Abstract: JF-10 detonation-driven high-enthalpy shock tunnel was re-built with a forward detonation cavity (FDC) driver and the experimental data from its performance tests are summarized and reported in this paper. Test-duration of high-enthalpy flows produced with the improved JF10 is found to be extended by two times under the condition that the FDC driver is about 40% shorter than the original one. The uniform pressure area of the thus-obtained hypersonic flows in a 500 mm diameter conical nozzle is about 700 mm in length and 400 mm in diameter. Incident shock wave decay in the driven section appears to be much less by comparing with the original JF-10 shock tunnel. The performance improvement of JF-10 high-enthalpy shock tunnel was demonstrated to be very successful and high quality hypervelocity flows can be generated for aero-thermochemistry experiments.
Content Type Journal Article Category Research Article Pages 29-36
DOI 10.1260/1759-3107.2.1.29
Authors
Z. Jiang, State Key Laboratory of High Temperature Gasdynamics Institute of Mechanics, Chinese Academy of Sciences, Beijing, 100190, China J. Lin, State Key Laboratory of High Temperature Gasdynamics Institute of Mechanics, Chinese Academy of Sciences, Beijing, 100190, China W. Zhao, State Key Laboratory of High Temperature Gasdynamics Institute of Mechanics, Chinese Academy of Sciences, Beijing, 100190, China
Journal International Journal of Hypersonics
Abstract: Effect of injection angle in mixing and combustion in a scramjet combustor is numerically simulated. Three dimensional Navier Stokes equations alongwith k-ε turbulence model are solved using commercial CFD software. Both infinitely fast rate kinetics and single step finite rate kinetics are used to model chemical kinetics. Turbulence chemistry interaction is modeled by Eddy Dissipation Concept (EDC). Good agreement between the computed and experimental results for angular injection (30°) and perpendicular injection forms the basis of further analysis. More flow blockage has caused significant upstream interaction for perpendicular injection and terminal shock is seen to anchor upstream of combustor step; while for angular injection, the flow field is predominantly supersonic. Single step finite rate chemistry show comparatively low pressure and lesser upstream interaction because of the presence of backward reaction. Thermochemical variables are analysed to study the effect of angle of injection on nature of combustion (whether premixed or diffusive), heat release pattern, combustion efficiency etc.
Content Type Journal Article Category Research Article Pages 15-28
DOI 10.1260/1759-3107.2.1.15
Authors
M S R Chandra Murty, Directorate of Computational Dynamics, Defence Research and Development Laboratory Hyderabad, 500058, India Debasis Chakraborty, Directorate of Computational Dynamics, Defence Research and Development Laboratory Hyderabad, 500058, India
Journal International Journal of Hypersonics
Abstract: In this paper, the effectiveness of counterflowing jets as heat-reduction devices for large-angle blunt cones flying at hypersonic Mach numbers is numerically simulated with various coolant jets. Different jet conditions have been chosen to investigate the effect of the counterflow jet on the surrounding flow field of nose cone. The compressible, unsteady, axisymmetric Navier-Stokes equations are solved with SST turbulence model for free stream Mach number of 5.75 at 0° angle of attack with and without gas injection. The coolant gas (air, Carbon Dioxide, and helium) is chosen to inject into the hypersonic flow at the nose of the model. The numerical results presented the surface heat reduction for different coolant jets. According to the investigation of various conditions of opposing jets, important phenomena of flow field and some effective jet conditions are found.
Content Type Journal Article Category Research Article Pages 1-14
DOI 10.1260/1759-3107.2.1.1
Authors
M. Barzegar Gerdroodbary, Department of Mechanical Engineering, Iran University of Science & Technology, Narmak, Tehran 16844, Iran M. A. Fayazbakhsh, School of Engineering Science, Simon Fraser University, Surrey, BC, Canada
Journal International Journal of Hypersonics
Abstract: The Euler equations are solved on unstructured triangular meshes for hypersonic flow over double-wedge geometries. The driving algorithm is an upwind biased cell centered finite volume method. AUSM+ method is used to split the fluxes. Edney (1968) studied the shock interactions by impinging an externally generated planar oblique shock on the bow shock generated by a cylinder. Depending upon the parametric conditions Edney classified the shock interactions in different types. Two interaction topologies, namely Type-VI and Type-V and the transition from Type-VI to Type-V are studied in details. Both six-shock and seven-shock configurations of Type-V interaction are presented.
Content Type Journal Article Category Research Article Pages 225-244
DOI 10.1260/1759-3107.1.4.225
Authors
Pabitra Halder, SM Laboratory, CSIR-CMERI, Durgapur, India Kalyan P. Sinhamahapatra, Department of Aerospace Engineering, Indian Institute of Technology, Kharagpur, India Navtej Singh, Department of Aerospace Engineering, Indian Institute of Technology, Kharagpur, India
Journal International Journal of Hypersonics
Print ISSN 1759-3107
Journal Volume Volume 1
Journal Issue Volume 1, Number 4 / December 2010 PubDate: Fri, 03 Feb 2012 22:00:01 GMT
Abstract: This paper presents the results of numerical and experimental investigations on sonic jet (nitrogen) injections into a hypersonic cross-flow of air. The simulations aim at finding suitable injection conditions consistent with the experimental facility at IISc from the viewpoint of combustion. Numerical results under predicted the measured jet penetration depth for normal injection. It is found from numerical simulations that the total pressure loss in case of oblique injection is less than the normal injection.
Content Type Journal Article Category Research Article Pages 245-262
DOI 10.1260/1759-3107.1.4.245
Authors
Ratan Joarder, Department of Aerospace Engineering, Indian Institute of Science, Bangalore, India G. Jagadeesh, Department of Aerospace Engineering, Indian Institute of Science, Bangalore, India
Journal International Journal of Hypersonics
Print ISSN 1759-3107
Journal Volume Volume 1
Journal Issue Volume 1, Number 4 / December 2010 PubDate: Fri, 03 Feb 2012 22:00:01 GMT
Abstract: Amodel for plasma assisted combustion of ethylene-air mixtures at conditions typical for scramjet combustion chamber is developed combining classical mechanisms of thermal combustion with non-thermal plasma chemistry. Numerical simulations showed that sufficiently strong reduction of ignition induction time at a reasonable energy cost can be realized with help of filamentary discharges. Starting from the discharge region, the gas mixture is heated due to exothermic reactions involving atomic oxygen and secondary chemical radicals. Temperature increment to the end of this stage for ethylene-air mixture is relatively small. An important effect of this stage is not heating but production of transient species. Then, a period with slow growth of temperature follows, which terminates by fast combustion. Processes causing the first fast growth of gas temperature are analyzed, and intermediate species controlling acceleration of ignition are determined numerically for plasma assisted combustion of stoichiometric mixture of ethylene with air. The value of the calculated induction time defined as a moment of the fast combustion is rather sensitive to the particular combustion mechanism adopted. This manifests a necessity to refine combustion mechanisms for conditions typical for scramjet combustion chamber with plasma initiation - one atmosphere pressure, static gas temperature around 700 K and appearance of atomic oxygen&dR.
Content Type Journal Article Category Research Article Pages 209-224
DOI 10.1260/1759-3107.1.4.209
Authors
Maxim A. Deminsky, State Research Center RF "Kurchatov Institute", Moscow, Russia Igor V. Kochetov, State Research Center RF Troitsk Institute for Innovation and Thermonuclear Research (TRINITI), Troitsk, Moscow region, Russia Anatoly P. Napartovich, State Research Center RF Troitsk Institute for Innovation and Thermonuclear Research (TRINITI), Troitsk, Moscow region, Russia Sergey B. Leonov, Joint Institute for High Temperature RAS, Moscow, Russia
Journal International Journal of Hypersonics
Print ISSN 1759-3107
Journal Volume Volume 1
Journal Issue Volume 1, Number 4 / December 2010 PubDate: Fri, 03 Feb 2012 22:00:01 GMT
Abstract: Experiments were performed in the T4 shock tunnel to investigate the self ignition of hydrogen in a supersonic air stream. Hydrogen was injected into the flow over an inclined flat plate for oncoming Mach numbers of 7.9 to 8.0. The nozzle-supply enthalpy was kept between 3.1 and 3.4 MJ/kg and two different pressure levels were used in the tests. Measurements of surface pressures were used to infer the location of ignition but only small pressure increases were obtained when combustion occurred. Therefore multiple tests at nominally the same condition were used so that statistical methods could be used to identify the ignition lengths. The ignition lengths of the hydrogen air mixture directly behind a strong leading edge shock indicate that Pergament's method is able to predict the ignition length to within 35% for the observed autoignition over the range of conditions tested.
Content Type Journal Article Category Research Article Pages 199-208
DOI 10.1260/1759-3107.1.4.199
Authors
Rainer M. Kirchhartz, Centre for Hypersonics, School of Mechanical and Mining Engineering, The University of Queensland, Brisbane, 4072, Australia Allan Paull, Centre for Hypersonics, School of Mechanical and Mining Engineering, The University of Queensland, Brisbane, 4072, Australia David J. Mee, Centre for Hypersonics, School of Mechanical and Mining Engineering, The University of Queensland, Brisbane, 4072, Australia Herbert Olivier, RWTH Aachen University, Aachen, 52060, Germany
Journal International Journal of Hypersonics
Print ISSN 1759-3107
Journal Volume Volume 1
Journal Issue Volume 1, Number 4 / December 2010 PubDate: Fri, 03 Feb 2012 22:00:00 GMT
Abstract: Tests were performed in a non-axisymmetric, single nozzle, strut-based ejector to investigate mass flow entrainment, choking mechanisms and stream mixing as a function of primary (strut nozzle) to secondary (duct inlet) flow stagnation pressure ratio. Experimental results show a mass flow choke in the mixing duct rather than a traditional aerodynamic choke in the strut gap. The stream mixing length was constant for lower primary flow pressures, whereas mixing length varied with pressure at higher values. A companion numerical study was performed using Reynolds Averaged Navier-Stokes solutions to investigate several turbulence models. Based on both 2-D and 3-D simulation results, compressibility correction to conventional incompressible twoequation models was required for capturing the supersonic ejector mixing phenomena. The Baldwin-Lomax and the SST two-equation models were capable of capturing the essential flow features. Even with compressibility correction, the k-Σ model could not reproduce wall-dominated phenomena such as mixing duct pressure recovery.
Content Type Journal Article Category Research Article Pages 181-198
DOI 10.1260/1759-3107.1.3.181
Authors
M. S. Balasubramanyam, Propulsion Research Center, University of Alabama in Huntsville, Huntsville, AL 35899 D. Lineberry, Propulsion Research Center, University of Alabama in Huntsville, Huntsville, AL 35899 C. P. Chen, Department of Chemical & Materials Engineering, University of Alabama in Huntsville, Huntsville, AL 35899 D. B. Landrum, Department of Mechanical and Aerospace Engineering, University of Alabama in Huntsville, Huntsville, AL 35899
Journal International Journal of Hypersonics
Abstract: An experimental activity is currently underway to develop a tunable diode laser absorption tomography (TDLAT) technique to measure flow properties in a supersonic combustion wind tunnel. The present study simulates the reconstruction of spectroscopic measurements to determine the effects of data collection geometry and ambient water vapor on reconstruction accuracy. The study also proposes a way to remove the effects of ambient water vapor absorption. Overall, results show that TDLAT data collection time can be significantly reduced while maintaining reconstruction accuracy by taking fewer projections with a high density of rays and by correcting for ambient water vapor absorption.
Content Type Journal Article Category Research Article Pages 173-180
DOI 10.1260/1759-3107.1.3.173
Authors
Elizabeth F. Martin, Department of Mechanical and Aerospace Engineering, University of Virginia, Charlottesville, Virginia, 22904, USA Christopher P. Goyne, Department of Mechanical and Aerospace Engineering, University of Virginia, Charlottesville, Virginia, 22904, USA Glenn S. Diskin, Chemistry and Dynamics Branch, NASA Langley Research Center
Journal International Journal of Hypersonics
Abstract: In this paper, the mechanism of detonation to quasi-detonation transition was theoretically discussed, a new physical model to simulate quasi-detonation was proposed, and one-dimensional numerical simulation was conducted. This study first demonstrates that the quasi-detonation is of thermal choking. If the conditions of thermal choking are satisfied by the chemical release of energy and some disturbances, the supersonic flow is then unable to accept the additional thermal energy, and the CJ detonation becomes the unstable quasi-detonation precipitately. The kinetic energy loss consumed by this transition process is first considered in this new physical model. The numerical results are in good agreement with previous experimental observations qualitatively, which demonstrates that the quasi-detonation model is physically correct and the study is fundamentally important for detonation and supersonic combustion research.
Content Type Journal Article Category Research Article Pages 145-156
DOI 10.1260/1759-3107.1.3.145
Authors
Yunfeng Liu, Laboratory of High Temperature Gasdynamics, Chinese Academy of Sciences Institute of Mechanics, CAS, Beijing 100190, China Zonglin Jiang, Laboratory of High Temperature Gasdynamics, Chinese Academy of Sciences Institute of Mechanics, CAS, Beijing 100190, China
Journal International Journal of Hypersonics
Abstract: Valid Physical Processes from Numerical Discontinuities in Computational Fluid Dynamics
Content Type Journal Article Category Research Article Pages 157-172
DOI 10.1260/1759-3107.1.3.157
Authors
Kun Xu, Department of Mathematics, The Hong Kong University of Science and Technology, Kowloon, Hong Kong Quanhua Sun, LHD, Institute of Mechanics, Chinese Academy of Sciences, Beijing, 100190, China Pubing Yu, Department of Mathematics, The Hong Kong University of Science and Technology, Kowloon, Hong Kong
Journal International Journal of Hypersonics
Abstract: Hypersonic laminar flow past a rearward facing step has been numerically investigated using computational fluid dynamics (CFD). The flow parameters were : total specific enthalpy ho ≈7.6 MJ/kg; unit Reynolds number Re ≈1.82 × 106 1/m ; and Mach number M∞ ≈7.6. A detailed grid independent study has been carried out to investigate the sensitivity of the surface heat flux in the regions of separation and reattachment. The nature of the flow in the close vicinity of the step is particularly emphasised. The influence of real-gas effects such as the thermal and chemical non-equilibrium are studied using Park's two-temperature model and finite-rate chemistry models respectively. The numerical results are then compared with the available experimental data of surface heat flux measurements.
Content Type Journal Article Category Research Article Pages 115-134
DOI 10.1260/1759-3107.1.2.115
Authors
Deepak N Ramanath, School of Engineering & IT, University of New South Wales, Australian Defence Force Academy, Canberra 2600, Australia Sudhir L Gai, School of Engineering & IT, University of New South Wales, Australian Defence Force Academy, Canberra 2600, Australia Andrew J Neely, School of Engineering & IT, University of New South Wales, Australian Defence Force Academy, Canberra 2600, Australia
Journal International Journal of Hypersonics
Print ISSN 1759-3107
Journal Volume Volume 1
Journal Issue Volume 1, Number 2 / June 2010 PubDate: Wed, 03 Nov 2010 15:16:24 GMT
Abstract: Reentry vehicle often uses reaction control system for providing necessary critical forces during reentry. The effect of lateral jet due to reaction control system on aerodynamic characteristics especially stability aspect of a reentry vehicle has been investigated .CFD analysis of the interaction of a lateral jet on Leeward side has been investigated on a reentry body with a circular sonic air jet injected normally in to hypersonic flow from the body surface. A systematic numerical analysis for a lateral jet interaction was performed at angles of attack 0, 2, 5, 8, and 10 degrees at free stream Mach number 8.1. The methodology has also be verified by comparing the pressure distribution at an angle of attack of 20 degrees with experimental data. The effect of lateral jet on aerodynamic characteristic has been studied by comparing the variation of aerodynamic characteristic along the length with and without lateral jet. The aerodynamic coefficients variation of aerodynamic characteristics along the axial length is studied with and without jet. Force coefficients, pitching moment coefficient and centre of pressure are studied with jet and without jet conditions by varying the angle of attack. The Mach contours, velocity vectors and path lines represent the effect of lateral jet. The computational results are validated here to compare the experimental results obtained from open literature.
Content Type Journal Article Category Research Article Pages 135-143
Abstract: The effects of aerospike geometry on the drag reduction and heat transfer rates for a large-angle blunt cone flying at hypersonic Mach numbers are investigated in a hypersonic shock tunnel. Two spike geometries are considered. The first is a plain spike with a conical tip and the second is a telescopic aerospike fitted with discs of decreasing diameter in the direction opposite to the flow direction. These aerospikes are fitted to a 120° apex-angle blunt cone and results are investigated at free stream Mach numbers of 5.75 and 7.9 for different angles of attack. The aerodynamic forces are measured using an accelerometer-based force balance system and the heat transfer rates are measured using platinum thin film sensors. It is found that the telescopic aerospike has better drag reduction performance at angles of attack beyond 2° while the performance of the plain aerospike is better for angles of attack closer to zero degrees.
Content Type Journal Article Category Research Article Pages 93-114
DOI 10.1260/1759-3107.1.2.93
Authors
S Srinath, Department of Aerospace Engineering, Indian Institute of Science, Bangalore-560012, India K. P. J Reddy, Department of Aerospace Engineering, Indian Institute of Science, Bangalore-560012, India
Journal International Journal of Hypersonics
Print ISSN 1759-3107
Journal Volume Volume 1
Journal Issue Volume 1, Number 2 / June 2010 PubDate: Wed, 03 Nov 2010 15:16:23 GMT
Abstract: The Short Duration Propulsion Test and Evaluation (SDPTE) Program and the Hy-V Program have recently been combined with the aim of examining the influence of ground test facilities on scramjet performance. The combined program includes both research and educational activities that are being conducted by a consortium of university, industry and government participants. The objectives of the combined program are to; 1) Resolve ground testing issues related to the effects of test medium on dual-mode scramjet engine performance, 2) Resolve ground testing issues related to the duration of the test flow on dual-mode scramjet engine performance, and 3) Educate and motivate a new generation of aerospace engineers through student participation and research. This paper provides an overview and status of the combined program but focuses on objectives 1) and 2). The ground testing issues associated with these objectives are being examined using a range of facilities. These include a continuous-flow direct connect facility, a freejet blowdown facility and an impulse facility. By testing in a range of facilities, the effects of combustion generated test medium vitiates and test flow duration on the operation of two dual-mode scramjet flowpaths will be examined. The experiments will focus on flowpath operation at conditions equivalent to a flight Mach number of 5. However, some Mach 7 freejet testing will also take place. The program will conclude with a Mach 5 flight experiment of both scramjet flowpaths aboard a sounding rocket such that differences between ground and flight performance data can also be isolated.
Content Type Journal Article Category Research Article Pages 77-92
DOI 10.1260/1759-3107.1.2.77
Authors
Christopher P. Goyne, University of Virginia, Charlottesville, VA, 22904-4248, USA Daniel Cresci, Alliant Techsystems Inc. (ATK GASL), Ronkonkoma, NY, 11779, USA Thomas P. Fetterhoff, AEDC, Arnold AFB, TN, 37389-9011, USA
Journal International Journal of Hypersonics
Print ISSN 1759-3107
Journal Volume Volume 1
Journal Issue Volume 1, Number 2 / June 2010 PubDate: Wed, 03 Nov 2010 15:16:23 GMT
Abstract: Editorial
Content Type Journal Article Category Editorial Pages i-ii
DOI 10.1260/1759-3107.1.1.i
Authors
K. P. J. Reddy, Department of Aerospace Engineering, Indian Institute of Science, Bangalore-560012, India
Journal International Journal of Hypersonics
Abstract: Drag reduction by counterflow supersonic jet for a 60-degree apex angle blunt cone in high enthalpy flow is investigated in a free piston driven hypersonic shock tunnel, HST3. For flow Mach number of 8 with specific flow enthalpy of 5 MJ/kg, it has been observed that the drag force decreases with increase in the ratio of supersonic jet total pressure to the freestream pitot pressure until the critical injection pressure ratio is reached. Maximum percentage drag reduction of 44 is measured at the critical injection pressure ratio of 22.36. Further increase in injection pressure ratio has reduced the percentage drag reduction. Experimentally obtained drag signals portray the change in nature of the flowfield, around the model, across the critical injection pressure ratio.
Content Type Journal Article Category Research Article Pages 69-76
DOI 10.1260/1759-3107.1.1.69
Authors
Vinayak Kulkarni, Department of Mechanical Engineering, Indian Institute of Technology Guwahati, INDIA K. P. J. Reddy, Department of Aerospace Engineering, Indian Institute of Science, Bangalore, INDIA
Journal International Journal of Hypersonics
Abstract: The net drag on a non-fuelled, internal compression, constant-area combustor scramjet engine was measured using a single-component accelerometer balance in a shock tunnel at a freestream Mach number of 8 and a flow total enthalpy of 1.35 MJ/kg. The flow fields of the model were simulated using a commercial CFD code in order to understand the aerodynamics of the engine and compare with the measured net drag co-efficient. Measured and computed values of the net drag co-efficient were found to be in a good agreement.
Content Type Journal Article Category Research Article Pages 59-68
DOI 10.1260/1759-3107.1.1.59
Authors
K. K. N. Anbuselvan, Department of Aerospace Engineering, Indian Institute of Technology Bombay, Powai, Mumbai 400076, INDIA V. Menezes, Department of Aerospace Engineering, Indian Institute of Technology Bombay, Powai, Mumbai 400076, INDIA K. S. N. Abhinav Kumar, Department of Aerospace Engineering, Indian Institute of Technology Bombay, Powai, Mumbai 400076, INDIA
Journal International Journal of Hypersonics
Abstract: Recent advances in hypersonic research activities have led to the development in ground based test facilities for simulating near-realistic flight conditions in the laboratory. Basic performance data such as drag, lift, thrust etc., are usually obtained from ground-based experimentation for physical understanding of the flow fields and subsequent prototype developments. Hypersonic flows are produced for a very short time in impulse facilities such as shock tunnels, free-piston shock tunnels and expansion tubes. The test times are in the order of few milliseconds in the shock tunnel and for the expansion tubes, it is still less (~50 μs). The force measurements in these facilities are very challenging due the need of fast response sensors and instrumentations. In this review paper, two force measurement techniques are discussed that are best suited for short duration facilities namely; accelerometer balance and stress-wave force balance system.
Content Type Journal Article Category Research Article Pages 31-58
DOI 10.1260/1759-3107.1.1.31
Authors
Niranjan Sahoo, Department of Mechanical Engineering, Indian Institute of Technology Guwahati, Guwahati - 781 039, INDIA K. P. J. Reddy, Department of Aerospace Engineering, Indian Institute of Science, Bangalore - 560 012, INDIA
Journal International Journal of Hypersonics
Abstract: Numerical simulations were carried out for various reacting flow fields related to scramjet propulsion system using commercial CFD Software. Three-dimensional Navier stokes equations were solved along with the K-ε turbulence model. Modeling of the turbulence chemistry interaction is done through infinitely fast rate chemical kinetics. The software was validated extensively by comparing different experimental conditions for scramjet combustor to find its error band and range of application. Good agreement between the experimental and computational values were obtained for the scramjet combustor flow field with strut, pylon and cavity injection system with both hydrogen and hydrocarbon fuel. The validated CFD tool was applied in the design exercise of flight-sized kerosene fuel scramjet combustor of a hypersonic airbreathing mission. Significant improvement of combustion efficiency and thrust could be achieved by relocating the fuel injection system through the analysis of various thermochemical variables in the scramjet combustor.
Content Type Journal Article Category Research Article Pages 13-30
DOI 10.1260/1759-3107.1.1.13
Authors
Debasis Chakraborty, Directorate of Computational Dynamics, Defence Research and Development Laboratory, Hyderabad - 500058, INDIA
Journal International Journal of Hypersonics
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