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The recent trend of using Graphics Processing Units (GPUs) for high performance computations is driven by the high ratio of price performance for these units, complemented by their cost effectiveness. Such kinds of units are increasingly being deployed not only as accelerators for supercomputer installations, but also in GPU-enabled nodes in Grid and Cloud installations. At first glance computational fluid dynamics (CFD) solvers match perfectly to GPU resources, because these solvers make intensive calculations and use relatively small memory. Nevertheless, there are scarce results about the practical use of this serious advantage of GPU over CPU, especially for calculations of viscous, compressible, heat conductive gas flows with double precision accuracy. In our work we present calculation of unsteady, viscous, compressible and heat conductive gas with double precision accuracy using GPU-enabled version of the algorithm SIMPLE-TS, written on standard OpenCL. As a test case we model the flow past a square in a microchannel at supersonic speed with Mach number \(\mathrm{{M}}=2.43\) on AMD Radeon HD 7950 GPU and achieve 90 GFlops, which is 46 times faster than the CPU serial code run on Intel Xeon X5560.

We present an exact analytic solution of a nonlinear Schrodinger field interacting with a moving potential (obstacle) at supersonic speed. We discover conditions under which the field can form a stable shape-invariant structure localized around the obstacle-a dressing effect that protects the field against excitations by the obstacle. Such an effect demonstrates the existence of frictionless motion beyond the conventionally defined critical velocity.

This report describes a computational study to understand the unsteady free-flight aerodynamics and flight dynamics of a finned projectile with jet control. Numerical simulations have been performed for this projectile with jet maneuver using an advanced coupled computational fluid dynamics (CFD)/rigid body dynamics (RBD) technique. The coupled CFD/RBD method allows time-accurate virtual fly-out simulations of projectiles and simultaneously predicts the aerodynamics and the flight dynamics in an integrated manner. The control is provided by a transient pulsed jet that is activated only once in the beginning for a fixed duration and then turned off for the remaining trajectory computation. Computed results show the actual time-dependent response of the flight vehicle and the resulting unsteady aerodynamics and flight dynamics with jet control for a variety of conditions.

We conducted flow visualization test with liquid crystal film to obtain the characteristics of boundary layer transition on a three-dimensional warped wing at supersonic speed using NAL 2m*2m transonic wind tunnel. Present paper describes how we selected suitable liquid crystal film for our test condition, how we judged transition position from color pattern of the film and some results of present test.

Bulgarian. Russian, English summary, [Zapryanov Zapryan; Запрянов Запрян]

平成29年度衝撃波シンポジウム (2018年3月7日-9日. 東北大学片平さくらホール ）仙台市, 宮城県, (2018年3月7日-9日. 宇宙航空研究開発機構角田宇宙センター ）角田市, 宮城県, Symposium on Shock Waves in Japan, FY2017 (March 7-9, 2018. Sakura Hall), Sendai, Miyagi, (March 7-9, 2018. Kakuda Space Center, Japan Aerospace Exploration Agency (JAXA)(KSPC)), Kakuda, Miyagi, 資料番号: SA6000119147, レポート番号: P-17【若】

Moving mass control implements the maneuver control of vehicle through moving the movable slide inside vehicle in order to move the mass center position. In this paper, take missile as an example, based on the derivation of six degree of freedom (6-dof) model of mass moment missile, combined with the law of parameter variation of aerodynamic and speed during missile flight, combined with virtual simulation technology, to establish a virtual reality simulation for high supersonic speed vehicle's control of moving mass model, and provide necessary foundation for the next further study of moving mass control.

The structure of the shock layer and the abrupt aerodynamic forces acting on a sphere moving in an atmosphere with inhomogeneous inclusions of various shapes are considered.

The unsteady flow in the neighborhood of the stagnation line on a sphere traveling at supersonic speed through a plane layer of diatomic gas with elevated temperature and nonequilibrium excitation of the molecular vibrations is investigated. (The source of the inhomogeneity could be a gas discharge [1].) The problem is solved using the viscous shock layer model which makes it possible to take molecular transport processes into account and analyze the unsteady heat transfer. Such flows were previously calculated in [2] within the framework of the inviscid gas model.

Pulse infrared laser launched by the prepositional laser fuse system reaches the target first. Laser echo pulse signal then arrive the surface of photodiode laser detection receiver, and penetrates through the boundary layer - shear layer - shock wave - atmosphere - shock wave - shear layer - boundary layer under supersonic velocity, finally, focuses on the photosensitive surface. As shock wave has an effect on the prepositional laser fuse in the optical path, a great deviation has been brought between distance determining methods on static state and in the course of the actual flight. In this paper, a model of the shock wave's influence on the prepositional laser fuse's optical distance under supersonic velocity is presented. Based on a certain projectile, the optical path difference is analyzed. The optical path differences under different projectile velocities are figured out with the MATLAB software, which will provide a feasible error analysis method for the exploration of laser fuse or laser weapon under supersonic velocity.

This paper describes the development of a test rig dedicated to tests of artillery-type models and rotating missiles, in order to simulate the roll, yaw and pitch movements of the models. The factors affecting the design of the system are outlined. A special measurement technique is introduced in order to determine at the same time the rotations around the three main axes of the model. Furthermore, a theoretical description based on a six-degree-of- freedom approach is developed for the motion of the model. Part of the investigation concentrates on the characterization of the roll, yaw and pitch damping behavior of the models in use for the calculation the dynamic behavior of the models used during the spin-down phase of flight. Experiments from the first test campaigns are presented. The initial experiments using the test rig have been encouraging and have demonstrated its ability to study the dynamics of configurations at high spin rates.

The interaction of a concentrated vortex and a shock wave may occur in many cases in the operational environment of supersonic aircraft and missiles. The interaction may be the result of vortices created by the forward components of a supersonic vehicle (for example canards, wings, etc) convecting downstream and interacting, for example, with shock waves present in front of air intakes or other wings. The objective of the present work is to conduct an experimental studies simulating the interaction of streamwise wing-tip vortices with shock waves in supersonic flows. Flow-field visualizations show the different behavior (for example: unsteady vortex breakdown). To elucidate the complex flow containing shock waves, backflow and large-scale fluctuations, measurements using a laser Doppler velocimeter (LDV) were carried out.

The present paper describes the numerical methods to compute flows around the NAL Scaled Supersonic Airplane in ascending flight and its separation process from the rocket booster. The Euler equations are solved on unstructured tetrahedral grid. A seamless procedure from CAD data to grid generation is developed to treat the complex configuration of the NAL airplane piggybacked on a rocket for ascending flight. The separation process of the airplane from the rocket is efficiently simulated by the overset unstructured grid method. Comparisons of the computed results with the experiments show good agreements in the lift and pitching moment histories. It is also shown that the small parts attached between the rocket and the airplane significantly change the aerodynamic coefficients of the airplane at transonic regime.

Physical Review Letters

Stability analysis on a supersonic natural laminar flow (NLF) wing designed by JAXA was performed at both flight and wind tunnel test conditions as a joint research program between ONERA and JAXA. Both parties validated the NLF wing effect by confirming suppression of crossflow instability at design point, using an e(N) method with fixed β strategy proposed by ONERA. Although JAXA’s flight test vehicle and wind tunnel test model had almost same averaged roughness height of about 1micrometer, it was shown that surface roughness had little influence on measured transition location at both flight and wind tunnel test conditions. Furthermore, JAXA found that the pressure distribution measured in flight test did not completely coincide to the target pressure distribution (Cp (Target)) for the NLF wing design. It revealed that measured transition location at outer wing region was not delayed in flight test, compared with the transition location predicted with the Cp (Target) which had large gain to delay the transition. Influence of Reynolds number on transition location was also studied for the target pressure distribution Cp (Target). Consequently, chordwise transition location, (x/c)Tr., rapidly moves from mid-chord location to forward location (near leading edge) for increasing Reynolds numbers, according to the change of instability of boundary layer from Tollmien-Schlichting instability to crossflow instability., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0062081000, レポート番号: JAXA-RR-12-009E

In this article, the interaction of a normal shock with a yawed wedge moving at supersonic speed has been considered. The vorticity distribution of a particle over the diffracted shock wave for various combinations of yawed angles, Mach number of the shock wave and Mach number of the moving wedge have been obtained. Further triple point angle χ in Mach reflection has been calculated for the various parameters.

The supersonic flow around a reentry capsule with and without the jet of a thruster has been studied using a Navier-Stokes method. While previous studies addressed mainly the basic problem, in this paper the jet / airflow interaction has been investigated on a real capsule configuration. Results obtained demonstrate that the flow field and the surface pressure field are changed significantly by the jet. As expected from earlier basic studies, the force applied on the vehicle by the thruster is increased significantly by these effects.

High incidence flows about a Space-Plane under investigation at the National Aerospace Laboratory of Japan were numerically simulated at supersonic speed.The basic equations used here are the Reynolds-averaged thin-layer Navier-Stokes equations. The convective terms are evaluated by Roe’s flux difference splitting while the viscous terms are centrally differenced and treated separately from the convective terms.The resulting finite difference equations are integrated in time employing the LU-ADI factorization algorithm. The eddy-viscosity necessary for turbulent flow computation is evaluated by a modified Baldwin-Lomax model. The Space-Plane model consists of fuselage, strake-wing, and tail with angle of inclination of 30 deg. The computational grid around the model has an O-C topology with 110 streamwise, 103 circumferential, and 49 normal points. Computation was carried out at a Mach number of 1.5, angles of attack ranging from 0 to 45 deg, and a Reynolds number of 3 ×106. Strong shock waves as well as leading-edge vortices in the complicated flowfields are captured in detail by the present numerical simulation. Some of the computational results including force coefficients and surface pressure distributions were also compared with experimental data. These comparisons show that the numerical predictions are in good agreement with NAL’s supersonic wind tunnel test data.