Hypersonic Flow Field Research Articles

Research on Aeroheating of Hypersonic Reentry Vehicle Base Flow Fields

The structure of the base flow of a hypersonic reentry vehicle and the resulting base pressure and heat transfer have been studied by numerical study. The compressible Navier-Stokes equations are solved by the finite-volume method. SST k-ω turbulence model is used, and comparisons are made with flight test. Attention was focused on assessing the effects of angle of attack and Mach number. It was found that angle of attack can significantly alter the wake flow structure and reentry vehicle base pressure and heating distributions. The results of the simulation may provide a theoretical basis for the design of the thermal protection system of hypersonic reentry vehicles.

Receptivity characteristics of a hypersonic boundary layer under freestream slow acoustic wave with different amplitudes

An unsteady hypersonic blunt wedge flowfield under freestream slow acoustic wave is solved by direct numerical simulation (DNS), and the receptivity characteristics of the boundary layer under slow acoustic disturbance with different amplitudes are analyzed. Results show that slow acoustic wave in the freestream not only has a significant effect on the aerothermodynamics characteristics of hypersonic flowfield and boundary layer, but also on the bow shock characteristics. It shows that the maximum amplitudes of the disturbance modes in the boundary layer are positively correlated with the disturbance amplitude of the freestream. The bigger the amplitude of the freestream disturbance gets, the earlier the high-frequency disturbance mode exists, and the stronger the high frequency interference is. The larger amplitude of freestream disturbance wave can accelerate the generation of the higher order harmonic modes in boundary layer. Furthermore, the transformation position and the times of the dominant mode in the boundary layer differ under various amplitude of freestream disturbance, and compared with the smaller amplitude’s slow acoustic disturbance in the freestream, the larger amplitude disturbance can accelerate the first transformation of the dominant mode in boundary layer, and also delay its second transformation.

Numerical modeling of a glow discharge through a supersonic bow shock in air

The interaction between a glow discharge and the bow shock of a Mach 3 air flow around a truncated conical model with a central spike is modeled, and comparison is made with prior experimental results. The KRONOS workflow for plasma modeling in flow fields, which has recently been developed at ONERA, was used for the modeling. Based on the quasi-neutral approximation, it couples hypersonic and reactive flow fields with electron chemistry, including the effect of non-Maxwellian electron energy distribution function. The model used for the discharge involves 12 species and 82 reactions, including ionization, electronic and vibrational excitation, and attachment. The simulations reproduce the main features of the discharge observed experimentally well, in particular, the very recognizable topology of the discharge. It was found from the simulations that behind the bow shock, in the afterglow, the negative ion flow ensures the electrical conduction and the establishment of the glow discharge. The influence of kinetic rates on the voltage-current characteristics is discussed.

A wall grid scale criterion for hypersonic aerodynamic heating calculation

The theory of molecular motion was used to study the wall cell Reynolds number method, and it found that the height of the first grid point off wall is related to the molecular mean free path near wall. A wall grid scale criterion based on the molecular mean free path at the stagnation point was proposed by considering both the flow parameters and the wall temperature boundaries. The criterion was verified through numerical calculations and comparative analyses of the wall heat flux in different hypersonic flow fields. Research results draw the following conclusions: The optimal height of the first grid point off wall is a molecular mean free path. The influence of height of the first grid point off wall on aerodynamic heat is much greater than that of aerodynamic force. The criterion is not only effective for the wall heat flux predictions in air conditions, but also useful to that of non air environments. With the weakening of aerodynamic heating along the flow direction, the effect of wall grid scale on the wall heat flux gradually decreases.

Thermal Protection System Response Uncertainty of a Hypersonic Inflatable Aerodynamic Decelerator

The objective of this paper is to investigate the uncertainty in the bondline temperature response of a flexible thermal protection system subject to uncertain parameters in the hypersonic flowfield and thermal response modeling of a hypersonic inflatable aerodynamic decelerator configuration for ballistic Mars entry. An inflatable decelerator with a 10 m major diameter is selected for this study based on the forebody dimensions scaled from the 6 m test article that was tested in the NASA Ames Research Center’s National Full-Scale Aerodynamics Complex facility. A global nonlinear sensitivity analysis study for the bondline temperature uncertainty shows that the dimension of uncertain parameters can reduce from 22 to 8. An uncertainty analysis of the bondline temperature in the reduced dimensions indicates that the bondline temperature varies by as much as 125% above the nominal prediction and exceeds the temperature limit of 400°C. The largest uncertainty occurred at 70 s in the trajectory before separation of the inflatable decelerator for transition to a secondary descent technology. The main contributors to the bondline temperature uncertainty are the insulator and outer fabric conductivities, as well as the freestream density. The thickness and initial density of the insulator layer, closest to the gas barrier layer, are also shown to be significant contributors to the bondline temperature uncertainty, especially earlier in the trajectory.

Fabrication, calibration and proof experiments in hypersonic wind tunnel for a novel MEMS skin friction sensor

MEMS skin friction sensor, used to measure skin friction and its distribution, possesses smaller volume, higher sensitivity, better stability and other merits, and is very important for air vehicles. As we all know, normal load is much larger than skin friction measured in hypersonic flow field, thus MEMS skin friction sensor applied in hypersonic conditions presents higher requirements. Basing on MEMS technologies in existence and hypersonic flow field traits, a novel skin friction measuring method and meso-scopic sensor-head microstructure, which adopt that floating element even with measured wall and signal output microstructure being isolated from hypersonic flow field, are presented. Sensor-head design, weak capacitance signal detection, MEMS process technology, static calibration and proof experiments in hypersonic wind tunnel for sensor prototypes are implemented. The results of prototype tape-out and proof experiments indicate that: the process technology scheme that meso-scopic sensor-head being divided, fabricated separately and integrated by micro-assembly is feasible, the repeatability precisions of static calibration for prototypes are better than 1.5%; sensor prototypes present stable output, better return-to-zero, higher sensitivity, and repeatability precision better than 2% in proof experiments, and provide a preliminarily feasible experimental research method for skin friction measuring in hypersonic wind tunnel.

Stability Analysis of a Fully Coupled Implicit Scheme for Inviscid Chemical Non-Equilibrium Flows

AbstractVon Neumann stability theory is applied to analyze the stability of a fully coupled implicit (FCI) scheme based on the lower-upper symmetric Gauss-Seidel (LU-SGS) method for inviscid chemical non-equilibrium flows. The FCI scheme shows excellent stability except the case of the flows involving strong recombination reactions, and can weaken or even eliminate the instability resulting from the stiffness problem, which occurs in the subsonic high-temperature region of the hypersonic flow field. In addition, when the full Jacobian of chemical source term is diagonalized, the stability of the FCI scheme relies heavily on the flow conditions. Especially in the case of high temperature and subsonic state, theCFLnumber satisfying the stability is very small. Moreover, we also consider the effect of the space step, and demonstrate that the stability of the FCI scheme with the diagonalized Jacobian can be improved by reducing the space step. Therefore, we propose an improved method on the grid distribution according to the flow conditions. Numerical tests validate sufficiently the foregoing analyses. Based on the improved grid, theCFLnumber can be quickly ramped up to large values for convergence acceleration.

Improved quantitative visualization of hypervelocity flow through wavefront estimation based on shadow casting of sinusoidal gratings.

A simple noninterferometric optical probe is developed to estimate wavefront distortion suffered by a plane wave in its passage through density variations in a hypersonic flow obstructed by a test model in a typical shock tunnel. The probe has a plane light wave trans-illuminating the flow and casting a shadow of a continuous-tone sinusoidal grating. Through a geometrical optics, eikonal approximation to the distorted wavefront, a bilinear approximation to it is related to the location-dependent shift (distortion) suffered by the grating, which can be read out space-continuously from the projected grating image. The processing of the grating shadow is done through an efficient Fourier fringe analysis scheme, either with a windowed or global Fourier transform (WFT and FT). For comparison, wavefront slopes are also estimated from shadows of random-dot patterns, processed through cross correlation. The measured slopes are suitably unwrapped by using a discrete cosine transform (DCT)-based phase unwrapping procedure, and also through iterative procedures. The unwrapped phase information is used in an iterative scheme, for a full quantitative recovery of density distribution in the shock around the model, through refraction tomographic inversion. Hypersonic flow field parameters around a missile-shaped body at a free-stream Mach number of ∼8 measured using this technique are compared with the numerically estimated values. It is shown that, while processing a wavefront with small space-bandwidth product (SBP) the FT inversion gave accurate results with computational efficiency; computation-intensive WFT was needed for similar results when dealing with larger SBP wavefronts.

Effects of compressibility and Knudsen number on the aero optics in hypersonic flow fields

Tsien summarized the similarity in hypersonic flows, and related Knudsen number to Mach number and Reynolds number. Recently, a path-based problem, aero-optical effect, arises in hypersonic flows, and it concerns about the compressibility and the Knudsen number of the gas flows, which differs from the Tsien’s focus to some extent. In this paper, the similarity of hypersonic aero optics is theoretically studied, and both flow fields and induced aero-optical effect after flows pass through a cylinder are predicted by a well-accepted particle-based method, direct simulation Monte-Carlo (DSMC) method. The results show that the optical distortions are inversely proportional to the Knudsen number, while the compressibility plays an important role in the optical degradations. Hence, it is confirmed that the effects of Mach number and Knudsen number on the aero-optical effect induced by hypersonic flows are of great significance. Besides, since the Knudsen number is defined straightforwardly based on the optically active region, the physics is clearer than any other similarity criteria.

Flap effectiveness appraisal for winged re-entry vehicles

The interactions between shock waves and boundary layer are commonplace in hypersonic aerodynamics. They represent a very challenging design issue for hypersonic vehicle. A typical example of shock wave boundary layer interaction is the flowfield past aerodynamic surfaces during control. As a consequence, such flow interaction phenomena influence both vehicle aerodynamics and aerothermodynamics. In this framework, the present research effort describes the numerical activity performed to simulate the flowfield past a deflected flap in hypersonic flowfield conditions for a winged re-entry vehicle.

Rarefaction Effect on the Aerodynamic Performance of Scramjet Inlet in Hypersonic Flow

Accurate prediction of the performances of the scramjet inlet is very imperative to design the advance hypersonic vehicles which usually operate in a low-density fly environment. However, the majority of corresponding researches related to the inlet performance are limited in the continuum region by solving the conventional Navier-Stokes equation (NS solver). In this paper, the primary objective is to examine the rarefaction effect on the aerodynamic performance of scramjet inlet based on the parallel DSMC. First, both the DSMC and NS methods are compared by simulating the low-density hypersonic flow around a scramjet inlet which has been tested in the experiment before. It can be found that NS solver fails to capture the rarefaction flow structure, while the DSMC's solutions match the experimental data well. Then, several hypersonic flow fields with different degrees of rarefaction by changing the flight height are investigated by using DSMC. Some convincing quantitative numerical results related to the flow fields are obtained. The corresponding inlet performance's parameters, such as total pressure recovery coefficient, pressure ratio and flow coefficient, are also compared in details. The numerical results show that a thick viscous layer, which is one of the predominant features of this type of flow field, grows to a thickness almost larger than the height of duct passage with the increase of the rarefaction. Furthermore, shock-boundary layer interactions near the juncture move forward which will greatly affect the inlet performances. These conclusions are very useful to the design of future configuration of hypersonic scramjet inlets.

Computational Approach for Aeroheating with Thermally Coupled Fields

In this paper, a coupling methodology was proposed to solve simultaneously the external flowfield of a hypersonic reentry vehicle/capsule and the resultant heating on the vehicle’s structure and internal components with buoyancy-driven very low-speed flows resulting from conjugate heat transfer through the thermal protection system of the vehicle/capsule. Two flow solvers were employed to solve the hypersonic flowfield and the very low-speed flowfield, respectively. The coupling was done through the exchange of appropriate interface data between the two solvers. One of them was a density-based solver, and the other was a pressure-based solver. Validations of the individual solvers were done with two benchmark cases. The proposed coupling method was proved through simulating two conjugate heat transfer problems. Finally, an aerothermal–man–machine system exposed in a hypersonic flow was computed using the proposed coupling method.

Numerical Study of Spallation Phenomenon in an Arc-Jet Environment

It is hypothesized that spallation could affect the aerodynamic heating rates of reentry vehicles using ablative heat shields. To investigate spallation effects, a code is developed to compute the dynamics of spalled particles. The code uses a finite-rate chemistry model to study the chemical interactions of the particles with the flow field. The spallation code is one-way coupled to a computational fluid dynamics solver that models the hypersonic flow field around an ablative sample. Spalled particle behavior is numerically studied in both argon and air flow fields.

Large carbon cluster thin film gauges for measuring aerodynamic heat transfer rates in hypersonic shock tunnels

Different types of Large Carbon Cluster (LCC) layers are synthesized by a single-step pyrolysis technique at various ratios of precursor mixture. The aim is to develop a fast responsive and stable thermal gauge based on a LCC layer which has relatively good electrical conduction in order to use it in the hypersonic flow field. The thermoelectric property of the LCC layer has been studied. It is found that these carbon clusters are sensitive to temperature changes. Therefore suitable thermal gauges were developed for blunt cone bodies and were tested in hypersonic shock tunnels at a flow Mach number of 6.8 to measure aerodynamic heating. The LCC layer of this thermal gauge encounters high shear forces and a hostile environment for test duration in the range of a millisecond. The results are favorable to use large carbon clusters as a better sensor than a conventional platinum thin film gauge in view of fast responsiveness and stability.

S1910302 極超音速流れ場における非平衡原子・分子過程解析コード開発

S1910302 極超音速流れ場における非平衡原子・分子過程解析コード開発

Quantitative estimation of density variation in high-speed flows through inversion of the measured wavefront distortion

A simple method employing an optical probe is presented to measure density variations in a hypersonic flow obstructed by a test model in a typical shock tunnel. The probe has a plane light wave trans-illuminating the flow and casting a shadow of a random dot pattern. Local slopes of the distorted wavefront are obtained from shifts of the dots in the pattern. Local shifts in the dots are accurately measured by cross-correlating local shifted shadows with the corresponding unshifted originals. The measured slopes are suitably unwrapped by using a discrete cosine transform based phase unwrapping procedure and also through iterative procedures. The unwrapped phase information is used in an iterative scheme for a full quantitative recovery of density distribution in the shock around the model through refraction tomographic inversion. Hypersonic flow field parameters around a missile shaped body at a free-stream Mach number of 5.8 measured using this technique are compared with the numerically estimated values. (C) 2014 Society of Photo-Optical Instrumentation Engineers (SPIE)

Conjugate Heat Transfer Analysis for a Finite Thickness Cylinder in a Hypersonic Flow

Prediction of surface heating rates is of prime importance for the hypersonic flow regime. Experimental and conventional computational efforts overlook the heat transfer phenomenon in the solid due to the rigid assumptions involved in the solution methodologies. In order to address this fact, conjugate heat transfer (CHT) studies are carried out using various coupling techniques to examine their implementation abilities. Three types of solution methodologies are adopted, namely, decoupled, strongly coupled, and loosely coupled analysis. This study is also focused on looking into the effect of a hypersonic flow field on wall heat flux for a finite thickness insulating cylinder at moderately large time scales. Increase in wall temperature and decrease in surface heat flux have been noticed using strong and loose coupling techniques with an increase in simulation time. Decoupled fluid and solid domain analysis is found to be useful for typical shock tunnel test durations (∼1 ms) while investigations with loose coupling techniques are advisable for time scales corresponding to flight testing (∼1 s). Efforts are also made to reason the discrimination in prediction of stagnation point heat flux using conventional computational and experimental analysis.

Numerical Simulation on Air Mass Capture Control of Hypersonic Inlet by Laser Energy

Mechanism of hypersonic inlet performance promoting by “virtual cowl” induced by laser energy is introduced, and the physical model of interaction between laser energy and shocks in hypersonic flow field is established. Through comparing results in this paper with the work performed by Macheret et al.(Princeton University) and analyzing effects of laser energy deposited in hypersonic flow, numerical program and the feasibility of model are validated. At Mach 6, with no laser energy addition, air mass capture ratioKmis 71.87% in this paper, whileKmis 73.17% in the reference. Energy addition module can simulate the formation of shocks induced by laser energy, and the interaction process between laser energy and compression ramp shocks properly. Curves of mass capture ratioKmagrees well with that of the reference.Kmis a little higher than data in the reference, but the variable trend is consistent with each other when laser power is 1kW. When laser power is 0.5kW,Kmvaries a little, and the peak value is 77.36%, which only increases 7.64%, while the value is 7.28% in the reference.

Assessment of Real Gas Effects on Approximate and Boundary Layer Equations for Hypersonic Laminar Flow over Axisymmetric Bodies

Two efficient computational procedures based on the boundary layer equations and approximate relations areassessedin prediction of the laminar hypersonic flowfield for both the perfect gas and equilibrium air around the axisymmetric blunt body configurations. For the boundary layer procedure, the boundary layer equationsutilize the integral matrix solution algorithm for the blunt nose and after body region by using a space marching technique. The integral matrix procedure able us to create accurate and smooth results using the minimum grid in the boundary layer and minimize the computational costs. Applying the approximate method creates a robust and efficient code for heating calculations over the blunt bodies which flies in hypersonic regimes. These algorithms are highly appropriate to design of hypersonic reentry vehicles. The effects of real gas on the flowfield characteristics are also studied in two procedures.

S1910103 高次精度スキームを用いた大気圏突入カプセル後方部の強制遷移流れ場計算([S191-01]大気突入・減速技術(1),宇宙工学部門)

An aeroheating measurement test campaign on an Apollo capsule model with turbulent boundary layers was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. When flow became turbulent flow, heat flux on fore and rear body was 1.5-2 times larger than one in laminar flow. In order to reproduce the heat flux by numerical simulation, we constructed high order CFD code towards LES in hypersonic flow. Since robustness near shock and high resolution in boundary layer is needed for LES in hypersonic flow, we employed recently improved WENO method. We calculated the hypersonic flow (M_∞ = 17) around cylinder and investigated the robustness for strong shock by this method. We could obtain smooth pressure distribution and good agreement with the calculated heat flux by NASA LAURA code. The hypersonic flow field around HTV-R, which is a manned space capsule under development by JAXA was also calculated towards predicting heat flux on after body.