The Full Flowpath Analysis of a Hypersonic Vehicle

A numerical study has been carried out to investigate the full flow path and aerodynamic characteristics of a hypersonic vehicle at a 7.0 free stream Mach number. Results indicate that the inlet started and unstarted operations have remarkable effects on the flow pattern of the full flow path. When the inlet operates in a started mode, the transverse pressure gradient generated by the forebody alters the air captured characteristics and the entering flow quality of the inlet. Furthermore, the expansion process of the nozzle jet flow is obviously affected by the external flow field around the afterbody with the cross section shape transiting from a near rectangle at the exit of the nozzle to a near triangle at the tail of the vehicle. When the inlet operates in an unstarted mode, the aerodynamic instability can be observed in the full flow path of the vehicle. Due to the oscillation of the external compressed shock wave and nozzle jet flow, the aerodynamic characteristics of the vehicle vary periodically with the lift-drag ratio changing from 0.25 to 2.09. Finally, by comparing to the experimental data, the reliability of the CFD is verified.

Experimental study on surface arc plasma actuation-based hypersonic boundary layer transition flow control

Effective control of hypersonic transition is essential.In order to avoid affecting the structural proflle of the aircraft,as well as reducing power consumption and electromagnetic interference,a low-frequency surface arc plasma disturbance experiment to promote hypersonic transition was carried out in theΦ0.25 m double-throat Ludwieg tube wind tunnel at Huazhong University of Science and Technology.Contacting printed circuit board sensors and non-contact focused laser differential interferometry testing technology were used in combination.Experimental results showed that the low-frequency surface arc plasma actuation had obvious stimulation effects on the second-mode unstable wave and could promote boundary layer transition by changing the spectral characteristics of the second-mode unstable wave.At the same time,the plasma actuation could promote energy exchange between the second-mode unstable wave and other unstable waves.Finally,the corresponding control mechanism is discussed.

Analysis of Unstarted Hypersonic Flow Unsteadiness Based on Schlieren Image Processing

The unstarted flow field in a hypersonic inlet model at a design point of Ma 6 is studied experimentally.The time-resolved spatial flow characteristics of the separation shock oscillation,which is induced by the unstarted flow,are analyzed based on a high-speed Schlieren system and an image processing method.The motion of the separation shock detected by the shock-detection algorithm is compared to the results of fast-response wall-pressure measurements,and good agreement is demonstrated by comparing the frequency components in the power spectral density contours between shock oscillation and pressure fluctuation.The hysteresis of the pressure and separation shock during oscillation cycles is observed from the time history of the shock motion,which means that the unsteady flow pattern of the unstarted hypersonic flow can be accurately clarified by time-resolved Schlieren image processing.These results convincingly demonstrate that the shock-detection technique is successfully applied to an unstarted hypersonic flow case.

NON-STATIONARY EFFECTS IN HYPERSONIC NONUNIFORM DUSTY-GAS FLOW PAST A BLUNT BODY

In the framework of the two-fluid model, a hypersonic flow of a nonuniform dusty gas with low inertial (non-depositing) particles around a blunt body is considered. The particle mass concentration is assumed to be small, so that the effect of particles on the carrier phase is significant only inside the boundary layer where the particles accumulate. Stepshaped and harmonic nonuniformities of the particle concentration ahead of the bow shock wave are considered and the corresponding nonstationary distributions of the particle concentration in the shock layer are studied. On the basis of numerical study of nonstationary two-phase boundary layer equations derived by the matched asymptotic expansion method, the effects of free-stream particle concentration nonuniformities on the thermal flux, and the friction coefficient in the neighborhood of stagnation point are investigated, in particular, the most “dangerous” nonuniformity periods are found.

Investigation of hypersonic flows through a cavity with sweepback angle in near space using the DSMC method

Near space has been paid more and more attentionin recent years due to its military application value.However,flow characteristics of some fundamental configurations(e.g.,the cavity)in near space have rarely been investigated due to rarefied gas effects,which make the numerical simulation methods based on continuous flow hypothesis lose validity.In this work,the direct simulation Monte Carlo(DSMC),one of the most successful particle simulation methods in treating rarefied gas dynamics,is employed to explore flow characteristics of a hypersonic cavity with sweepback angle in near space by considering a variety of cases,such as the cavity at a wide range of altitudes 20-60 km,the cavity at freestream Mach numbers of 6-20,and the cavity with a sweepback angle of 30°-90°.By analyzing the simulation results,flow characteristics are obtained and meanwhile some interesting phenomena are also found.The primary recirculation region,which occupies the most area of the cavity,causes pressure and temperature stratification due to rotational motion of fluid inside it,whereas the pressure and temperature in the secondary recirculation region,which is a small vortex and locates at the lower left corner of the cavity,change slightly due to low-speed movement of fluid inside it.With the increase of altitude,both the primary and secondary recirculation regions contract greatly and it causes them to separate.A notable finding is that rotation direction of the secondary recirculation region would be reversed at a higher altitude.The overall effect of increasing the Mach number is that the velocity,pressure,and temperature within the cavity increase uniformly.The maximum pressure nearby the trailing edge of the cavity decreases rapidly as the sweepback angle increases,whereas the influence of sweepback angle on velocity distribution and maximum temperature within the cavity is slight.

Development of a Synchronization Method for Fluid-Thermal Study of Hypersonic Flow

A synchronization method is developed for the fluid-thermal study of hypersonic flow.Different from conventional loosely/tightly coupled methods which separately deal with the flow field and the structure temperature field,the presented method expresses the governing equations in a unified framework so that the two fields can be calculated simultaneously.For efficiently solving the unified equations,the finite volume method together with the dual-time stepping approach is employed.Like in the flow field,the local time step is also used in the temperature field,which is determined from thermal conductivity spectral radii.In order to treat the fluid-structure interface more conveniently,an expanded virtual boundary is introduced.For validation,several fluid-thermal hypersonic flow problems are simulated.The computed results are compared with those obtained from the coupled methods and the experiment.In the continuous heating problems,the stagnation temperatures predicted by both the coupled and synchronization methods are in good agreements with the experimental data.In the unsteady flowthermal hypersonic flows,the stagnation heat fluxes predicted by the presented method and tightly coupled method are basically the same,which agree better with the experimental data than those predicted by the loosely coupled method.In terms of prediction of the stagnation temperature,the synchronization method shows better accuracy than the tightly coupled method.

Prediction of hypersonic boundary layer transition on sharp wedge flow considering variable specific heat

When the air temperature reaches 600 K or higher, vibration is excited. The specific heat is not a constant but a function of temperature. Under this condition, the transition position of hypersonic sharp wedge boundary layer is predicted by using the improved eN method considering variable specific heat. The transition positions with different Mach numbers of oncoming flow, half wedge angles, and wall conditions are computed condition, the nearer to the Mach number The results show that for the same oncoming flow condition and wall transition positions of hypersonic sharp wedge boundary layer move much leading edge than those of the flat plate. The greater the oncoming flow the closer the transition position to the leading edge.

An Empirical Method for Prediction of Hypersonic Rarefied Flow-Field Structure

Numerical simulations are presented about the effects of gas rarefaction on hypersonic flow field.Due to the extremely difficult experiment,limited wind-tunnel conditions and high cost,most problems in rarefied flow regime are investigated through numerical methods,in which the direct simulation Monte-Carlo(DSMC)method is widely adopted.And the unstructured DSMC method is employed here.Flows around a vertical plate at a given velocity 7 500 m/s are simulated.For gas rarefaction is judged by the free-stream Knudsen number(Kn),two vital factors are considered:molecular number density and the plate′s length.Cases in which Kn varies from 0.035 to13.36 are simulated.Flow characters in the whole rarefied regime are described,and flow-field structure affected by Knis analyzed.Then,the dimensionless position D*of a certain velocity in the stagnation line is chosen as the marker of flow field to measure its variation.Through flow-field tracing and least-square numerical method analyzing,it is proved that hypersonic rarefied flow field expands outward linearly with the increase of 1/2Kn.An empirical method is proposed,which can be used for the prediction of the hypersonic flow-field structure at a given inflow velocity,especially the shock wave position.

Stable Runge-Kutta discontinuous Galerkin solver for hypersonic rarefied gaseous flow based on 2D Boltzmann kinetic model equations

A stable high-order Runge-Kutta discontinuous Galerkin(RKDG) scheme that strictly preserves positivity of the solution is designed to solve the Boltzmann kinetic equation with model collision integrals. Stability is kept by accuracy of velocity discretization, conservative calculation of the discrete collision relaxation term, and a limiter. By keeping the time step smaller than the local mean collision time and forcing positivity values of velocity distribution functions on certain points, the limiter can preserve positivity of solutions to the cell average velocity distribution functions. Verification is performed with a normal shock wave at a Mach number 2.05, a hypersonic flow about a two-dimensional(2D) cylinder at Mach numbers 6.0 and 12.0, and an unsteady shock tube flow. The results show that, the scheme is stable and accurate to capture shock structures in steady and unsteady hypersonic rarefied gaseous flows. Compared with two widely used limiters, the current limiter has the advantage of easy implementation and ability of minimizing the influence of accuracy of the original RKDG method.

NATURE OF THE SURFACE HEAT TRANSFER FLUCTUATION IN A HYPERSONIC SEPARATED TURBULENT FLOW

This paper presents the results of an experimental study of the unsteady nature of a hypersonic sepa- rated turbulent flow.The nominal test conditions were a freestream Mach number of 7.8 and a unit Reynolds number of 3.5x10^7/m.The separated flow was generated using finite span forward facing steps.An array of flush mounted high spatial resolution and fast response platinum film resistance thermometers was used to make mul- ti-channel measurements of the fluctuating surface heat trtansfer within the separated flow.Conditional sampling ana- lysis of the signals shows that the root of separation shock wave consists of a series of compression wave extending over a streamwise length about one half of the incoming boundary layer thickness.The compression waves con- verge into a single leading shock beyond the boundary layer.The shock structure is unsteady and undergoes large-scale motion in the streamwise direction.The length scale of the motion is about 22 percent of the upstream influence length of the separation shock wave.There exists a wide band of frequency of oscillations of the shock system.Most of the frequencies are in the range of 1-3 kHz.The heat transfer fluctuates intermittently between the undisturbed level and the disturbed level within the range of motion of the separation shock wave.This inter mittent phenomenon is considered as the consequence of the large-scale shock system oscillations.Downstream of the range of shock wave motion there is a separated region where the flow experiences continuous compression and no intermittency phenomenon is observed.

SEPARATED FLOW AROUND BENT-NOSE BICONIC IN HYPERSONIC FLOW

This article investigates the separated flow around the bent-nose biconic, in which the second ordered TVD(Total Variation Diminishing) method is applied. The governing cqua-tion is a thin layer Navier-Stokes equation. The surface pressure distribution at the leeward side is a little higher in the vicinity of the symmetrical plane. This seems to be caused by the development of the vortex at the leeward side due to the separated flow. The location of a separation line and the size are in very good agreement with the experiment. Also the secondary separation is captured very clearly.

Explicit Large Time Stepping with Third-Order Exponential Time Differencing Scheme for Hypersonic Chemical Non-Equilibrium Flow

In this paper, a third-order exponential time differencing scheme, named ETDRK3, was investigated for large time stepping in the computation of hypersonic non-equilibrium flow. The second-order Harten-TVD scheme was used for the spatial discretization. The efficient implementation of the scheme with diagonalization of Jacobin matrix was established and carried out for the semi-cylindrical around flow. Current observations showed that the numerical results were in good agreement with those obtained by the classical explicit three-stage Runge-Kutta scheme (RK3) and implicit LU scheme. Efficiency assessments promised the effectiveness of the ETDRK3 scheme. The rationality of the application of this scheme was proved by its preferable accuracy and efficiency.

Application of Nano Technique in Measuring Supersonic/Hypersonic Flow

Turbulence, universally exist in nature and human activities, is a kind of three-dimensional, irregular, unsteady flow. Ever since 19th century when people started to investigated turbulent flow technically, they have never dropped the po-tent and intuitionistic experimental method. Recently, with the development of aviation and aerospace industry, espe-cially with the increase desire of supersonic and hypersonic flight, the mechanism of high speed and compressible flow has become hot topic of fluid research, resulting in development of measurement method and technique. When encoun-tering compressible high flow, traditional techniques, such as schilieren, shadow and interference, cannot measure fine flow structures. Fortunately, multiple-discipline integration of nano technique, laser technique and imaging technique provides a new design for fluid measurement。Nano-tracer planar laser scattering (NPLS) is a new flow visualization technique, which was developed by the authors’ group in 2005, and it can visualize time correctional flow structure in a cross-section of instantaneous 3D supersonic flow at high spatiotemporal resolution. Many studies have demonstrated that NPLS is a powerful tool to study supersonic turbulence.

AN ANALYTIC SOLUTION ON HYPERSONIC FLOW OVER AN ARBTTRARY SLENDER BODY WITH NEAR POWER-LAW PROFILE

On the basis of a self-similar solution as well as of the assumption of the 'Transserse Motion',a general linear theory on hypersonic flow over a general slender body is set up in this paper By means of this theory, the problem concerned can he put into a universal system of O.D Eqs .which can be integrated manerically in adyance

Series Solution of Non-Similarity Boundary-Layer Flow in Porous Medium

This paper aims to present complete series solution of non-similarity boundary-layer flow of an incompressible viscous fluid over a porous wedge. The corresponding nonlinear partial differential equations are solved analytically by means of the homotopy analysis method (HAM). An auxiliary parameter is introduced to ensure the convergence of solution series. As a result, series solutions valid for all physical parameters in the whole domain are given. Then, the effects of physical parameters γ and Prandtl number Pr on the local Nusselt number and momentum thickness are investigated. To the best of our knowledge, it is the first time that the series solutions of this kind of non-similarity boundary-layer flows are reported.

Controlling secondary flow in high-lift low-pressure turbine using boundary-layer slot suction

The design of high-lift Low-Pressure Turbines(LPTs)causes the separation of the boundary layer on the suction side of the blade and leads to a strong secondary flow.This present study aims to minimize secondary losses through endwall slot suction and incoming wakes in a front-loaded high-lift LPT cascade with Zweifel of 1.58 under low Reynolds number of 25000.Two slotted schemes for the boundary layer of the endwall were designed(Plan A and Plan B),and the effects of suction mass flow on secondary flow were studied.The underlying physics of the endwall boundary layer of the suction and secondary flow under unsteady wakes was discussed.The results show that slot suction at the endwall boundary layer can significantly suppress the secondary flow by removing low-momentum fluids.Plans A and B significantly reduced the secondary kinetic energy by 44.2%and 36.9%,respectively,compared with the baseline cascade at the suction mass flow ratios of 1%.With an increase in the mass flow ratio of suction,the secondary flow was gradually reduced in both Plans A and B.It is more beneficial to control the secondary flow to destroy the intersection of the pressure side and suction side of the horseshoe vortex before it develops into a passage vortex.Under unsteady wakes,the combined effects of incoming wakes and endwall boundary layer suction can further suppress the secondary flow at the suction mass flow ratios of 2%for Plan A,because the positive and negative vorticity inside upstream wakes accelerated the mixing of the main flow and secondary flow and thus increased the energy of secondary vortices.

Experimental study on shock interaction control of double wedge in high-enthalpy hypersonic flow subject to plasma synthetic jet

The hypersonic shock-shock interaction flow field at double-wedge geometries controlled by plasma synthetic jet actuator is experimentally studied in a Ma = 8 high-enthalpy shock tunnel with the purpose of exploring a novel technique for reducing surface heat flux in a real flight environment. The results demonstrate that increasing the discharge energy is advantageous in eliminating the shock wave, shifting the shock wave interaction point, and shortening the control response time. The oblique shock wave can be completely removed when the actuator's discharge energy grows from 0.4 J to 11.5 J, and the displacement of the shock wave interaction point increases by 124.56%, while the controlled response time is shortened by 30 μs. Besides, the reduction in diameter of the jet exit is firstly proved to have a negative impact on energy deposition in a working environment with incoming flow, which reduces the discharge energy and hence decreases the control effect. The shock wave control response time lengthens when the jet exits away from the second wedge. Along with comparing the change in wall heat flux at the second wedge over time, the control effect of plasma synthetic jet actuator with and without inflation is also analyzed. When plasma synthetic jet works in inflatable mode, both the ability to eliminate shock waves and the shifting effect of the shock wave interaction point are increased significantly, and the wall heat flux is also reduced.

Aerodynamic map for soft and hard hypersonic level flight in near space

In this note, we design a velocity-altitude map for hypersonic level flight in near space of altitude 20-100 km. This map displays aerodynamic-related parameters associated with near space level flight, schematically or quantitatively. Various physical conditions for the near-space level flight are then characterized, including laminar or turbulent flow, rarefaction or continuous flow, aerodynamic heating, as well as conditions for sustaining level flight with and without orbital effect. This map allows one to identify conditions to have soft flight or hard flight, and this identification would be helpful for making correct planning on detailed studies of aerodynamics or making initial design of near space vehicles.

A reduced-order model for fast predicting ionized flows of hypersonic vehicles along flight trajectory

An improved Reduced-Order Model(ROM)is proposed based on a flow-solution preprocessing operation and a fast sampling strategy to efficiently and accurately predict ionized hypersonic flows.This ROM is generated in low-dimensional space by performing the Proper Orthogonal Decomposition(POD)on snapshots and is coupled with the Radial Basis Function(RBF)to achieve fast prediction speed.However,due to the disparate scales in the ionized flow field,the conventional ROM usually generates spurious negative errors.Here,this issue is addressed by performing flow-solution preprocessing in logarithmic space to improve the conventional ROM.Then,extra orthogonal polynomials are introduced in the RBF interpolation to achieve additional improvement of the prediction accuracy.In addition,to construct high-efficiency snapshots,a trajectory-constrained adaptive sampling strategy based on convex hull optimization is developed.To evaluate the performance of the proposed fast prediction method,two hypersonic vehicles with classic configurations,i.e.a wave-rider and a reentry capsule,are used to validate the proposed method.Both two cases show that the proposed fast prediction method has high accuracy near the vehicle surface and the free-stream region where the flow field is smooth.Compared with the conventional ROM prediction,the prediction results are significantly improved by the proposed method around the discontinuities,e.g.the shock wave and the ionized layer.As a result,the proposed fast prediction method reduces the error of the conventional ROM by at least 45%,with a speedup of approximately 2.0×105compared to the Computational Fluid Dynamic(CFD)simulations.These test cases demonstrate that the method developed here is efficient and accurate for predicting ionized hypersonic flows.

Hypersonic Shock Wave/Boundary Layer Interactions by a Third-Order Optimized Symmetric WENO Scheme

A novel third-order optimized symmetric weighted essentially non-oscillatory(WENO-OS3)scheme is used to simulate the hypersonic shock wave/boundary layer interactions.Firstly,the scheme is presented with the achievement of low dissipation in smooth region and robust shock-capturing capabilities in discontinuities.The Maxwell slip boundary conditions are employed to consider the rarefied effect near the surface.Secondly,several validating tests are given to show the good resolution of the WENO-OS3 scheme and the feasibility of the Maxwell slip boundary conditions.Finally,hypersonic flows around the hollow cylinder truncated flare(HCTF)and the25°/55°sharp double cone are studied.Discussions are made on the characteristics of the hypersonic shock wave/boundary layer interactions with and without the consideration of the slip effect.The results indicate that the scheme has a good capability in predicting heat transfer with a high resolution for describing fluid structures.With the slip boundary conditions,the separation region at the corner is smaller and the prediction is more accurate than that with no-slip boundary conditions.