Entrance Mach Number Research Articles

Experimental Evaluation of Isolated Intake for Hypersonic Air- Breathing Vehicle

Isolated intake of an air-breathing conceptual hypersonic demonstrator vehicle has been experimentally characterized simulating the entry Mach number and the entry flow direction. A 1:8 scale intake model having provision to vary the entry area of the internal compression portion by means of actuating the intake cowl has been tested in the supersonic freestream Mach numbers 3, 3.44 and 3.9. Tests were carried out in the 0.3 m x 0.3 m trisonic wind tunnel at National Aerospace Laboratories, Bangalore. The result pertaining to M¥ = 3.44 is discussed in this paper. The effect of cowl position on the intake starting characteristics and performance are analyzed using wall pressure measurements, schlieren and oil flow visualization. Intake unstart occurred before the cowl reaching to its fully open position with the original hinge location. Essentially the strength of cowl leading edge shock increases as the cowl opening angle increases. The strong shock-boundary layer interaction in the convergent portion of intake aids the growth of the separation bubble in the flow path and that cause the intake to unstart. Improvement of intake starting by means of shifting the cowl hinge location downstream was attempted. Shifting the cowl hinge location downstream has improved the intake starting characteristics. On the improved configuration, the effect of combustion induced back pressure and the effect of cowl opening speed on intake unstart have been analyzed and quantified.

Experimental investigation on the ignition limits of a kerosene-fueled cavity in model scramjet combustor

The ignition limits of liquid kerosene injected upstream of a cavity flameholder using a spark plug are investigated in a model scramjet combustor with the isolator entrance Mach number of 2.52, the total pressure of 1.6 MPa, and stagnation temperature of 1486 K. The effects of injection pressure and mixing distance on ignition are studied. Flame dynamics and the evolution of the flow field during the ignition process are acquired through high-speed photography and Schlieren photography. For near-field injection, successful ignition of kerosene can only be achieved when the injection pressure is extremely low or in a higher injection pressure range. The ignition is failed due to the fuel-rich local equivalence ratio in the cavity recirculation zone when the injection pressure is in the middle of injection pressure range. For far-field injection, the ignition trend is totally opposite compared with near-field injection. The ratio of the fuel that the cavity can react completely to that consumed by the mainstream is extremely low due to the minimal mass exchange rate between the cavity and the mainstream. The local equivalence ratio in cavity is the controlling factor determining whether successful ignition could be achieved. Only local ignition with low combustion efficiency is achieved in the present single-cavity combustor. The flame could not spread to the mainstream due to the lower turbulence level weakening the interaction of combustion products in the recirculation zone with unburned reactants in the supersonic flow.

Shock train/glancing shock/boundary layer interaction in a curved isolator with sidewall contraction

Herein, a curved isolator with sidewall contraction of a hypersonic scramjet is extracted and simplified. The flow characteristics in the isolator with an entrance Mach number of 3.46 are studied. The results show that the lateral contraction of the sidewall generates a strong shock and its subsequent reflected shock, which induce glancing shock wave/boundary layer interaction (GSWBLI), rolling up large-scale vortices adjacent to the symmetry plane. As the downstream backpressure increases, the shock train propagates upstream and interacts with the glancing shock and vortices inevitably, forming a typical shock train/glancing shock/boundary layer interaction (STGSBLI) phenomenon. Different from the shock train in a straight isolator, it can be divided into two distinct parts, i.e., the center part and the side part. For the center part, it behaves in a quasi-two-dimensional manner, which is primarily associated with STGSBLI, forcing the low-momentum subsonic flow to principally accumulate near the symmetry plane and weakening the spanwise pressure gradient in the center part. For the side part, it locates behind the glancing and reflected shocks and is much shorter than the center part, resulting in a streamwise extension of the supersonic flow region near the sidewall. As the shock train moves upstream, the quasi-two-dimensional region enlarges and extends to the sidewall. It is demonstrated that the aforementioned two parts of the shock train can also be discovered in curved sidewall-contraction isolators with different centerlines. Therefore, the STGSBLI is the dominant and universal physical phenomenon in isolators with sidewall contraction.

Effect of Centrifugal Force on Gas Flow in a Supersonic Turbine

Abstract To clarify the loss mechanism in supersonic turbines, the authors experimentally and numerically studied the flow condition between the rotor blades of a supersonic turbine for liquid rocket engines. The entrance Mach number was 1.94, and the turning angle of the blades was 120 deg. Shock waves were created at the leading edge of the blade. The Mach number and total pressure decreased in the passage between the blades, where the flow condition was restricted by the centrifugal force. Such degradation, which has also been reported in past studies, is an inherent feature of high-speed, large-turning-angle blades. Here, it was found that a convergent–divergent configuration of the passage between the blades suppressed the performance degradation of the supersonic turbine.

Experimental study of ignition and combustion characteristics of ethylene in cavity-based supersonic combustor at low stagnation temperature and pressure

At low inflow Mach number without throttling, since low static temperature the ignition and flame holding is still a big challenge in supersonic combustor. In this paper, the ignition and flame transient development in ethylene fueled cavity-based supersonic combustor are experimentally investigated at low conditions of stagnation temperature 950 K, pressure 0.82 MPa, and the isolator entrance Mach number 2.0. The results show ethylene can achieve successful ignition and stable combustion at low inflow Mach number with global equivalence ratios (ER) varying from 0.12 to 0.3. The combustion mode could be either supersonic or subsonic which depends on the ER. The combustion mode is supersonic stable combustion mode at ER = 0.12, 0.18 and 0.24 while the mode is subsonic stable combustion mode at ER = 0.27, 0.3. Owing to the better fuel/air mixing, ethylene injected from upstream position burns more intensively than that of downstream position at the same ER. It is also found that combustion and shockwaves are strengthened with each other until the balance between the pressure of shockwaves and the backpressure of combustion. This balance makes the flow structures and combustion keep stable. Comparing the combustion and shockwaves of different ER and injector position, the mutual promotion between combustion and shock wave is further demonstrated.

Combustion modes periodical transition in a hydrogen-fueled scramjet combustor with rear-wall-expansion cavity flameholder

Cavity-stabilized combustion modes periodical transition in a laboratory hydrogen-fueled scramjet combustor were captured by high speed imaging. Experiments were performed with an isolator entrance Mach number of 2.92, and a sonic transverse fuel jet upstream of the cavity was employed. The reproducibility of the results had been tested in several repeated experiments. It is observed that, under a fixed combustor entry flow parameters and fuel equivalence ratio, after the flame has been fully developed, the flame structure may periodically transform between two different combustion modes. Specifically, under a moderate equivalence ratio, the flame structure switches between the cavity shear-layer stabilized mode and the combined shear-layer/recirculation stabilized mode, which shows an apparent periodicity with a period of about 7.5 ms. The formation mechanism for this novel phenomena is analyzed and some suppositions are given. When increasing the equivalence ratio to a high level, the flame structure shows a quasi-periodic low-frequency oscillation and the combustion mode changes between the combined shear-layer/recirculation stabilized mode and the cavity-assisted jet wake stabilized mode. At last, the formation mechanism and characteristics of the combined shear-layer/recirculation stabilized mode are detailed analyzed.

Experimental study on the coupling process of flow and combustion phenomena during pilot ignition transients

The pilot ignition of kerosene under different pre-established hydrogen flames is performed in a model scramjet combustor with the isolator entrance Mach number of 2.52, the total pressure of 1.6 MPa, and stagnation temperature of 1486 K, which corresponds to Ma 5.5 flight condition. The flame dynamics and evolvement of the precombustion shock train during pilot ignition transients are acquired through high-speed photography and schlieren photography. The flame stabilization mode of pilot hydrogen belongs to cavity-recirculation-region stabilized combustion with smaller equivalence ratio. While the flame stabilization mode turns into jet-wake stabilized combustion as the equivalence ratio of pilot hydrogen increases. Liquid kerosene can not be ignited by cavity-recirculation-region stabilized flame, while it can be ignited successfully by pilot hydrogen flame stabilized in the jet-wake. Thermal choking resulted from combustion of the pilot fuel should be large enough to produce and sustain the precombustion shock train in the combustor for igniting liquid kerosene. There is a positive feedback mechanism between combustion and flow dynamics. The backpressure caused by combustion produces and sustains the precombustion shock train, which in turn increases combustion intensity. The positive feedback mechanism lasts until a new balance between combustion and flow phenomena is achieved. Boundary-layer separation is the main reason for the upstream propagation of flame, which is a result of the large adverse pressure gradient caused by heat release.

Experimental study on the effects of simulated flight conditions on ignition and flame stabilization in a supersonic combustor

Effects of simulated flight conditions on ignition and flame stabilization in a kerosene fueled scramjet combustor were experimentally investigated in the present paper. Simulated flight conditions were Mach 4.5 and Mach 4.0, and the isolator entrance Mach number was 2.0. Wall pressure and flame emission were made during the experiments in an attempt to better understand the combustion characteristics, air throttling was used to enhance the flame stabilization. The results showed that, the kerosene ignition was achieved successfully under the Mach 4.5 flight condition, that of Mach 4.0 flight condition was blowout before the kerosene pressure reached the max pressure. The flame in the combustor without air throttling during Mach 4.0 flight condition only existed in the cavity and near the top wall before it was blown off, that of Mach 4.5 flight condition had spread into the isolator. The flame stabilization was also only achieved under the Mach 4.5 flight condition with the aid of air throttling, the flame in the combustor with air throttling during the Mach 4.0 flight condition was blowout before the throttling air was removed. Ignition and flame stabilization were easy to be achieved under the higher flight condition.

Investigations on flowfield behavior and resistance backpressure characteristics of supersonic cascade with boundary layer suction

A numerical study aiming at the enhancement of the operating range of supersonic cascade via boundary layer suction is presented. The purpose of this paper is to improve the maximum backpressure ratio of the flowfield by suction control and to learn the flowfield behavior and characteristics with different suction control parameters. An S-shaped supersonic cascade developed and tested at DLR (DLR-PAV-1.5) that operates with a relative entrance Mach number of 1.5 is selected in this paper. A two-dimensional multiblock NS solver has been successfully applied to the cascade flow which is validated using the data of the experiment of Schreiber. An in-deep numerical study has been carried out to investigate the effects of suction locations, widths and angles by comparing the operating range and aerodynamic performance parameters of the supersonic cascade. The results indicate that the maximum backpressure ratio can be enhanced by 17.4% with a loss of 3% mainstream mass flow by suction slot in 60% or 70% length of the axial chord. As the width of the suction slot increases to 7 mm, the maximum backpressure ratio is enhanced by 24%, and the total pressure recovery coefficient is improved to 95.3%. The range of suction slot angle that can effectively improve the maximum backpressure ratio is 60–90° and −90 to −60°. This paper provides guidance for the supersonic cascade to improve its performance parameters through suction control.

Flame stabilization mechanism study in a hydrogen-fueled model supersonic combustor under different air inflow conditions

Flame stabilization mechanisms of a hydrogen-fueled laboratory model scramjet combustor are numerically studied. The combustor consists of a sonic fuel jet injected paralleling into a supersonic air flow at the base of a wedge-shaped strut. The numerical calculations are based on an Open Source Field Operation and Manipulation (OpenFOAM) solver using the large eddy simulation method. A skeletal mechanism for hydrogen-air chemistry including 9 species 27 reactions is used to depict the combustion dynamics. Three cases with air stagnation temperature of 568 K, 460 K, and 960 K, are respectively studied under the same equivalence ratio and entrance Mach number. Damköler number based on the subgrid turbulent mixing time scale and the instantaneous chemical time scale is used to distinguish the role of mixture self-igniting from flame propagation in different air inflows. Flame stabilization in such strut-based combustor is compared with that of cavity-based one. It is found that the basic working principles for the two types of combustor are similar. However, the strut shows more simplicity in flame stabilization locations.

Experimental investigation on laser-induced plasma ignition of hydrocarbon fuel in scramjet engine at takeover flight conditions

Laser-induced plasma ignition of an ethylene fuelled cavity is successfully conducted in a model scramjet engine combustor with dual cavities. The simulated flight condition corresponds to takeover flight Mach 4, with isolator entrance Mach number of 2.1, the total pressure of 0.65 MPa and stagnation temperature of 947 K. Ethylene is injected 35 mm upstream of cavity flameholder from four orifices with 2-mm-diameter. The 1064 nm laser beam, from a Q-switched Nd:YAG laser source running at 10 Hz and 940 mJ per pulse, is focused into cavity for ignition. High speed photography is used to capture the transient ignition process. The laser-induced gas breakdown, flame kernel generation and propagation are all recorded and ensuing stable supersonic combustion is established in cavity. The highly ionized plasma zone is almost round at starting, and then the surface of the flame kernel is wrinkled severely in 150 μs after the laser pulse due to the strong turbulence flow in cavity. The flame kernel is found rotating anti-clockwise and gradually moves upstream as the entrainment of circulation flow in cavity. The flame is stabilized at the corner of the cavity for about 200 μs, and then spreads from leading edge to trailing edge via the under part of shear layer to fully fill the entire cavity. The corner recirculation zone of cavity is of great importance for flame spreading. Eventually, a cavity shear-layer stabilized combustion is established in the supersonic flow roughly 2.9 ms after the laser pulse. Both the temporal evolution of normalized chemiluminescence intensity and normalized flame area show that the entire ignition process can be divided into four stages, which are referred as turbulent dissipation stage, combustion enhancement stage, reverting stage and combustion stabilization stage. The results show promising potentials of laser induced plasma for ignition in real scramjets.

Numerical and experimental study of supersonic combustion chamber at entrance Mach number 4. Part 1: Non-reacting flow

Results of joint experimental and numerical study are presented for a non-reacting 3D flow in a channel with backward-facing steps (BFS) located on top and bottom walls. Experiments were conducted in a hot-short wind tunnel at entrance Mach number 4, total pressure 9.5 bar and total temperature 2120K. Argon was injected into the channel from eight round holes, placed on channel walls upstream of the channel expansion. Simulations performed in a 3D RANS approach with k-ω SST turbulence model give a reasonable agreement with experimental data and allow evaluating the pressure flowfield and argon mass concentration in the channel.

Effect of the configuration of the rear wall of the cavity on combustion in a supersonic combustor

Results of an experimental study of a supersonic combustor with a solid (continuous) and discrete (discontinuous) rear wall of the cavity are reported. The tests are performed in a short-duration wind tunnel with an attached pipeline at the entrance Mach number of 3, total temperatures of 2550–3140 K, and static pressures of 178–195 kPa. Heated kerosene is used as a fuel. Data on ignition conditions and kerosene combustion efficiency are obtained for different cavity configurations. The drag of internal elements of the combustor, which form recess stabilizers, is estimated. An overall advantage of the combustor with a discrete rear wall of the cavity over a solid rear wall is demonstrated.

Dynamic combustion characteristics in a rectangular supersonic combustor with single-side expansion

Dynamic combustion characteristics of a rectangular scramjet combustor with single-side expansion were studied experimentally and numerically. Experiments were implemented with an isolator entrance Mach number of 3.46, and an air stagnation temperature of 1430 K. Ethylene was utilized to fuel the combustor over an equivalence ratio range of 0.20 < φ < 0.63. Results indicated that the combustion modes varied from different equivalence ratios. For an intermediate φ = 0.375, an intermittent dynamic combustion occurred. During the dynamic process, the flame sometimes stabilized in the jet wake of the top cavity, and at other time it oscillated between dual parallel cavities. The pseudo-shock train traveled periodically along the length of the combustor, and the penetration depths of the two injectors exchanged. Quantitative analysis illustrated that the average frequency of unsteady combustion was approximately 200 Hz. The reason for the occurrence of the self-sustained dynamic process was related to the interactions between the shock-induced separated region and heat release.

Development of body force model for steady inlet distortions in high-speed multistage compressor

This study aims at establishing a three-dimensional numerical model, compressor aerodynamic performance analysis model, to simulate the impact of complicated distorted flow on multistage axial flow compressor based on the body force model. The model solves the compressible three-dimensional Euler equations, which are modified to include source terms representing the effect of the blade rows. In this study, the association between blade source terms and entry Mach number together with attack angle could be established with the deviation angle model and loss model. In this paper, compressor aerodynamic performance analysis model is used to evaluate the effect of inlet circumferential total pressure distortion and swirl distortion on a five-stage high-pressure compressor. Calculated operating maps for compressor agree well with the experimental results. Meanwhile, the traveling process of inlet distortions in the multistage compressor is correctly revealed. The wide application prospect of the model can be seen in the area of inlet distortion problems.

Investigation on Ignition Enhancement Mechanism in a Scramjet Combustor with Dual Cavity

The spark ignition and flame stabilization of kerosene in a dual-cavity scramjet combustor is investigated with the isolator entrance Mach number of 2.52, total pressure of 1.6 MPa, and stagnation temperature of 1486 K, which corresponds to flight condition. The effects of dual-cavity configuration on ignition and flameholding were analyzed. Two distinct cavity-organized flame regimes were found: cavity flame and cavity shear-layer flame. The driving force of the transition of flame stabilization mode from cavity flame to cavity shear-layer flame is the high pressure resulted from combustion, which occurs in the cavity. The dual-cavity flameholding scheme can promote the ignition ability. First, the downstream cavity could stabilize the flame convected from the upstream cavity. Second, compared with the upstream cavity, the downstream cavity is a more suitable ignition location because the fuel has a longer distance for evaporation and mixing. The boundary-layer separation and upstream moving of precombustion shock train are responsible for the contraflow flame propagation process. Also, an unstable boundary-layer/cavity-stabilized combustion is found during the flame transition process. This phenomenon results from the interaction between the cavity recirculation zones and the separation zone immediately upstream of T1 cavity, which is caused by the strong shock/boundary-layer interactions, as the shock train is pushed upstream.

Engine performance analysis and optimization of a dual-mode scramjet with varied inlet conditions

A dual-mode scramjet can operate in a wide range of flight conditions. Higher thrust can be generated by adopting suitable combustion modes. Based on the net thrust, an analysis and preliminary optimal design of a kerosene-fueled parameterized dual-mode scramjet at a crucial flight Mach number of 6 were investigated by using a modified quasi-one-dimensional method and simulated annealing strategy. Engine structure and heat release distributions, affecting the engine thrust, were chosen as analytical parameters for varied inlet conditions (isolator entrance Mach number: 1.5–3.5). Results show that different optimal heat release distributions and structural conditions can be obtained at five different inlet conditions. The highest net thrust of the parameterized dual-mode engine can be achieved by a subsonic combustion mode at an isolator entrance Mach number of 2.5. Additionally, the effects of heat release and scramjet structure on net thrust have been discussed. The present results and the developed analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets. A dual-mode scramjet can achieve high thrust in a wide range of flight conditions. The present work uses modified quasi-1-D analysis model and simulated annealing strategy to investigate the engine performance of kerosene-fueled parameterized dual-mode scramjet at a crucial flight Mach number of 6. Engine structure and heat release distributions were optimized to obtain optimal engine net thrust for varied inlet conditions. The highest net thrust can be achieved by subsonic combustion mode at an isolator entrance Mach number of 2.5. The present results and analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets.

Experimental Study of Combustor–Isolator Interactions in a Dual-Combustion Ramjet

Combustor–isolator interactions are investigated in a direct-connect dual-combustion ramjet experiment. The test apparatus contained an annular isolator, a constant cross-sectional area cylindrical supersonic combustor, and a subsonic-burning gas generator. The gas generator, in which a liquid hydrocarbon fuel was injected, produced a fuel-rich hot-gas mixture that is mixed with the annular supersonic airflow in the supersonic combustor. A precombustion compression field was generated, both in the isolator and the combustor, by the heat release. The annular isolator entrance Mach number was varied as 1.78, 1.98, and 2.23. The wall static pressure distributions measured in the isolator and the combustor are presented and analyzed to determine the precombustion shock-train length. The new correlating equations and the corresponding theoretical model that can predict the precombustion shock-train length and the wall pressure distribution are derived therefrom.

A mechanism of combustion mode transition for hydrogen fueled scramjet

The combustion mode transition will occur repeatedly during the real flight of a hydrogen fueled scramjet. Under the assumption of inviscid flow, a series of the equations have been developed to identify combustion mode transition in an analytical way. The processes of heating and accelerating induced mode transition are investigated, respectively. This paper presents an analytical solution to understand the conditions for the transition between ramjet-scramjet operation modes. Besides, this paper also presents a qualitative explanation of the mode transition processes in the physical point of view. The mode transition is closely related to the thermal choking effects and shock wave motion. The process of mode transition is discontinuous and very fast. There is a sudden change with the combustor entry Mach number in this process.

Analyse the electrical characteristics of an MagnetoHydroDynamic generator for maximizing the thermal efficiency

In this study, a Faraday type MagnetoHydroDynamic (MHD) generator is studied to consider the effect of electrical characteristics to the thermal efficiency. The generator performance is specified by optimizing the cycle efficiency with respect to the load parameter and by optimizing output power density with respect to seed fraction and operating pressure. As the calculation results, the value of load parameter, which maximizes the thermodynamic efficiency, is independent of the regenerator efficiency, but dependent on Mach number and the compressor efficiency. It can also be seen that there is no need for a high entrance Mach number more than 5 because the increases in thermal efficiency are insignificant.