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In the present study, a 3-D study of the aerodynamic coefficients of an ejection seat system is performed. The analysis is performed using the Reynolds Averaged Navier-Stokes equations where an unstructured grid of the polyhedral cells is used. The aerodynamic coefficients are calculated using a density-based solver and the standard k-ε turbulence model. The investigation is performed at Mach number (Ma) = 1.25 by varying the angle of attack (α) from -15° to 15° with the increment of 5° where the yaw angle (β) is fixed at 0°. According to the findings, the magnitude of the axial force coefficient (CX) increases as α decreases whereas the value of the normal force coefficient (CZ) decreases with the increase of α. Similarly, the value of the pitching moment coefficient (Cm) decreases with the increase of α.

AbstractThis research endeavor seeks to analyze the nose shapes of a wedge and a cone through simulations conducted with the ANSYS-FLUENT software. When designing missiles and aerodynamic vehicles for high speed, the selection of nose shape is thought to be the most crucial. The aerodynamic bodies traveling at supersonic speeds experience a pronounced influence of fluid compressibility, wherein high-speed fluid particles impart their kinetic energy as heat upon interacting with the frontal surface. The primary difficulty is that missiles and aerodynamic vehicles with high aerodynamic loads depend heavily on the shape of their noses. Therefore, the purpose of this research work is to compare the shape of the wedge and the conical nose at Mach number 2.0. A numerical simulation is used to estimate aerodynamic properties such as downstream Mach number, shock wave angle, pressure ratio, density ratio, temperature ratio, and aerodynamic coefficient, and the results are compared with theoretical results, which are obtained to solve oblique shock relation and Taylor–Maccoll equation. The conical-shaped nose model exhibits superior aerodynamic characteristics compared to the wedged-shaped nose. [ABSTRACT FROM AUTHOR]

Shock wave formation in aircraft drastically retards the maneuverability of aircraft by increasing drag, destabilizing buffet, and diminishing engine efficiency. Conventional method of shock wave strength reduction by structural modifications are cumbersome and lack scalability and the repurposability is difficult to incorporate. The aim of this study was to investigate the effectiveness of nanomaterial coatings in reducing shock wave strength around airfoil surfaces. A nanocoating of graphene was applied to the surface of a TsAGI S-12 airfoil and the aerodynamic characteristics were evaluated. The flow visualization captured from the wind tunnel experiments at various Mach reveals 18% reduction in shock wave angle. The reduced shock wave angle and pressure coefficient confirmed that flow stability increased with increase in Mach, emphasizing that the technique is highly effective for supersonic speeds. Furthermore, to investigate the role of surface roughness, AFM measurements were performed, revealing that the surface roughness decreased from 10 to 2 nm. Notably, this decrease in surface roughness has a direct and vital influence on the reduction of shock wave strength. Specifically, the smoother surface offered by the nanocoating minimizes the irregular reflections that occur on rough surfaces, thereby dissipating the shock wave more effectively. The observed reduction in shock wave angle, surface drag, and static pressure has transpired as a consequence of the reduced surface roughness upon nanomaterial coating. Nanomaterial surface coating is a simple and highly effective method for shock wave strength reduction with scalability that can be readily employed in the aircraft industry. [ABSTRACT FROM AUTHOR]

Obvious aeroelastic deformation occurs in spinning projectiles with large slenderness ratio, which seriously affects flight stability and maneuverability. This paper investigates the aeroelastic response of spinning projectiles with large slenderness ratio under supersonic speed. Based on a dynamic mesh method, an unsteady numerical simulation method is developed to study the aeroelasticity of spinning projectiles by coupling aerodynamics and structural dynamics. The numerical simulation method is well validated by the experimental results of AGARD 445.6 wing flutter. Then, the aeroelastic response of spinning projectiles with large slenderness ratio is numerically explored under different flight conditions. The aeroelastic response is obtained, revealing the presence of beat vibrations and variations in response frequency. Furthermore, the influence mechanism of flight conditions on the aeroelastic response is analyzed. The results suggest that the coupling of the first two modes of the projectile caused by the spinning motion leads to the occurrence of beat vibrations in the aeroelastic response; the coupling degree of the first two modes decreases as the angle of attack increases and it increases with the increase in spinning speed; and the time−averaged deformation caused by the time−averaged aerodynamic force is beneficial to the convergence of the aeroelastic response of spinning projectiles, while the rotation−induced Magnus effect is counterproductive.

Copyright of Journal of China University of Petroleum is the property of China University of Petroleum and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Research on supersonic transport seeks to overcome two main challenges, to mitigate sonic boom and improve aero- dynamic performance, in order to become environmentally viable and economically practicable. To achieve these goals, this research proposes a waverider configuration to solve the challenge of designing commercial supersonic aircraft. The waverider is famous for providing a higher lift-to-drag ratio than that provided by other conventional designs for super- sonic and hypersonic flights. The study of three types of supersonic waveriders, cone-derived, power-law, and osculating-cone waverider, is introduced for low-supersonic designs. The flight condition includes a speed of Mach 1.5 at an altitude of 18,000 m. An unstructured grid with an inviscid Euler equation solver is used to calculate the flow field of the vehicle. The ground noise post-processing uses an augmented Burgers equation solver to assist in far-field signature evaluation. The result of a cone-derived waverider provided significant volume with low overpressure, while the power-law waverider showed the advantages of aerodynamic performance and ample design space. The osculating-cone waverider obtained the merits from both the cone-derived and power-law waveriders. The waverider is expected to have a quality that satisfies the design objectives of supersonic aircraft: high aerodynamic performance and sonic boom mitigation. [ABSTRACT FROM AUTHOR]

area of interest for flow control due to coherent vortical structures existing on the leeward side of wing. In the present research, a numerical study was carried out by using two different turbulence models at Mach No. 1.6, to identify the influence of turbulence modeling and wing on overall pressure distribution and onset of flow separation. Fully structured multi-block grid was used with the grid refinement near the wall to resolve the vortical flow structures accurately. Different turbulence models used were SA (Spalart-Allmaras) and kω-SST (Shear-stress Transport). An under prediction in pressure distribution was observed by using SA Turbulence model, in comparison with known experimental data at fixed Mach No. The analysis of results presented showed that the height of primary vortex increases by increasing the angle of attack and by keeping the Mach No. as fixed. Moreover, the vortex formation on cambered wing seems less pronounced as compared with that of sharp wing at high angle of attack.

A Schiff base derived from 3,4-dimethoxybenzaldehyde and p-aminobenzoic acid (SBDA) was synthesized by novel supersonic speed gas impacting method. The morphology and geometric structure of the synthesized Schiff base SBDA were investigated, the results showed that the particles of SBDA were not completely regular and the HOMO and LUMO of SBDA optimized geometry structure using the semiempirical method PM3 were −9.115 eV and −1.191 eV, respectively, and some bond length of synthesized SBDA became short and bone angle became large compared to the reactant. The thermogravimetric analysis results indicated that SBDA began to decompose above 280°C; the decomposition temperature and thermogravimetric rate could vary with the different heating rate.

Obvious aeroelastic deformation occurs in spinning projectiles with large slenderness ratio, which seriously affects flight stability and maneuverability. This paper investigates the aeroelastic response of spinning projectiles with large slenderness ratio under supersonic speed. Based on a dynamic mesh method, an unsteady numerical simulation method is developed to study the aeroelasticity of spinning projectiles by coupling aerodynamics and structural dynamics. The numerical simulation method is well validated by the experimental results of AGARD 445.6 wing flutter. Then, the aeroelastic response of spinning projectiles with large slenderness ratio is numerically explored under different flight conditions. The aeroelastic response is obtained, revealing the presence of beat vibrations and variations in response frequency. Furthermore, the influence mechanism of flight conditions on the aeroelastic response is analyzed. The results suggest that the coupling of the first two modes of the projectile caused by the spinning motion leads to the occurrence of beat vibrations in the aeroelastic response; the coupling degree of the first two modes decreases as the angle of attack increases and it increases with the increase in spinning speed; and the time−averaged deformation caused by the time−averaged aerodynamic force is beneficial to the convergence of the aeroelastic response of spinning projectiles, while the rotation−induced Magnus effect is counterproductive. [ABSTRACT FROM AUTHOR]

Summary: Supersonic combustion ramjet engine is more fascinating among all the air‐breathing engines. Due to its higher thrust to weight ratio, researchers are more interested to get the superior combustion performance at the optimum boundary conditions. The flow field characteristics and combustion performance have been analysed with the help of Ansys 14.0 software. Generic scramjet combustor of German Aerospace Center (DLR) has been taken into consideration for comparison purpose and off design analysis has been conducted to investigate and analyse the changes. Two dimensional compressible Reynolds Averaged Navier‐Stokes (RANS) turbulence model has been opted with the finite‐rate/eddy‐dissipation reaction model. K‐ε two equation turbulence model has been selected to reach up to reasonable accuracy. Validation of the present work has been done with the help of both non‐reacting and reacting type data from open literature. To choose the appropriate meshing of the computational model three different types of mesh elements, that is, coarse, medium and fine has been analysed and also grid independence analysis is performed. The present article objective is to get optimum boundary condition by changing the incoming air temperature and pressure at constant Mach number to connect the bridge between incoming air temperature and pressure to the change in velocity throughout the combustion chamber. The detailed understanding and explanation have been done by varying the temperature range of incoming air because of its major impact on combustion performance. Nonetheless, a small variation of air pressure will also discuss to observe the parameters which majorly influence while doing performance analysis. At the end the Optimum boundary condition for the present computation work is observed to be at 833 K temperature with 115 299 Pa pressure. [ABSTRACT FROM AUTHOR]

Experiments were carried out to bring out the effects of the single vertical tail (SVT) and the twin vertical tail (TVT) of a typical winged reentry vehicle on aerodynamic coefficients at supersonic speed (M=2.95) for the range of angles of attack α varying from -5 to 25 deg. Toward this, the tunnel was calibrated for the Mach number (flow uniformity) and the flow angularity. As a second level of calibration, experiments were carried out on a standard Experiment Re-Entry Space System reentry model. The results indicated that the obtained aerodynamic coefficients were well within 3% of the literature values. Flowfield studies through schlieren flow visualization illustrated almost similar flow patterns for both SVT and TVT configuration for the tested α. However, flow separation on the leeward side at high α was observed to be strong in the case of the TVT configuration as compared to the SVT configuration. The aerodynamic coefficients such as CA, CN, and L/D for the TVT configuration were 9, 11, and 13% higher than those of the SVT configuration. On the other hand, the SVT configuration was observed to be more stable, whereas mixed stability was seen in the TVT configuration. As far as the stability margin was concerned, the SVT configuration showed the stability margin as high as 33%, whereas the TVT configuration exhibited only a 14% stability margin. Beyond α=10  deg, the TVT configuration was found to be neutrally stable. [ABSTRACT FROM AUTHOR]

Existing theories are reviewed and compared. From the Tangent Wedge formula a simple expression for the normal force on thin wings is derived and compared with some existing experimental data, which cover the whole supersonic-hypersonic range of Mach numbers. [ABSTRACT FROM AUTHOR]

The response and control of the transverse vibration of an axially moving free-free beam with high speed subjected to the thermal shock is studied. Considering the thermal shock and the control force, the transverse vibration equation and boundary conditions are formulated by the Hamilton principle. The instant linearized equations are set up based on Galerkin' s method for the approximation solution. Considering the thermal shock, the displacement responses are calculated, where the axially speed is 2 Ma. Then the controller is designed based on LQR method. The optimal Q and R matrices are selected by the weighting coefficient method. The displacement responses for the control method are simulated by a numerical example. The results show that the control method can suppress the transverse vibration effectively [ABSTRACT FROM AUTHOR]

CFD (Computational Fluid Dynamics) Analysis of Delta wing configurations used for supersonic flight regime is a key area of interest for flow control due to coherent vortical structures existing on the leeward side of wing. In the present research, a numerical study was carried out by using two different turbulence models at Mach No. 1.6, to identify the influence of turbulence modeling and wing on overall pressure distribution and onset of flow separation. Fully structured multi-block grid was used with the grid refinement near the wall to resolve the vortical flow structures accurately. Different turbulence models used were SA (Spalart-Allmaras) and kω-SST (Shear-stress Transport). An under prediction in pressure distribution was observed by using SA Turbulence model, in comparison with known experimental data at fixed Mach No. The analysis of results presented showed that the height of primary vortex increases by increasing the angle of attack and by keeping the Mach No. as fixed. Moreover, the vortex formation on cambered wing seems less pronounced as compared with that of sharp wing at high angle of attack. [ABSTRACT FROM AUTHOR]

The first supersonic business jet to enter the market will not fare its competition from rival supersonic designs under development. Its true competitors are the then current generation of highly evolved high-subsonic business jets when compared an economic grounds. For a price tag of $1m for the now conception of very light jets, ranging up to $45m for the highest-performing 'race-horse'-like corporate jets. this breed of aircraft is able to accommodate the needs of most executives, VIPs, officials, from corporate transportation to cargo services of civil to military origin. Understanding the state of modern business class aircraft and their market is essential in gaining base knowledge required foe any supersonic business jet endeavor aiming at a prospective market. The key descriptors for this marketplace are market potential, market productivity, and market drivers, altogether being a measure fur growth and consumer demand. Such common denominator is used to gain the understanding necessary to ascertain and visualise the top level implications regarding any supersonic business case. Having assembled an understanding of the key descriptors for business aviation, the study first analyses the flight operation of traditional subsonic and high-subsonic business jets. Such perceptive is then complemented with the peculiarities associated with supersonic operation, ultimately defining the supersonic solution space consisting of market viability, efficiency, and its overall flight performance. Consequently a vehicle development strategy and mission specification are suggested for the first generation of supersonic business jets (SSBJ) and supersonic cargo jots (SSCJ). [ABSTRACT FROM AUTHOR]

The first generation of supersonic commercial transportation has seen three serious attempts to arrive at an economically and environmentally viable aircraft. The US B2707-2001300 design was cancelled early before a prototype could emerge; the Russian Tu-144 design succeeded to become the first supersonic transport butspanned only a few years of restricted airline service; the Anglo- French Concorde endured more than 27 glamorous airline service years until the last of its species was retired on 30 August 2003. This first generation was followed by a second generation of supersonic commercial transport projects in the time period between 1986 until about 1999, designs which did not proceed towards the production hardware stage. This study critically examines the anatomy of two generations of supersonic commercial transport design failures and successes in order to arrive at lessons learned free of 'wishful thinking'. The design conditions leading to the identification of the product 'solution space' for an economically and environmentally acceptable supersonic commercial transport are discussed. Having assembled an understanding of the product metrics valid for supersonic commercial transports, the paper then provides an outlook for the first generation of supersonic corporate and cargo jet projects. This first generation of supersonic business jet (SSBJ) and supersonic cargo jet (SSCJ) projects spans a period of nearly two decades of development, starting from 1988 until today. The present study identifies that the product development metrics of this class of aircraft is radically different compared to the metrics alid for supersoniccommercial transports. The challenges in VIP transportation and dedicated freight transportation at supersonic speeds are portrayed leading to two principal trains of thought targeting the development of the first supersonic business jet and/or supersonic cargo jet hardware: the development based on a new airframe, and alternatively the development based on an existing airframe. [ABSTRACT FROM AUTHOR]

This study presents the Doppler shift behaviour on a supersonic wireless channel. Firstly, the authors present the instrumented projectile and the flight parameters influencing the Doppler shift of the RF carrier. In a second part, they start by analysing experimental results from a non-dedicated measurement campaign. Doppler shift is extracted from the transmitted minimum shift keying digital data thanks to a classical phase-tracking algorithm. To explain the particular shape of the Doppler curve obtained during the flight, they carry on with the study of multipath components in the propagation environment by performing a ray-tracing simulation to identify them and evaluate their contribution. The study is completed by the analysis of the two remaining flight dynamics effects, namely spinning and nutation, on the Doppler spread. They finally draw some conclusions and present the remaining steps required to obtain a realistic supersonic speed ground-projectile channel model. [ABSTRACT FROM AUTHOR]

In this paper we develop a dynamical model of the propagating nonlinear localized excitations, supersonic kinks, in the cation layer in a silicate mica crystal. We start from purely electrostatic Coulomb interaction and add the Ziegler-Biersack-Littmark short-range repulsive potential and the periodic potential produced by other atoms of the lattice. The proposed approach allows the construction of supersonic kinks which can propagate in the lattice within a large range of energies and velocities. Due to the presence of the short-range repulsive component in the potential, the interparticle distances in the lattice kinks with high energy are limited by physically reasonable values. The introduction of the periodic lattice potential results in the important feature that the kinks propagate with the single velocity and single energy, which are independent on the excitation conditions. The unique average velocity of the supersonic kinks on the periodic substrate potential we relate with the kink amplitude of the relative particle displacements, which is determined by the interatomic distance corresponding to the minimum of the total, interparticle plus substrate, lattice potential. The found kinks are ultradiscrete and can be described with the "magic wave number" q=2π/3a, which was previously revealed in the nonlinear sinusoidal waves and supersonic kinks in the Fermi-Pasta-Ulam lattice. The extreme discreteness of the observed supersonic kinks, with basically two particles moving at the same time, allows the detailed interpretation of their double-kink structure, which is not possible for the multikinks without an account for the lattice discreteness. Analytical calculations of the displacement patterns and energies of the supersonic kinks are confirmed by numerical simulations. The computed energy of the found supersonic kinks in the considered realistic lattice potential is in a good agreement with the experimental evidence for the transport of localized energetic excitations in silicate mica crystals between the points of ^{40}K recoil and subsequent sputtering.

Support interferences of low-aspect-ratio flying-wing at M=0.6, 0.9 1.5 are calculated by means of numerical simulation methods, both near-field interference and far-field interference are concerned, and attempt is made to analyse mean flow distortion distribution for the model. Results show that, drag coefficient and pitching moment coefficient of far-field interference account about 30% of the total support interference while lift coefficient account about 20% at M=0.6. For Mach number 0.9, angle of attack 2°, drag coefficient of far-field interference account 40%, and far-field interference make the vortex breakdown location move ahead at the angle of attack 18°. Far-field interference can be ignored at M=1.5. It's hard to get the distortion of Mach number and angle of attack (Δ M and Δ α) at subsonic speed (M=0.6); when upstream become M=0.9, support interference affects the shock position, and the surface pressure coefficient differences are sensitive with Mach number, support interference decrease the Mach number for about 0.02 and decrease the angle of attack for 0.1° on the model surface. [ABSTRACT FROM AUTHOR]

Copyright of Journal of Ballistics / Dandao Xuebao is the property of Journal of Ballistics Editorial Department and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

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.