Axisymmetric Supersonic Inlet Research Articles

Source of buzz instability in a supersonic air inlet

An axisymmetric supersonic inlet is computationally studied at a freestream Mach number of 2.0. The Reynolds-Averaged Navier-Stokes (RANS) equations are solved in an unsteady state and the SST k-ω turbulence model is used in this study. Several throttling ratios have been set to study the inlet flow field from supercritical to subcritical conditions, including little buzz and big buzz. Recently, studies have been about the new reasons for the buzz initiation beyond the accepted buzz criteria. The primary purpose of the current research is to discover the source of the buzz phenomenon in a supersonic inlet by examining the Ferri and Dailey criteria, which are the main criteria for predicting the buzz onset. The results showed that the low-speed flow above the vortex sheet, generated due to the intersection of the conical shock with normal shock, and the flow separation near the cowl lip are the main reasons for the buzz onset in the current inlet. To further study the source of buzz instability, the boundary layer suction was used on the cowl surface to investigate its effects on the buzz onset. According to the results, the bleed can remarkably postpone the buzz phenomenon and increase the last stable throttling ratio of the inlet from 65.0% to 77.5%. Furthermore, the big buzz at the throttling ratio of 80.0% is converted to the little buzz and the inlet performance is significantly improved after applying the bleed.

Buzz characteristics and separation bubble dynamics in supersonic intake

The flow field in an axisymmetric supersonic mixed-compression inlet is numerically solved to understand the separation bubble dynamics at different freestream Mach numbers of M∞ = 2.0, 3.0, 5.0, and throttling ratios of ψ = 0.515, 0.40, and 0.35. The two-dimensional compressible Reynolds Averaged Navier Stokes equations were solved in a finite volume framework. The computational results were validated with the wind tunnel experimental data for steady as well as unsteady operating conditions. Throttling conditions are simulated by varying the exit area of the model in the form of plug inserts. The analysis of the flow features in the complete computation domain indicates the formation of multiple separation bubbles near the exit plug as well as near the isolator entrance of the inlet. The frequency of unsteadiness (f) associated with the ramp surface pressure decreases with a decrease in throttling ratio ψ. The possibility of a common origin that is regardless of any buzz criteria is analyzed by considering separation and reattachment zones in the flow field in terms of wall shear stress variation. Inside the supersonic intake, the multiple shock waves interact with the wall boundary layer that can lead to a large separation bubble. This bubble is always unstable inside the intake flow passage and can generate the source of instability. The propagation of a pressure disturbance is analyzed by the temporal variation of the pressure on the ramp wall. The cross-correlation coefficient Cij(t) of the pressure signal on the ramp surface suggests the upstream propagation of the disturbance generated from the separation zone near the exit plug. The numerical Schlieren provides further information on the formation and propagation of the separation bubble near the exit plug.

Axisymmetric and three-dimensional flow simulation of a mixed compression supersonic air inlet

The flow through an axisymmetric supersonic mixed-compression air inlet has been simulated numerically to investigate the effects and the necessity of the three-dimensional (3D) modeling in comparison with the axisymmetric one. For this purpose, a supersonic inlet has been simulated numerically via axisymmetric and 3D CFD solvers, using the steady state Reynolds-averaged Navier-Stokes equations along with the SST k-ω turbulence model, for a free-stream Mach number of 2.0 and at zero degrees angle of attack. The grid for the 3D cases was a 14.4-degree sector, instead of a 360-degree domain one, with rotational periodic boundary condition for the side boundaries. The results show that both static and total pressure distributions match well with the experimental data for both the axisymmetric and the 3D simulations. If the prediction of performance parameters is the main goal of simulations, it seems that the axisymmetric simulation provides adequate accuracy, and the 3D simulation one is not the best choice. The 3D numerical simulation results in an in-depth study on the supersonic inlets, including the shock wave-boundary layer interaction, the location of the terminal normal shock wave, and consequently the separation point. For an axisymmetric supersonic inlet in an axisymmetric flow condition, 3D effects are not strong enough to have a significant influence on the inlet performance for all operational conditions. However, it seems that 3D effects play an important role in both critical and supercritical operating conditions during the steady state operation.

Построение туннельных осесимметричных сверхзвуковых воздухозаборников

Актуальность работы. Эффективность преобразования химической энергии топлива в механическое движение сверхзвукового летательного аппарата определяется потерями механической энергии на сопротивление летательного аппарата и потерями полного давления при торможении потока в воздухозаборнике воздушно-реактивного двигателя. Поэтому при прочих равных условиях более энергоэффективным будет тот летательный аппарат, у которого эти потери меньше. Существенное повышение энергоэффективности ожидается от применения осесимметричных изоэнтропических туннельных воздухозаборников. В рамках модели невязкого течения предлагается численная методика построения указанных воздухозаборников с профилированным центральным телом и цилиндрической обечайкой. Цель работы. В настоящее время отсутствуют способы профилирования осесимметричных изоэнтропических туннельных воздухозаборников, не имеющих элементов, выступающих за цилиндрический корпус летательного аппарата. Поэтому целью исследования является разработка методики построения таких воздухозаборников. Методы исследования. Используется численная реализация метода характеристик при условии изоэнтропичности течения. Предлагаемая методика состоит из двух задач. В первой задаче от угловой точки строится контур центрального тела с заданной ординатой точки фокусировки характеристик. Во второй задаче производится построение оставшейся части контура до другой угловой точки, при этом используется обращенное течение в кольцевом сопле с цилиндрической обечайкой. Одновременно с решением второй задачи определяется положение обечайки. Центральное тело может содержать на краях по угловой точке. При использовании промежуточной линии тока центральное тело будет гладким. Результаты. Создана методика для расчета семейства контуров осесимметричных туннельных сверхзвуковых воздухозаборников, геометрические характеристики которых однозначно описываются исходными данными.

Study of buzz phenomenon using visualization of external shock structure

An experimental study was carried out on an axisymmetric supersonic inlet with external compression in order to investigate the buzz phenomenon at different angles of attack and mass flow rates. The model was equipped with accurate and high-frequency pressure sensors, and the tests were conducted at Mach numbers varying from 1.8 to 2.5, for various angles of attack. Shadowgraph visualization technique, together with a high-speed camera, was used to provide the visual description of the shock structure in front of the inlet and to study the characteristics of buzz. Furthermore, pressure distribution over the spike surface was measured using several pressure sensors. Frequency of the buzz and shock displacement were measured by inspection of visualization pictures in each test. The obtained data from shadowgraphs were compared with those obtained from pressure measurements, and good agreement was found between them. The results revealed that for a moderate value of mass flow rate, the frequency of shock oscillation decreases as Mach number increases. Further, by increasing angle of attack, the shock displacement of oscillation will increase. At non-zero angles of attack, the displacement and frequency of shock motion show different behaviors on the leeward and windward sides of the body.

A Study on Blended Inlet Body Design for a High Supersonic Unmanned Aerial Vehicle

The design process of blended inlet body (BIB) for the preliminary design of a near-space high supersonic unmanned aerial vehicle (HSUAV) is presented. The mass flow rate and cowl area of inlet at a design point are obtained according to the cruise condition of the HSUAV. A mixed-compression axisymmetric supersonic inlet section with a fixed geometry reasonably matching the high supersonic cruise state is created by using the inviscid theory of aerodynamics. The inlet section is optimized and used as a baseline section for the BIB design. Three BIB concepts for the HSUAV are proposed, and their internal aerodynamic characteristics of inlet are evaluated using Euler computational fluid dynamics (Euler CFD) solver. The preferred concept is identified, in which the straight leading edge of the baseline HSUAV configuration is modified into the convex leading edge to accommodate the inlet and meet the requirements of the cowl area to capture the sufficient air flow. The total recovery of inlet for the preferred BIB concept and the aerodynamic characteristics of the modified HSUAV configuration are verified using Navier-Stokes computational fluid dynamics (NS CFD) solver. The validation indicates that the preferred BIB concept can meet both the requirements of the inlet and aerodynamic performance of the HSUAV.

Effects of Bleed Position on the Stability of a Supersonic Inlet

Effects of various boundary-layer bleed locations on the stability of an axisymmetric supersonic inlet has been investigated experimentally. The bleeds were located on the inlet compression cone and were designed to improve the stability margin of the inlet. The main objective of the study was to identify the boundary-layer bleed location and its effects on the improvement of the stability margin of the inlet. In addition, a buzz precursor detection methodology based on the rms level has been introduced. Experiments have been carried out on a mixed-compression inlet with a design Mach number of 2.0 and three different bleed locations at three freestream Mach numbers of 1.8, 2.0, and 2.2. All tests were conducted at an angle of attack of 0 deg. The results show that, based on the individual design of an inlet, a location can be found that the spilled normal shock stands at this location and consequently separation starts at the buzz onset. If the bleed slot is located in this place, the subcritical performance and stability margin of the inlet can be enhanced.

Simulations and Experiments of a Dual-Stream Low-Boom Supersonic Inlet

An axisymmetric supersonic inlet was designed to deliver high-quality subsonic flow to a turbofan engine, with a relaxed external compression centerbody to minimize the external overpressures. This study investigates the fluid dynamic performance and angle of attack, or pitch, sensitivity of two similar inlet geometries. The computations employ three-dimensional Reynolds-averaged Navier–Stokes simulations, while the experimental data were derived from the NASA Glenn Research Center at Lewis Field supersonic tunnel. Test conditions were based on a free-stream Mach number of 1.66 and pitch angles from to . Instrumentation includes stagnation pressure rakes in the core and bypass streams, as well as centerbody surface pressures. The two geometries were found to deliver high recovery flow to the engine face at zero pitch. At finite pitch angles, lower recovery flow was observed near the cowl, caused by the change in shock structure, but was successfully diverted to the bypass stream, and good levels of core stream recovery were maintained. The experiments showed a fuller boundary layer on the centerbody surface, while the outer core boundary layer was thinner than predicted. The measured core stream recovery at 5 deg pitch was 0.5–1.0% higher than in the simulations over a range of core flow rates.

Morphing Supersonic Inlet with Deforming Air Cell

A mixed-compression supersonic inlet with a variable throat can solve the contradiction of the starting capability at low Mach numbers and the total pressure recovery performance at high Mach numbers. However, the variation of the throat area is rather difficult for an axisymmetric supersonic inlet. A new method based on a deforming air cell to control the throat area of an axisymmetric inlet is brought forward and investigated by both experimental and computational methods. The morphing behavior of the deforming air cell in a supersonic wind tunnel is also obtained. The results indicate that the contraction ratio of the inlet can be effectively controlled by the cooperation of the air cell profile under different charging pressures and the cowl inner surface. As compared with a fixed geometry inlet, the total pressure recovery and the corresponding maximum backpressure of the morphing inlet are increased by 25.7 and 31.8%, respectively. In addition, the air cell can withstand the impact of strong oscillatory flows when the inlet experiences buzz.

Experimental Investigation of Effects of Mach Number on the Flow Instability in a Supersonic Inlet

An extensive wind tunnel tests were conducted on an axisymmetric supersonic inlet at Mach numbers from 1.8 to 2.2 and at different values of mass flow rates. Frequencies of the buzz were achieved from the pressure data as well as the high speed shadowgraph pictures. For each Mach number, two main frequencies for the buzz were obtained. The inlet at its design condition was stable, but when the mass flow rate was reduced, at first the shock wave started to oscillate with a small amplitude which is matched the Ferri criterion (little buzz). In this situation, both high- and low-frequency oscillations occurred; however, the high frequency one was dominant, but the oscillations seemed to be irregular. By further decreasing the inlet mass flow, the shock oscillations got a single low frequency and high amplitude, which in agreement with the Dailey criterion (big buzz). At a free stream Mach number of 2.2, the frequencies of these two kinds of instabilities are about 554 and 137 Hz, respectively. At this Mach number for moderate mass flow, an oscillation with high frequency/large amplitude and constant period was observed. The oscillation has the combined characteristics of the little buzz and the big buzz, which is named “added buzz” here. It is shown that at fairly constant mass flow, the frequency of the buzz is independent of the Mach number. It is further found that the buzz instability affects the external flow as well as the internal one with the same frequency.

An experimental study of buzz instability in an axisymmetric supersonic inlet

An experimental study was carried out on an axisymmetric supersonic inlet with external compression. The scope of this study was to investigate the general characteristics of the inlet buzz under various design and off-design conditions. The model was equipped with accurate and high frequency pressure sensors and the tests were conducted at Mach numbers varying from 1.8 to 2.5, at various angles of attack, and at different mass flow rates. Shadowgraph flow visualization, together with a high speed camera, was used to capture the external shock structure in front of the inlet. Frequencies of buzz were obtained from both the shadowgraph pictures and analysis of the pressure data. The amplitude of the shock wave motion was measured from the visualization pictures, too. The results show that for low mass flow rates, buzz frequency is almost independent of Mach number, while at a higher mass flow rate, the frequency of the shock oscillation increases with Mach number. Further, for each free stream Mach number, as the mass flow rate decreased, the amplitude of the shock movement grew, but its frequency decreased.

Numerical simulation of flows around axisymmetric inlet with bleed regions

A numerical simulation of flows in an axisymmetric supersonic inlet with bleed regions is performed. An existing code which solves the Reynolds Averaged Navier-Stokes equations and the two-equation turbulence model equations is converted into an axisymmetric code. In addition, a bleed boundary condition model has been applied to the code. In this paper, the modified code is validated by comparing numerical results against experimental data and other computational results for flows on a bump and over an oblique shock with bleed region. Using the code, numerical simulation is performed for the flows in an inlet with multiple bleed regions.

Performance Improvement of a Supersonic External Compression Inlet by Heat Source Addition

Heat source addition to the axisymmetric supersonic inlet may improve the performance parameters, which will increase the inlet efficiency. In this investigation the heat has been added to the flow field at some distance ahead of an axisymmetric inlet by adding an imaginary thermal source upstream of cowl lip. The effect of heat addition on the drag coefficient, mass flow rate and the overall efficiency of the inlet have been investigated. The results show that heat addition causes flow separation, hence to prevent this phenomena, roughness has been added on the spike surface. However, heat addition reduces the drag coefficient and the inlet mass flow rate considerably. Furthermore, the effects of position, size, and shape on the inlet performance were studied. It is found that the thermal source deflects the flow streamlines. By improper location of the thermal source, the optimum condition has been obtained. For the optimum condition, the drag coefficient is considerably reduced and the inlet mass flow rate and its efficiency have been increased slightly. The optimum shape of the heat source is obtained too. Keywords—Drag coefficient, heat source, performance parameters, supersonic inlet.

THE DEVELOPMENT OF FLOW DISTORTIONS AT THE FAN FACE FOR AN AXISYMMETRIC SUPERSONIC INLET

As a parallel effort to complement the experimental study on the aeroacoustics of an axisymmetric supersonic inlet, a computational study was conducted to provide a better understanding of the mechanism that generates the circumferential flow distortion at the fan face. A three-dimensional, compressible Navier–Stokes code using a finite volume method was used to calculate the flow field of the inlet with the struts located near the fan face. At 60% design fan speed, the computational results revealed a large region of three-dimensional separation on the centerbody downstream of the inlet throat. The influence of the struts on this separated flow resulted in a strong secondary flow pattern, creating a region of large circumferential distortion at the fan face near the centerbody. While the distortion caused by the strut wake shows significant reduction in a distance of half a strut chord downstream from the trailing edge of the strut, the distortion created by the secondary flow is much more persistent, and shows little sign of decay as it is converted downstream.

EFFECTS OF CHOKING ON THE AEROACOUSTICS OF AN AXISYMMETRIC SUPERSONIC INLET

An experimental study was conducted to determine the effect of choking on the aerodynamic and acoustic performance of a supersonic inlet. The investigated inlet was a prototype model of a mixed compression, axisymmetric supersonic inlet designed for the high speed civil transport aircraft. A 10·4 cm (4·1 in) turbofan engine simulator was used in conjunction with the inlet. The inlet was tested with the centerbody in the fully extended position at different fan speeds. Results show that “soft choking”, as characterized by a reduction in forward propagating fan noise, can occur when the Mach number at the inlet throat exceeds 0·5. In the forward sector (0°–60° from the inlet axis), the overall sound pressure level was reduced by about 7 dB as the fan speed increased from 50,000 to 70,000 rpm, due to the increase in the Mach number at the inlet throat. Additional comparison was made between the inlet configurations with the centerbody fully extended and fully retracted at a fan speed of 50,000 rpm. The results show that the higher Mach number at the inlet throat for the full retracted centerbody configuration was successful in reducing the overall sound pressure level by about 4 dB between 0° and 30° angular sector. While there is no measurable difference in the total pressure recovery for the two inlet configurations, there is, however, a significant increase in the circumferential flow distortion at the fan face for the higher throat Mach number test case.

Navier-Stokes solutions for mixed compression axisymmetric inlet flow with terminal shock

AN inlet flow with mixed external/internal compressions was simulated numerically for an axisymmetric supersonic inlet model by solving the three-dimensional, unsteady, compressible Navier-Stokes equations. A study was focused on the critical flow condition with the terminal shock positioned at the inlet throat. The computed results compared favorably with available wind-tunnel data. Contents For advanced supersonic aircraft, integrated design procedures capable of handling both the external and internal flowfields are required to treat a complete inlet system such as those shown in Fig. 1. The inviscid flow analysis of Ref. 1 indicates that viscous effects must be accounted for in the computation to achieve reasonable agreement with experimental data. For the critical flow condition, in particular, the treatment of boundary-layer bleed is needed to reduce the flow separation and to stabilize the terminal shock in the inlet throat. The present study extends the inviscid inlet flow analysis of Ref. 1 but focuses on the viscous effects on the inlet flow in the critical mode of operation and on the treatment of the boundary-layer bleed condition. A combined implicit (alternating direction scheme in the streamwise and radial directions) and explicit (in the azimuthal direction) approach of Ref. 1 is adopted to solve the unsteady, compressible Navier-Stokes equations enclosed by the Baldwin-Lomax turbulence model.2 The grid is divided into two zones of the H type covering the external (freestream) and internal (inlet) flowfields (Fig. 1). The computations were carried out in each zone interchangeably. The two zones are coupled on a common boundary along a line extending from the cowl lip to the upstream boundary. The no-slip condition for the viscous flow was imposed on the inlet surfaces. During the initial stage of time-iterativ e computation, the pressure at the inlet exit and the pressure differential across the bleed slot between the inlet and the plenum were raised gradually until the experimentally measured engine face pressure and inlet bleed-mass flow ratio were attained. The porous bleed holes of the experimental model were modeled geometrically by a single slot (Fig. 2) of the same length as the extent of the bleed holes. For the purpose of simulating the actual bleed through the porous holes, the gradient of pressure differential in the streamwise direction over the bleed slot was prescribed as the boundary condition and set equal to the experimental data.

Variable Geometry for Supersonic Mixed-Compression Inlets

Study of two-dimension al and axisymmetric supersonic mixed-compression inlet systems has shown that the geometry of both systems can be varied to provide adequate transonic airflow to satisfy the airflow demand of most jet engines. Collapsing geometry systems for both types of inlet systems provide a generous amount of transonic airflow for any design Mach number inlet system. However, the mechnaical practicality of collapsing centerbodies for axisymmetric inlet systems is doubtful. Therefore, translating centerbody axisymmetric inlets with auxiliary airflow systems to augment the transonic airflow capability are an attractive alternative. Estimates show that the capture mass-flow ratio at Mach number 1.0 can be increased approximately 0.20 for a very short axisymmetric inlet system designed for Mach number 2.37. With this increase in mass-flow ratio, even variablecycle engine transonic airflow demand can be matched without oversizing the inlet at the design Mach number.

A control system concept for an axisymmetric supersonic inlet.

A control system concept for an axisymmetric supersonic inlet.