exhaust nozzle. The theoretical formulation is based on the Farve-averaged conservation equations of mass, momentum, energy, and species concentration, and accommodates finite-rate chemical kinetics and variable thermo-physical properties. Turbulence closure is achieved using a combined scheme comprising of a low-Reynolds number k-e two-equation model and Sarkar’s compressible turbulence model. Detailed flow phenomena such as inlet flow aerodynamics, flame evolution, and acoustic excitation as well as their interactions, are investigated. Mechanisms responsible for driving the inlet buzz are identified and quantified for the engine operating at subcritical conditions.

超音速機の機体/推進系統合設計におけるインテークランプの性能評価

超音速機の機体/推進系統合設計におけるインテークランプの性能評価

Numerical Investigation of Supersonic Ram-Jet Intake Operating Characteristics

Numerical Investigation of Supersonic Ram-Jet Intake Operating Characteristics