Hypersonic Gun Tunnel Research Articles

Start-Up Analysis of a Hypersonic Short-Duration Facility

This paper analyzes the nozzle start-up of the von Kármán Institute Longshot, an impulse hypersonic gun tunnel, using a combination of the Lagrangian L1D code (used to model the initial compression process of the test gas into a reservoir by an inertial piston) and unsteady Reynolds-averaged Navier–Stokes simulations to model the flow expansion through the appended nozzle. The influence of initial nozzle flow conditions on the duration of the start-up is investigated based on various simulated experiments, and it is verified that the shortest establishment time is obtained for the lowest initial pressure. The applicability of the numerical scheme is proved through a comparison of the inlet and outlet total and static conditions against experimental results at relevant conditions. Besides, the right running and left running shocks constituting the start-up shock system are tracked through the nozzle, and their propagation velocities are reported. Momentum and thermal boundary-layer establishment times are reported based on the temperature and velocity profiles predicted at the nozzle exit. Finally, the flow establishment time is compared against several criteria available in the literature. Additionally, new establishment and quasi-steady behavior criteria are presented, derived from the freestream static pressure and temperature evolution.

Experimental study on the influence of attitude angle on the aero-optical effects of a hypersonic optical dome

The aero-optical effects of a hypersonic dome with cooling jet device can be mitigated by micro vortex generators (MVGs) at 0°angle of attack (AoA) and 0°angle of sideslip (AoS). However, the attitude angle of an aircraft is constantly changing during actual flight. In this research, experiment was carried out at the KD-01 hypersonic gun tunnel at Mach 6.0. Using the schlieren technique, the flow field density distribution of a hypersonic optical dome under different pressure ratios of jet (PRJ) was observed. The optical path difference (OPD) was obtained by background-oriented schlieren (BOS). Experimental results showed that a 4°AoA could increase the OPD value more obviously with or without MVGs; and that the use of MVGs under 4°AoS would aggravate the aero-optical effects. These findings place new demands on the design and installation of MVGs.

Toluene-based planar laser-induced fluorescence imaging of temperature in hypersonic flows

Planar laser-induced fluorescence imaging is carried out in a hypersonic gun tunnel at a freestream Mach number of 8.9 and Reynolds number of \(47.4 \times 10^6\,\hbox {m}^{-1}\) (\(N_2\) is the test gas). The fluorescence of toluene \((C_7H_8)\) is correlated with the red shift of the emission spectra with increasing temperature. A two-colour approach is used where, following an excitation at 266 nm, emission spectra at two different bands are captured in separate runs using two different filters. Two different flow fields are investigated using this method: (i) hypersonic flow past a blunt nose, which is characterised by a bow shock with strong entropy effects, and (ii) an attached shock-wave/boundary-layer interaction induced by a flare located further downstream on the same blunt cylinder body. Measurements from as low as the freestream temperature of \(68.3\) K all the way up to \(380\) K \((T_{\infty }-5.6T_{\infty })\) are obtained. The uncertainty at the higher temperature level is approximately \(\pm 15\) %, while at the low end of the temperature, an additional \(\pm 15\) % uncertainty is expected. Application of the technique is further challenged at high temperatures due to the exponentially reduced fluorescence quantum yields and the occurrence of toluene pyrolysis near the stagnation region (\(T_\mathrm{o}=1150\) K). Overall, results are found to be within \(10\) % agreement with the expected distributions, thus demonstrating suitability of the technique for hypersonic flow thermometry applications in low-enthalpy facilities.

Flow control effectiveness of jets, strakes, and flares at hypersonic speeds

In the present article, the effects of three important flow control techniques, i.e. control by strakes, jets, and flares, on the aerodynamic characteristics and performance of a number of body configurations at high speeds are reviewed. The results are based on previous studies conducted by the author in a hypersonic gun tunnel at a Mach number of 8.2 and a Re∞/ m=9 200 000. Air was used as the working gas. The tests employed schlieren photography to study the overall flow field. Quantitative studies have been made by pressure measurements. The forces were measured with a three-component balance equipped with semiconductor strain gauges. The surface flow was studied using the oil-dot technique.

極超音速流中のランプ誘起衝撃波/遷移境界層干渉の実験的研究

An experimental investigation of a shock-wave/transitional-boundary-layer interaction was conducted on a flat plate with an aft-mounted ramp in a hypersonic gun tunnel. Test Mach number was 10, the Reynolds number range ReL based on the model plate length was from 1.1 to 3.2 × 105, the attack angle of the model was zero, and the ramp deflection angle θw was varied from 10 to 30deg in increments of 2. The effects of ReL and θw on the characteristics of the interaction region were investigated through both schlieren visualization and heat transfer measurement. The state of reattaching boundary layers were determined from these data. The following is obtained. A model that describes how the upstream influence of a shock-wave/boundary-layer interaction scales with the parameters of the incoming flow agrees with laminar data at small ReL-1/8Δθw, where Δθw is the incremental angle after incipient separation. The slope of the characteristic curve of the upstream influence is dependent on the type of the interaction. ReL-1/8 Δθw where the deviation from the model occurs and flow unsteadiness appears is found to become small with increasing ReL.

グロー放電トレーサー法を用いた極超音速圧縮コーナー流の研究

In this report, we apply high frequency impulsive glow discharge technique for obtaining the flow pattern upstream of a compression corner. For purpose of comparison, schlieren visualization using high speed video camera is also performed. These experiments are carried out in a hypersonic gun tunnel at Mach number 10 and Reynolds number based on the frontal plate chord length 2.1×105. The corner angles studied are in the range from 0 to 30deg at zero incidence. Time lines are obtained successfully even for separated flow cases and the boundary layer edges visualized by the discharge-tracer technique correlate well with those defined through the schlieren images. The extent of the upstream influence becomes large with increasing corner angles and the separated boundary layer is transitional to turbulent flow for the largest angle case tested. Free interaction similarity using Newtonian theory is presented and it seems reasonable with respect to the present experimental results.

A119 制御舵面を過ぎる極超音速流の挙動

Understanding of hypersonic viscous interaction phenomenon over control surfaces is of particular impor-tance for designers of hypersonic vehicles. An experimental study of the behaviour of hypersonic flow past flaps is presented in this report. Impulsive glow discharge technique and Schlieren visualization using high speed video camera are applied to examining the structure upstream of the hinge. These experiments are conducted in a hypersonic gun tunnel at Mach number 10 and at Reynolds numbers based on the chord length ranging 0.8 - 2.12 × 105. The flap deflection angles studied are in the range from 0 to 0.523 rad at zero incidence. Numerical simulations based on the Navier-Stokes equations are also performed. Time lines obtained by the dis-charge technique are reasonable and the structure evaluated by the time lines correlates well with those through Schlieren and computational results. Normalized upstream influence increases with increasing Reynolds number and increasing flap angle in the range studied.

Measurements of Capsule Wake Stabilization Times in a Hypersonic Gun Tunnel

The wake stabilization time behind a hypersonic vehicle is very important in experiments focused on the nearwake studies in pulsed facilities at high-enthalpy and high-Mach-number conditions. Observations of such wake structure have been very difficult to obtain, and therefore, the experimental discussion about the stabilization time itself has also been elusive. The wake structure stabilization time was formulated by using dimensional analysis and the present experimental results. The experiments were carried out by observing free shear layers for capsule models similar to the Mars Environmental Survey Pathfinder probe. To visualize the wake structure, an electrical discharge method was used. From these investigations, it was found that the stabilization time of the wake structure was related to the Reynolds number, the model size, and the freestream velocity. A Mach 10 gun tunnel was used in these experiments. Nomenclature A = projected area of model D = capsule maximum body diameter k = proportionality constant Rb = capsule maximum body radius R f = distance between free shear layer and symmetry axis Re =R eynolds number T = stabilization time of wake pattern measured from the beginning of stable uniform flow T0 = stabilization time of wake pattern measured from the beginning of flow obtained by a pulsed facility T1 = elapsed time from the beginning of flow obtained by a pulsed facility T2 = stabilization time of freestream measured from the beginning of flow U = freestream velocity α = constant µ = viscosity coefficient ρ = freestream density

Jet Control Effectiveness Studies on a Flat-Plate Body at Hypersonic Speeds

The effects of jet control on the aerodynamic characteristics and performance of a flat-plate body configuration have been studied experimentally. The study has been carried out in a hypersonic gun tunnel at a Mach number of 8.2 and a Reynolds number of 1,900,000, based on the flat-plate length-chord, using both sharp and blunted leading edges, at pitch angles 0, −5 and −10 deg. Air was used as the working gas for both the free stream and the sonic jet. The tests employed schlieren photography to study the overall flow field. Quantitative studies of the effect of the jet have been made by pressure measurements. The results were compared with numerical simulations performed using a 3-D time-marching Navier-Stokes code. Theoretical algorithms have been developed to predict the shape of the separation front and jet penetration height. The parameters, which determine the upstream extent and lateral spreading of the separation front around the transverse jet, and the magnitude of the separated region, are detailed.

B132 極超音速圧縮コーナー流の可視化

Understanding of hypersonic viscous interaction phenomenon over control surfaces and engine intakes is of particular importance for designers of hypersonic vehicles. However, experimental data of velocity fields around the compression corner which is an idealized configuration of these practical situations are sparse chiefly because it is difficult to obtain such data owing to short duration, high velocity, and rarefaction of the test flow in hypersonic ground facilities. In this report, we shall try to apply the high frequency impulsive glow discharge technique to obtaining the hypersonic velocity fields. The experiment is carried out in a hypersonic gun tunnel at Mach number 10 and Reynolds number based on the frontal plate chord length 2.8×105. Time lines are obtained successfully, even for separated flow cases.

Study on Stabilization Time of Wake Structure Behind Hypersonic Vehicles by Electrical Discharge Method

A stabilization time of a wake structure behind hypersonic vehicles is very important in experiments on the flow structure by using pulsed facilities of high enthalpy or of high Mach number. However, observations themselves of such wake structure have been very difficult, and therefore, the experimental discussion about the stabilization time itself has also been very difficult. In this study the stabilization time of the wake structure was formulated by using the dimensional analysis and the present experimental results. The experiment was carried out by observing separation points and free shear layers just after three re-entry capsule models. To visualize the wake structure, one of techniques called the electrical discharge method was utilized. From these investigations it was found that the stablization time of the wake structure was related to the Reynolds number, the model scale, and the freestream velocity. A hypersonic gun tunnel of Mach 10 was used in this experiment.

極超音速で再突入するＨＯＰＥ‐Ｘ周りの流れ場の放電発光法による研究

The shock shapes, boundary layers, and temperature layers around a model of HOPE-X traveling at a hypersonic speed were visualized utilizing the electric discharge method. The experiments were carried out under the condition that the angles of attack were 0°, 15°, 30°, and 45°. To visualize 3-D shock shapes around the model, the observation of shock shapes were conducted from the side and back of the model. From the experimental results, it was confirmed that the two kinds of shock shapes interacted with each other. The various kinds of shock shapes according to the model angles of attack were also obtained. Subsequently, the visualized results of temperature layer around the model were described. Furthermore, the boundary layers generated around the model were visualized and compared with the temperature layers. A hypersonic gun tunnel was used in these experiments. The main characteristics of the tunnel were the Mach number was 10 and the duration was 10ms.

Gun tunnel flow calibration: defining input conditions for hypersonic flow computations

The procedures used to calibrate a hypersonic gun tunnel nozzle flow are described. The values obtained using these methods are then used as input conditions for computations of hypersonic flow over a long slender test body. Only by completely specifying the flow field, including the effects of nozzle flow angularity, can the best agreement between experiment and computations be achieved.

Incipient separation on flared bodies at hypersonic speeds

AbstractThe aerodynamic effectiveness of a cylinder-flare body at zero incidence under laminar and turbulent boundary-layer conditions has been studied experimentally. Two nose geometries, namely a 10° half-angle sharp cone and a hemisphere, were used. The study has been carried out in a hypersonic gun tunnel at a Mach number of 8·2 and a Reynolds number of 158,100, based on the cylinder diameter, at flare angles 0°, 10°, 20°, 30° and 45°. The surface flow was studied using oil-dot and liquid crystal techniques. Some information regarding the shock layer was obtained from schlieren pictures. The effects of entropy layer and boundary-layer state on flare effectiveness were deduced from pressure measurements over the cylinder and the flare.

C404 極超音速流中における衝撃波と境界層との干渉に関する研究

Shock-wave boundary-layer interaction flow field near the sharp leading edge of a flat plate is investigated. Measurements of impact pressure distributions and Schlieren visualization are made under the flow conditions: M∞=10, 3<x<13 in a hypersonic gun tunnel. Preliminarily, a shape of a Pitot tube is improved to avoid unfavorable disturbance to flow field and to measure accurately for the short duration. The relationship between the measured impact pressure distributions and the structure of the flow field is discussed. The inviscid layer seems to be merged into the viscous layer at x= {-.7em} aisebox{1.1ex}{.} {.1em} aisebox{-0.2ex}{.} 10.

次世代スペースシャトルの試験機Ｘ‐３３周りの流れの放電発光法による可視化

Visualizations of flowfield around the X-33 test vehicle have been performed utilizing the electric discharge method. The electric discharge method has been invented and developed by the authors and others for visualizing three dimensional hypersonic flowfield phenomena. In this experiment the objectives of visualization are three dimensional shock shapes, temperature distributions, and flow patterns near the model surface. It is considered that these visualized results are useful for verifying the theoretical and numerical results of the flowfield around the model. These visualizations are carried out using the hypersonic gun tunnel whose main characteristics are that the Mach number is 10 and the duration is 10ms.

極超音速流中の衝撃波／境界層干渉場に関する放電発光法による研究

Investigations on the spatial flowfield of hypersonic shock-wave/boundary-layer interaction have been performed by utilizing the electric discharge method. The model geometry used in this experiment is composed of a flat plate and a fin. In general, visualizations of such spatial flowfield were very difficult or impossible. However, in these investigations utilizing the electric discharge method which is the method for visualizing three dimensional hypersonic flowfield phenomena, the spatial flowfield phenomena could be visualized qualitatively. The experiment is carried out using a hypersonic gun tunnel whose main characteristics are that the Mach number is 10 and the duration is 10ms.

Hypersonic control flap effectiveness

SummaryThe effects of flap deflection, configuration incidence and leading edge bluntness on boundary layer separation and transition have been studied experimentally. A quasi two-dimensional flat plate equipped with a full span trailing edge control flap has been employed for these tests.The studies have been carried out in a hypersonic gun tunnel at M∞ = 8·2 and Re∞/cm = 9·0 × 104. The flap deflection angles studied were in the range 0° ≤ β ≤ 30°. The incidence range was from zero to α = 10° (positive α is nose down). Leading edge bluntness effects were simulated by the introduction of a hemi-cylindrical leading edge.The flow structure was studied using high speed Schlieren photography as well as surface pressure and heat transfer measurements. Liquid crystals were employed to study the threedimensionality of the flow structure for selected configurations. Analytical theories have been developed to estimate the flap pressure and heat transfer levels for the sharp and blunt configurations. These are compared against experimental measurements.

Interactions between shock wave and hypersonic laminar boundary layer near convex corner

A hypersonic gun tunnel was used to investigate a Mach 8 laminar boundary layer interacting with an externally generated shock wave impinging in the neighborhood of a convex corner. Surface static pressure measurements and schlieren photographs were taken from planar models with corner angles of 5 and 10 deg;the shock generator was set to the same angles. The results are compared with data obtained in the same tunnel for an externally generated shock impinging on a flat-plate boundary layer, and for a boundary layer interacting with a cornerexpansion wave. When separation and reattachment both occur either upstream or downstream of the corner, the 5- and 10-deg interactions have similar properties; however, they differ markedly when separation is upstream and reattachment is downstream. The results also suggest that a convex corner will not completely suppress separation caused by an externally generated shock, even when impingement is close to the corner. However, the extent of separation is minimized when the separation point is located at the corner, corresponding to shock impingement 2-4 boundary-layer thicknesses downstream of the corner. Nevertheless, the overall interaction length still extends over many boundary-layer thicknesses.

New method for measurement of surface pressure using magnetic tape

A new method of surface pressure measurement has been developed using magnetic tape as a sensor—instead of the usual strain gage, piezogage or semiconductor —within a diaphragm-type pressure transducer. The method is based on the idea that a pressure value applied to the diaphragm is related to the deflection of the diaphragm, and the deflection can be related to the value of magnetization strength sensed in the tape. The merit of the new method is that it can be easily applied for measurement of pressure distributions on model surfaces because many pressure transducers are contained in a single sheet only about 1-mm thickness. The proposed method can be used in the high-temperature, short-duration measurement conditions, and so it is well suited to hypersonic wind-tunnel use. The sheet can be bonded onto a model surface. In this study, experiments using the new pressure sensor were performed in a hypersonic gun tunnel at a Mach number of 10 for a duration time of 10 ~2 s. Pressure distributions around two models in hypersonic flow were measured successfully.