An experimental investigation of hypervelocity flow in a conical nozzle

The flow in a conical nozzle is examined experimentally for a range of hypervelocity conditions in a free-piston shock tunnel. The pitot pressure levels compare reasonably well with an inviscid numerical prediction which includes a correction for the growth of the nozzle wall boundary layer. The size of the nozzle wall boundary layer seems to be well predicted by semi-empirical expressions developed for perfect gas flows, as do data from other free-piston shock tunnels.     

长试验时间爆轰驱动激波风洞技术研究

地面试验是先进高超声速飞行器研制的主要手段之一,获得满足高超声速气动实验研究的长时间高焓气流是发展激波风洞技术的关键难题之一.依据反向爆轰驱动方法,针对满足超燃试验有效时间的要求,讨论了爆轰驱动激波风洞运行缝合条件匹配、喷管起动激波干扰控制和激波管末端激波边界层相互作用等因素对激波风洞试验时间的制约及其相应的解决方法.应用这些延长试验时间的激波风洞创新技术,成功研制了基于反向爆轰驱动方法的超大型激波风洞,试验时间长达100ms,并有复现高超声速飞行条件的流动模拟能力.      

Transient three-dimensional startup side load analysis of a regeneratively cooled nozzle

The objective of this effort is to develop a computational methodology to capture the side load physics and to anchor the computed aerodynamic side loads with the available data by simulating the startup transient of a regeneratively cooled, high-aspect-ratio nozzle, hot-fired at sea level. The computational methodology is based on an unstructured-grid, pressure-based, reacting flow computational fluid dynamics and heat transfer formulation, and a transient inlet history based on an engine system simulation. Emphases were put on the effects of regenerative cooling on shock formation inside the nozzle, and ramp rate on side load reduction. The results show that three types of asymmetric shock physics incur strong side loads: the generation of combustion wave, shock transitions, and shock pulsations across the nozzle lip, albeit the combustion wave can be avoided with sparklers during hot-firing. Results from both regenerative cooled and adiabatic wall boundary conditions capture the early shock transitions with corresponding side loads matching the measured secondary side load. It is theorized that the first transition from free-shock separation to restricted-shock separation is caused by the Coanda effect. After which the regeneratively cooled wall enhances the Coanda effect such that the supersonic jet stays attached, while the hot adiabatic wall fights off the Coanda effect, and the supersonic jet becomes detached most of the time. As a result, the computed peak side load and dominant frequency due to shock pulsation across the nozzle lip associated with the regeneratively cooled wall boundary condition match those of the test, while those associated with the adiabatic wall boundary condition are much too low. Moreover, shorter ramp time results show that higher ramp rate has the potential in reducing the nozzle side loads.

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The application of coherent anti-Stokes Raman scattering to temperature measurements in a pulsed high enthalpy supersonic flow

Broadband single pulse coherent anti-Stokes Raman scattering (CARS) experiments employing a folded box phase matching geometry in a shock tunnel flow are presented. Rovibrational spectra of molecular nitrogen, produced at the exit of a pulsed supersonic nozzle for a range of flow enthalpies, are examined. Difficulties peculiar to the application of the optical technique to a high enthalpy pulsed flow facility are discussed and measurements of flow temperatures are presented. Theoretically calculated values for temperatures based upon algorithms used to determine shock tunnel flow conditions agree well with experimental measurements using the CARS technique.     

Numerical investigation of transient nozzle flow

Abstract. The starting process of two-dimensional and axisymmetric nozzle flows has been investigated numerically. Special attention has been paid to the early phase of the starting process and to the appearance of a strong secondary shock wave. For both cases, shock intensities and velocities are obtained and discussed. The flow evolution in the axisymmetric case is proved to be more complex and the transient starting process is slower than in the plane case. Finally, the effects of changing the nozzle angle and the incident shock wave Mach number on the transient flow are addressed. It is shown that a faster start-up can be induced either by decreasing the nozzle angle or increasing the Mach number of the incident shock wave. Received 16 November 2001 / Accepted 24 September 2002 / Published online 4 December 2002 Correspondence to:A.-S. Mouronval (e-mail: mouronv@coria.fr)     

Origin of flow asymmetry in planar nozzles with separation

An experimental investigation was conducted to study the mechanisms that lead to the origin of flow asymmetry in overexpanded planar nozzles, especially at low nozzle pressure ratios. Three Mach 2 planar nozzles with different divergent wall angles but same area-ratio were tested. For all three nozzles, a large portion of the dimensional pressure rise data across the separation shock shows the nature of boundary layer to be in the laminar/transitional state. Depending upon the local flow conditions, the flow can, therefore, experience either an early or a delayed separation on either wall. This can result in a free or a restricted shock separation condition on either wall which can initiate the beginning of flow asymmetry in nozzles at low nozzle pressure ratio. However, a higher nozzle wall angle was observed to prevent initiation of such a flow asymmetry. The present tests, therefore, indicate that in addition to the state of the boundary layer along the nozzle wall, the proximity of the separated shear layer to the nozzle walls also seems to play a dominant role in initiating conditions that favor the origin of flow asymmetry in nozzles. A significant drop in the shock unsteadiness levels was also indicated by increasing the wall angle.     

Modelling the complete operation of a free-piston shock tunnel for a low enthalpy condition

Only a limited number of free-stream flow properties can be measured in hypersonic impulse facilities at the nozzle exit. This poses challenges for experimenters when subsequently analysing experimental data obtained from these facilities. Typically in a reflected shock tunnel, a simple analysis that requires small amounts of computational resources is used to calculate quasi-steady gas properties. This simple analysis requires initial fill conditions and experimental measurements in analytical calculations of each major flow process, using forward coupling with minor corrections to include processes that are not directly modeled. However, this simplistic approach leads to an unknown level of discrepancy to the true flow properties. To explore the simple modelling techniques accuracy, this paper details the use of transient one and two-dimensional numerical simulations of a complete facility to obtain more refined free-stream flow properties from a free-piston reflected shock tunnel operating at low-enthalpy conditions. These calculations were verified by comparison to experimental data obtained from the facility. For the condition and facility investigated, the test conditions at nozzle exit produced with the simple modelling technique agree with the time and space averaged results from the complete facility calculations to within the accuracy of the experimental measurements.     

Reconstruction of velocity fields from wall pressure measurements in a shock wave/turbulent boundary layer interaction

We present here experimental results in a shock wave/turbulent boundary layer interaction at Mach number of 2.3 impinged by an oblique shock wave, with a deflection angle of 9.5°, as installed in the supersonic wind tunnel of the IUSTI laboratory, France. For such a shock intensity, strong unsteadiness are developing inside the separated zone involving very low frequencies associated with reflected shock motions.The present work consists in simultaneous PIV velocity fields and unsteady wall pressure measurements. The wall pressure and PIV measurements were used to characterize the pressure distribution at the wall in an axial direction, and the flow field associated. These results give access for the first time to the spatial-time correlation between wall pressure and velocity in a shock wave turbulent boundary layer interaction and show the feasibility of such coupling techniques in compressible flows. Linear Stochastic Estimation (LSE) coupled with Proper Orthogonal Decomposition (POD) has been applied to these measurements, and first results are presented here, showing the ability of these techniques to reproduce both the unsteady breathing of the recirculating bubble at low frequency and the Kelvin–Helmholtz instabilities developing at moderate frequency.     

Experimental study of shock tunnel flow with a stationary throat plug

A throat plug is a device to prevent fragments produced by the bursting of the primary diaphragm of a shock tunnel from entering the nozzle and damaging the model. An experimental study has been carried out to investigate the flow in the KAIST shock tunnel with two stationary throat plugs at primary shock velocities of 1.19 and 1.28 km/s. These shocks generate tailored conditions when helium and air are used as the driver and driven gases, respectively, at room temperature. Nozzle reservoir pressures and Pitot pressures at the nozzle exit are measured to examine the influence of the stationary throat plug on these properties. The throat plug that has an areal blockage of 6.2 % is found to have negligible effects on shock tunnel flow. However, the throat plug of 19.4 % blockage shows an appreciable influence by generating a pressure bump. Although it retards the establishment time of the nozzle flow, after the transient period, it does not cause deviation of the pressure from that for the clean nozzle configuration. It is found that the reduction of the cross-sectional area, which results from the presence of the throat plug, causes the pressure bump. By increasing the diameter of the driven tube in the region where the throat plug is located appropriately, the magnitude of the pressure bump is reduced significantly.     

Simulation of a complete reflected shock tunnel showing a vortex mechanism for flow contamination

Simulations of a complete reflected shock tunnel facility have been performed with the aim of providing a better understanding of the flow through these facilities. In particular, the analysis is focused on the premature contamination of the test flow with the driver gas. The axisymmetric simulations model the full geometry of the shock tunnel and incorporate an iris-based model of the primary diaphragm rupture mechanics, an ideal secondary diaphragm and account for turbulence in the shock tube boundary layer with the Baldwin-Lomax eddy viscosity model. Two operating conditions were examined: one resulting in an over-tailored mode of operation and the other resulting in approximately tailored operation. The accuracy of the simulations is assessed through comparison with experimental measurements of static pressure, pitot pressure and stagnation temperature. It is shown that the widely-accepted driver gas contamination mechanism in which driver gas ‘jets’ along the walls through action of the bifurcated foot of the reflected shock, does not directly transport the driver gas to the nozzle at these conditions. Instead, driver gas laden vortices are generated by the bifurcated reflected shock. These vortices prevent jetting of the driver gas along the walls and convect driver gas away from the shock tube wall and downstream into the nozzle. Additional vorticity generated by the interaction of the reflected shock and the contact surface enhances the process in the over-tailored case. However, the basic mechanism appears to operate in a similar way for both the over-tailored and the approximately tailored conditions.Communicated by R. R. Boyce     

Thrust shock vector control of an axisymmetric conical supersonic nozzle via secondary transverse gas injection

Transverse secondary gas injection into the supersonic flow of an axisymmetric convergent–divergent nozzle is investigated to describe the effects of the fluidic thrust vectoring within the framework of a small satellite launcher. Cold-flow dry-air experiments are performed in a supersonic wind tunnel using two identical supersonic conical nozzles with the different transverse injection port positions. The complex three-dimensional flow field generated by the supersonic cross-flows in these test nozzles was examined. Valuable experimental data were confronted and compared with the results obtained from the numerical simulations. Different nozzle models are numerically simulated under experimental conditions and then further investigated to determine which parameters significantly affect thrust vectoring. Effects which characterize the nozzle and thrust vectoring performances are established. The results indicate that with moderate secondary to primary mass flow rate ratios, ranging around 5 %, it is possible to achieve pertinent vector side forces. It is also revealed that injector positioning and geometry have a strong effect on the shock vector control system and nozzle performances.     

One method of profiling short plane nozzles

One of the possible methods is considered for profiling short plane nozzles for aerodynamic tubes. The nozzle has a straight sonic line, which allows the subsonic and supersonic sections to be constructed separately. The problem is solved numerically in the plane of a hodograph. In the subsonic region, Dirichlet's problem is formulated for Chaplygin's equation in a rectangle, one side of which is the sonic line. At the present time, two approaches have been defined in papers on calculations of a Laval nozzle, associated with the solution of the so-called “direct” and “inverse” problems (one has in mind a study of the flow in the interconnected region of sub- and supersonic flow). The direct problem determines the flow field in the case of a previously specified contour of the channel wall, the shape of which from technical considerations is obtained with certain geometry conditions. The direct problem can be applied in the construction of the Laval nozzle, if the contour of the inlet section of the channel (generally speaking, quite arbitrary) is chosen so successfully that neither shock compressions nor breakaway zones result in the flow. Although a strictly mathematical theory of the direct problem of the Laval nozzle is only being developed at present, there are still very effective numerical methods for its solution [1, 2]. In the inverse problem (which, by definition, is a problem of profiling), the contour of the nozzle is found with respect to a specified velocity distribution on the axis of symmetry. It is assumed that this quite arbitrary dependence can be selected from the condition of the absence of breakaway zones and shock compressions in the nozzle. By its formulation, the inverse problem is Cauchy's problem which, as is well-known, is incorrect in the classical sense in the ellipticity region — the subsonic section of the nozzle. At present, there are also efficient methods of solving the inverse nozzle problem [3], by interpreting it as an arbitrarily correct problem. Difficulties can arise in the inverse problem, in the provision of short (and, consequently, steep) nozzles because of the sharp increase of the error in the calculation. Together with the stated problems, a procedure can be evolved which is associated with the solution of the correctly posed problem for Chaplygin's equation in the plane of the hodograph. This approach is convenient in that it succeeds a priori in fulfilling the important condition of monotonicity of the velocity at the wall, ensuring (in the absence of shock compressions) nonseparability of the streamline flow at any Reynold's numbers.     

Simulation of the starting flow in a wedge–like nozzle

The starting process of the flow in a wedge-like expansion nozzle of a shock tunnel is simulated by an unsplit 2-D GRP scheme on an unstructured grid. The scheme is briefly outlined and results are presented and discussed in comparison to the experimental (shadowgraph) findings obtained by Amann. The simulated pattern of reflected and transmitted shock waves in the nozzle inlet region and inside the nozzle is found to agree well with the experimental data. Received 5 April 1996 / Accepted 16 June 1997     

The interaction of a shock wave with the boundary layer in a reflected shock tunnel

The influence of a nontotal reflection on the interaction of a reflected shock wave with the boundary layer in a reflected shock tunnel has been investigated. The calculating method of the velocity, the temperature and the Mach number profiles in the boundary layer in reflected shock fixed coordinates has been obtained. To account for equilibrium real gas effects of nitrogen, the numerical results show that the minimum Mach number in the boundary layer has been moved from the wall into the boundary layer with the increasing of the incident shock Mach number. The minimum Mach number, the shock angle in the bifurcated foot and the jet velocity along the wall to the end plate are reduced owing to the increasing of the area of nozzle throat. The numerical results are in good agreement with measurements.     

Investigation of the starting process in the shaped nozzle of a large-diameter shock tube

The characteristics of flow formation in the shock-wave starting process of a shaped nozzle with a developed subsonic section are analyzed on the basis of numerical calculations in the inviscid two-dimensional formulation. It is shown that in this case the use of steady-state boundary conditions in the throat section can lead to a significant error in determining the startup time. The results of the calculations are compared with experimental data obtained over a broad interval of the leading parameters. The results of the comparison help to explain the part played in the starting process by flow separation from the nozzle walls, which was not taken into account in the numerical investigation.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 5, pp. 88–95, September–October, 1986.The authors wish to express their thanks to R. I. Serikov and V. M. Khailov for the wind tunnel nozzle blowdown data and to V. P. Stulov for useful advice and discussions.     

双燃式超燃发动机冷态内流场的数值研究

研究了双燃式一体化通道(包含进气道、双燃式燃烧室和尾喷管)的冷态内流场特性.首次在激波风洞中对内流场进行纹影照相,用TVD格式求解三维全N-S方程对喷管和一体化通道进行分区数值模拟,并考察了几何参数对内流场的影响.结果表明对典型工况(h     

Behaviour of a shock train under the influence of boundary-layer suction by a normal slot

The interaction of a shock train with a normal suction slot is presented. It was found that when the pressure in the suction slot is smaller or equal to the static pressure of the incoming supersonic flow, the pressure gradient across the primary shock is sufficient to push some part of the near wall boundary layer through the suction slot. Due to the suction stabilized primary shock foot, the back pressure of the shock train can be increased until the shock train gradually changes into a single normal shock. During the experiments, the total pressure and therewith the Reynolds number of the flow were varied. The structure and pressure recovery within the shock train is analysed by means of Schlieren images and wall pressure measurements. Because the boundary layer is most important for the formation of a shock train, it has been measured by a Pitot probe. Additionally, computational fluid dynamics is used to investigate the shock boundary-layer interaction. Based on the experimental and numerical results, a simplified flow model is derived which explains the phenomenology of the transition of a shock train into a single shock and derives distinct criteria to maintain a suction enhanced normal shock. This flow model also yields the required suction mass flow in order to obtain a single normal shock in a viscous nozzle flow. Furthermore, it allows computation of the total pressure losses across a normal shock under the influence of boundary-layer suction.     

High enthalpy nozzle flow sensitivity study and effects on heat transfer

The sensitivity of the flow along the nozzle and in the test section of high enthalpy wind tunnels to the thermochemical response of the nozzle expansion process, as well as effects on the pressure and heat transfer distributions over the Electre blunt cone standard test model, are examined in the framework of properly characterizing the test section flow field in such facilities. Particularly sensitive to the thermochemical behaviour of the nozzle flow, in the facilities under consideration, are the static pressure, static temperature and Mach number, whereas stagnation point (pitot) pressure and heat transfer data or freestream velocity are inadequate for the characterization of the thermochemical state of the flow. The Electre and nozzle wall pressure data in the F4 arc jet wind tunnel suggest, in contrast to nonequilibrium computations, that the flow in the F4 nozzle is close to equilibrium. In the HEG and, to some extent, the T5 piston-driven shock tunnels, there are indications that significant heat losses occur in the reservoir. Lastly, simple semi-empirical formulations for stagnation point heating are shown to perform reasonably well in high enthalpy flow conditions.     

Shock structure in separated nozzle flows

In the case of high overexpansion, the exhaust jet of the supersonic nozzle of rocket engines separates from nozzle wall because of the large adverse pressure gradient. Correspondingly, to match the pressure of the separated flow region, an oblique shock is generated which evolves through the supersonic jet starting approximately at the separation point. This shock reflects on the nozzle axis with a Mach reflection. Thus, a peculiar Mach reflection takes place whose features depend on the upstream flow conditions, which are usually not uniform. The expected features of Mach reflection may become much difficult to predict, depending on the nozzle shape and the position of the separation point along the divergent section of the nozzle.      

Establishment of steady separated flow over a compression–corner in a free–piston shock tunnel

The time required to establish steady separated compression–corner flow is examined under hypervelocity conditions in a free-piston shock tunnel. This time is reasonably well described using previous perfect gas analyses. The results suggest that, provided the nozzle reservoir enthalpy is 20 MJ kg or less, there is sufficient time to establish steady separated flow before driver gas contamination becomes a significant problem in the present facility. Received 13 November 1996 / Accepted 20 May 1997