Dissertations / Theses: 'Hypersonic Aerodynamics'

1

Khorrami, Ahmad Farid. "Hypersonic aerodynamics on flat plates and thin aerofoils." Thesis, University of Oxford, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.292584.

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2

Hunt, Dillon C. "Measurement of ablation in transient hypersonic flows /." St. Lucia, Qld, 2001. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe16475.pdf.

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3

Singh, Amarjit. "Experimental study of slender vehicles at hypersonic speeds." Thesis, Cranfield University, 1996. http://hdl.handle.net/1826/4257.

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An experimental investigation of the hypersonic flow over (i) a wing-body configuration, (ii) a hemi-spherically blunted cone-cylinder body and (iii) a one-half- power-law body has been conducted for M,, = 8.2 and Re,, = 9.35x104 per cm. The tests were performed at model incidences, a=0,5 and 10° for flap deflection angles, (3 = 0,5,15, and 25° for the wing-body. The incidence ranged from -3 to 10° for the cone- cylinder and -5 to 15° for the power-law body. (i) The schlieren pictures showing top and side views of the model indicate that the body nose shock does not intersect the wing throughout the range of a under investigation. Detailed pressure measurements on the lower surface of the wing and flap along with the liquid crystal pictures suggest that the body nose shock does not strike the flap surfaces either. The wing leading edge shock is found to be attached at a=0 and 5° but detached at a= 10°. The liquid crystal pictures and surface pressure measurements indicated attached flow on the lower surface of the wing and flap for 13 =0 and 5° at all values of a under test. However at a= 0°, as the flap angle is increased to 15° the flow separates ahead of the hinge line. As incidence is increased the boundary layer becomes transitional giving rise to complex separation patterns around the flap hinge line. The spherically blunted body nose causes strong entropy layer effects over the wing and the trailing edge flap. A Navier-Stokes solution indicated a thick entropy layer of approximately constant thickness all around the cylindrical section of the body at zero incidence. However, at an incidence of 10° the layer tapers and becomes thinner under the body. The surface pressure over the wing and the plateau pressure for separated flow was found to increase from the root to the tip. This is partly because of the decrease in local Reynolds number across the span, however in the present case, entropy layer effects also affected separation. The entropy layer effects were found to reduce the peak pressures obtainable on the flap. The peak pressures, over the portion of the flap unaffected by entropy layer effects, could be estimated assuming quasi two dimensional flow. (ii) Force measurements were made for the blunted cone-cylinder alone as well as with the delta wing, with trailing-edge flap, attached to it. The lift, drag, and pitching moment characteristics for the cone-cylinder agree reasonably well with the modified Newtonian theory and the N-S results. The addition of a wing to the cone-cylinder body increases the lift as weil as the drag coefficient but there is an overall increase in the lift/drag ratio. The deflection of a flap from 0° to 25° increases the lift and drag coefficients at all the incidences tested. However, the lift/drag ratio is reduced showing the affects of separation over the wing. The experimental results on the wing-body are compared with the theoretical estimates based upon two-dimensional shock-expansion theory. (iii) The lift, and drag characteristics of a one-half-power-law body are compared with other existing results. The addition of strakes to the power-law body are found to improve its aerodynamic efficiency without any significant change in its pitching moment characteristics.

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4

Robinson, Matthew J. "Simultaneous lift, moment and thrust measurement on a scramjet in hypervelocity flow /." [St. Lucia, Qld.], 2003. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe17611.pdf.

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5

Babinsky, Holger. "A study of roughness in turbulent hypersonic boundary-layers." Thesis, Cranfield University, 1993. http://dspace.lib.cranfield.ac.uk/handle/1826/7586.

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The influence of large scale regular roughness on a Mach 5 turbulent boundary layer and a compression corner was investigated on axisymmetric wind tunnel models. Three types of roughness were examined; a series of square cavities at two different sizes and a 45 degree sawtooth. Typical sizes ranged from 50% to 100% of an undisturbed boundary layer thickness. The roughness was limited to a short region followed by a smooth surface. Compression corners were formed by 15° and 20° flares located downstream of the roughness. The flow in the wind tunnel was investigated in detail to obtain knowledge on operating conditions and flow quality. Liquid crystal thermography was developed for routine use in hypersonic blow-down wind tunnels with superior spatial resolution and experimental uncertainties in the range of traditional techniques. The effect on flow parameters downstream of the last roughness element were 7, found to differ significantly for the different quantities. Velocity profiles were found i, to be less full and skin friction was found to be reduced for all streamwise "~ distances. Surface heat transfer was increased in a short region limited to 1.5 boundary layer thicknesses behind the roughness whereas surface pressure was not affected. Sawtooth shaped roughness was found to cause a stronger j disturbance than square cavities of twice the size. Little influence of the roughness was noted on the flow over the compression corner. The flow over the 20° compression corner showed an increase in upstream influence for the sawtooth shaped roughness as well as the larger cavities. Surface pressure measurements did not indicate a separation in any case. Heat transfer measurements revealed a peak located approximately 0.25 boundary layer thicknesses behind the corner. No such feature was found in the surface pressure distributions. It is suggested that a small scale separation is located very close to the corner causing the peak in heat transfer at reattachment without any effect on surface pressures. The existence of such a separation has been confirmed by surface flow visualisations for both flares.

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6

Asproulis, Panagiotis. "High resolution numerical predictions of hypersonic flows on unstructured meshes." Thesis, Imperial College London, 1994. http://hdl.handle.net/10044/1/8357.

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7

Surah, Davinder. "Investigation of attachment line boundary layer characteristics in hypersonic flows." Thesis, Cranfield University, 2000. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.323921.

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8

Atcliffe, Phillip Arthur. "Effects of boundary layer separation and transition at hypersonic speeds." Thesis, Cranfield University, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.336458.

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9

Wilson, Althea Grace. "Numerical study of energy utilization in nozzle/plume flow-fields of high-speed air-breathing vehicles." Diss., Rolla, Mo. : Missouri University of Science and Technology, 2008. http://scholarsmine.mst.edu/thesis/pdf/Wilson_09007dcc804d881b.pdf.

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Thesis (M.S.)--Missouri University of Science and Technology, 2008.
Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed April 25, 2008) Includes bibliographical references (p. 57).

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10

Modlin, James Michael. "Hypersonic aerospace vehicle leading edge cooling using heat pipe, transpiration and film cooling techniques." Diss., Georgia Institute of Technology, 1991. http://hdl.handle.net/1853/16347.

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11

Axdahl, Erik Lee. "A study of premixed, shock-induced combustion with application to hypervelocity flight." Diss., Georgia Institute of Technology, 2013. http://hdl.handle.net/1853/50290.

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One of the current goals of research in hypersonic, airbreathing propulsion is access to higher Mach numbers. A strong driver of this goal is the desire to integrate a scramjet engine into a transatmospheric vehicle airframe in order to improve performance to low Earth orbit (LEO) or the performance of a semi-global transport. An engine concept designed to access hypervelocity speeds in excess of Mach 10 is the shock-induced combustion ramjet (i.e. shcramjet). This dissertation presents numerical studies simulating the physics of a shcramjet vehicle traveling at hypervelocity speeds with the goal of understanding the physics of fuel injection, wall autoignition mitigation, and combustion instability in this flow regime. This research presents several unique contributions to the literature. First, different classes of injection are compared at the same flow conditions to evaluate their suitability for forebody injection. A novel comparison methodology is presented that allows for a technically defensible means of identifying outperforming concepts. Second, potential wall cooling schemes are identified and simulated in a parametric manner in order to identify promising autoignition mitigation methods. Finally, the presence of instabilities in the shock-induced combustion zone of the flowpath are assessed and the analysis of fundamental physics of blunt-body premixed, shock-induced combustion is accelerated through the reformulation of the Navier Stokes equations into a rapid analysis framework. The usefulness of such a framework for conducting parametric studies is demonstrated.

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12

Morimoto, Hitoshi. "Trajectory optimization for a hypersonic vehicle with constraint." Diss., Georgia Institute of Technology, 1997. http://hdl.handle.net/1853/12076.

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13

Rohrschneider, Reuben R. "Variable-Fidelity Hypersonic Aeroelastic Analysis of Thin-Film Ballutes for Aerocapture." Diss., Georgia Institute of Technology, 2007. http://hdl.handle.net/1853/14590.

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Ballute hypersonic aerodynamic decelerators have been considered for aerocapture since the early 1980's. Recent technology advances in fabric and polymer materials as well as analysis capabilities lend credibility to the potential of ballute aerocapture. The concept of the thin-film ballute for aerocapture shows the potential for large mass savings over propulsive orbit insertion or rigid aeroshell aerocapture. Several technology hurdles have been identified, including the effects of coupled fluid structure interaction on ballute performance and survivability. To date, no aeroelastic solutions of thin-film ballutes in an environment relevant to aerocapture have been published. In this investigation, an aeroelastic solution methodology is presented along with the analysis codes selected for each discipline. Variable-fidelity aerodynamic tools are used due to the long run times for computational fluid dynamics or direct simulation Monte Carlo analyses. The improved serial staggered method is used to couple the disciplinary analyses in a time-accurate manner, and direct node-matching is used for data transfer. In addition, an engineering approximation has been developed as an addition to modified Newtonian analysis to include the first-order effects of damping due to the fluid, providing a rapid dynamic aeroelastic analysis suitable for conceptual design. Static aeroelastic solutions of a clamped ballute on a Titan aerocapture trajectory are presented using non-linear analysis in a representative environment on a flexible structure. Grid convergence is demonstrated for both structural and aerodynamic models used in this analysis. Static deformed shape, drag and stress level are predicted at multiple points along the representative Titan aerocapture trajectory. Results are presented for verification and validation cases of the structural dynamics and simplified aerodynamics tools. Solutions match experiment and other validated codes well. Contributions of this research include the development of a tool for aeroelastic analysis of thin-film ballutes which is used to compute the first high-fidelity aeroelastic solutions of thin-film ballutes using inviscid perfect-gas aerodynamics. Additionally, an aerodynamics tool that implements an engineering estimate of hypersonic aerodynamics with a moving boundary condition is developed and used to determine the flutter point of a thin-film ballute on a Titan aerocapture trajectory.

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14

Morrison, John William. "Auxiliary cooling in heat pipe cooled hypersonic wings." Thesis, Georgia Institute of Technology, 1990. http://hdl.handle.net/1853/17113.

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15

Bykerk, Tamas. "Low Speed Aerodynamics, Performance and Handling Qualities of a Hypersonic Waverider." Thesis, University of Sydney, 2020. https://hdl.handle.net/2123/23111.

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Interest in the development of high speed commercial aircraft capable of travelling in excess of Mach 5 is on the increase. This is mainly driven by the potential of these vehicles to significantly reduce long haul flight times, connecting antipodal city pairs such as Brussels and Sydney in less than three hours. The hypersonic waverider concept has received particular attention because of its ability to achieve comparatively high aerodynamic efficiency at hypersonic cruise, as the body shock is contained by the wing leading edges. As research into the extreme conditions at high Mach numbers takes priority, no full subsonic handling quality analysis of a waverider has ever been completed, nor has an understanding of the mechanisms behind both static and dynamic stability been developed. This gap in literature is addressed by this thesis, which presents a combination of results from computational fluid dynamics simulations and wind tunnel experiments, completing the first full subsonic handling quality analysis of a waverider.

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16

Garrett, Joseph Lee. "A comparison of flux-splitting algorithms for the Euler equations with equilibrium air chemistry." Thesis, Virginia Tech, 1989. http://hdl.handle.net/10919/44636.

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The use of flux-splitting techniques on the Euler equations is considered for high Mach number, high temperature flows in which the fluid is assumed to be inviscid air in equilibrium. Three different versions of real gas extensions to the Stegerâ Warming and Van Leer flux-vector splitting, and four different versions of real gas extensions to the Roe flux-difference splitting, are compared with regard to general applicability and ease of implementation in existing perfect gas g algorithms. Test computations are performed for the M = 5, high temperature flow over a 10-degree wedge and the M = 24.5 flow over a blunt body. Although there were minor differences between the computed results for the three types of flux-splitting algorithms considered, little variation is observed between different versions of the same algorithm.

Master of Science

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17

Higgins, Andrew J. "Investigation of detonation initiation by supersonic blunt bodies /." Thesis, Connect to this title online; UW restricted, 1996. http://hdl.handle.net/1773/10000.

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18

O'Kresik, Stephen R. "Design and optimization of a hypersonic test facility for sub-scale testing." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Dec%5FOKresik.pdf.

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Thesis (M.S. in Astronautical Engineering)--Naval Postgraduate School, December 2003.
Thesis advisor(s): Jose O. Sinibaldi, Garth V. Hobson. Includes bibliographical references (p. 69). Also available online.

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19

Taflin, David E. "Numerical simulation of unsteady hypersonic chemically reacting flow /." Thesis, Connect to this title online; UW restricted, 1995. http://hdl.handle.net/1773/9967.

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20

Wheatley, Vincent. "Modelling low-density flow in hypersonic impulse facilities /." [St. Lucia, Qld.], 2001. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe16173.pdf.

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21

Grant, Michael James. "Rapid simultaneous hypersonic aerodynamic and trajectory optimization for conceptual design." Diss., Georgia Institute of Technology, 2012. http://hdl.handle.net/1853/43685.

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Traditionally, the design of complex aerospace systems requires iteration among segregated disciplines such as aerodynamic modeling and trajectory optimization. Multidisciplinary design optimization algorithms have been developed to efficiently orchestrate the interaction among these disciplines during the design process. For example, vehicle capability is generally obtained through sequential iteration among vehicle shape, aerodynamic performance, and trajectory optimization routines in which aerodynamic performance is obtained from large pre-computed tables that are a function of angle of attack, sideslip, and flight conditions. This numerical approach segregates advancements in vehicle shape design from advancements in trajectory optimization. This investigation advances the state-of-the-art in conceptual hypersonic aerodynamic analysis and trajectory optimization by removing the source of iteration between aerodynamic and trajectory analyses and capitalizing on fundamental linkages across hypersonic solutions. Analytic aerodynamic relations, like those derived in this investigation, are possible in any flow regime in which the flowfield can be accurately described analytically. These relations eliminate the large aerodynamic tables that contribute to the segregation of disciplinary advancements. Within the limits of Newtonian flow theory, many of the analytic expressions derived in this investigation provide exact solutions that eliminate the computational error of approximate methods widely used today while simultaneously improving computational performance. To address the mathematical limit of analytic solutions, additional relations are developed that fundamentally alter the manner in which Newtonian aerodynamics are calculated. The resulting aerodynamic expressions provide an analytic mapping of vehicle shape to trajectory performance. This analytic mapping collapses the traditional, segregated design environment into a single, unified, mathematical framework which enables fast, specialized trajectory optimization methods to be extended to also include vehicle shape. A rapid trajectory optimization methodology suitable for this new, mathematically integrated design environment is also developed by relying on the continuation of solutions found via indirect methods. Examples demonstrate that families of optimal hypersonic trajectories can be quickly constructed for varying trajectory parameters, vehicle shapes, atmospheric properties, and gravity models to support design space exploration, trade studies, and vehicle requirements definition. These results validate the hypothesis that many hypersonic trajectory solutions are connected through fast indirect optimization methods. The extension of this trajectory optimization methodology to include vehicle shape through the development of analytic hypersonic aerodynamic relations enables the construction of a unified mathematical framework to perform rapid, simultaneous hypersonic aerodynamic and trajectory optimization. Performance comparisons relative to state-of-the-art methodologies illustrate the computational advantages of this new, unified design environment.

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22

Song, Dong Joo. "Hypersonic nonequilibrium flow over an ablating teflon surface." Diss., Virginia Polytechnic Institute and State University, 1986. http://hdl.handle.net/10919/71192.

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A complex chemical system of teflon/air mixture over an axisymmetric decoy at hypersonic reentry flight conditions has been analyzed by using the nonequilibrium viscous shock-layer method. The equilibrium catalytic wall boundary condition was used to obtain the species concentration at the wall. The species conservation equation for binary mixture (air/teflon) was solved to obtain the concentration of freestream air at the wall. Two test cases were chosen to demonstrate the capability of the current code. Due to lack of experimental or theoretical data, the surface measurable quantities from the current code(VSLTEF) were compared with the equivalent air injection and no-mass injection data obtained from VSL7S code. The current code predicts a higher total heat-transfer rate than that predicted by the seven species nonequilibrium air code (VSL7S) with the same injection rate due to the high diffusional heat-transfer rate. The wall pressure was not affected by blowing, while the skin-friction coefficient was decreased (i.e., 43 % reduction for teflon ablation case ; 53 % for nonequilibrium air injection case at 125 kft) when compared with that of no-mass injection case. A shock-layer peak temperature drop ( 1512° R for 125 kft altitude and 848°R for 175 kft altitude) was observed at both cases. The temperature drops were chiefly due to endothermic reactions (dissociation) of the teflon ablation species. Due to large blowing of teflon, the average molecular weight increased substantially and resulted in a reduction of the specific heat ratio γ and an increase in the Prandtl number at the wall. The impurity of sodium was the major source of free electrons near the wall at the end of the vehicle at 125 kft altitude; however, at 175 kft altitude NO⁺ was the major source of free electrons over the entire body. The peak concentration of Na⁺ increased along the body, but that of NO⁺ decreased at both altitudes; While the chemical reaction rate data used is believed to be the best currently available, uncertainties in this data as were cited by Cresswell et al.(1967) may lead to quantitative changes in the above teflon ablation results.
Ph. D.

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23

Owen, Andrew Kevin. "Experimental studies of the hypersonic, low density, aerodynamics of re-entry vehicles." Thesis, University of Oxford, 1997. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.298680.

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24

Smith, Andrew John Darwin. "The dynamic response of a wedge separated hypersonic flow and its effects on heat transfer." Thesis, University of Southampton, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.357111.

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25

Drauch, Gregory Andrew. "Hypersonic test facilities: requirements analysis and preliminary design." Thesis, Virginia Tech, 1990. http://hdl.handle.net/10919/41905.

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26

Fuller, Eric James. "Experimental and computational investigation of helium injection into air at supersonic and hypersonic speeds." Diss., Virginia Tech, 1992. http://hdl.handle.net/10919/39977.

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Experiments were performed with two different helium injector models at different injector transverse and yaw angles in order to determine the mixing rate and core penetration of the injectant and the flow field total pressure losses. when gaseous injection occurs into a supersonic freestream. Tested in the Virginia Tech supersonic tunnel. with a freestream Mach number of 3.0 and conditions corresponding to a freestream Reynolds number of 5.0 x 107 1m. was a single. sonic. 5X underexpanded, helium jet at a downstream angle of 30° relative to the freestream. This injector was rotated from 0° to _28° to test the effects of injector yaw. The second model was an array of three supersonic, 5X underexpanded helium injectors with an exit Mach number of 1.7 and a transverse angle of 15°. This model was tested in the NASA Langley Mach 6.0, High Reynolds number tunnel, with freestream conditions corresponding to a Reynolds number of 5.4 x 10⁷ /m. The injector array as tested at yaw angles of 0° and -15°. Surface flow visualization showed that significant flow asymmetries were produced by injector yaw. Nanosecond exposure shadowgraph pictures were taken, showing the gaseous injection plume to be unsteady, and further studies demonstrated this unsteadiness was related to shock waves orthogonal to the injectant bow shock, that were generated at a frequency of 30 kHz. The primary data technique used, was a concentration probe which measured the molar concentration of helium in the flow field. Concentration data and other meanflow data was taken at several downstream axial stations and yielded contours of helium concentration, total pressure, Mach number, velocity, and mass flux, as well as the static properties. From these contour plots, the various mixing rates for each case were determined. The injectant mixing rates, expressed as the maximum concentration decay, and mixing distances were found to be unaffected by injector yaw, in the Mach 3.0 experiments, but were adversely affected by injector yaw in the Mach 6.0 experiments. One promising aspect of injector yaw was the that as the yaw angle was increased, lateral motion of the injectant plume became significant, and the turbulent mixing region area increased by approximately 34%. Comparisons of the 15° transverse angled injection into a Mach 6.0 flow to previous experiments with 15° injection into a Mach 3.0 freestream, demonstrated that there is a significant decrease in initial mixing, at Mach 6.0, resulting in a much longer mixing distance. From a parametric computational study of the Mach 6.0 experiments, the effects of adjacent injectors was found to decrease lateral spreading while increasing the vertical penetration of the injectant plume, and marginally increasing the injectant core decay rate. Matching of the computational results to the experimental results was best achieved when using the Baldwin-Lomax turbulence model without the Degani-Schiff modification.
Ph. D.

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27

Shi, Yijian. "Off-design waverider flowfield CFD simulation /." free to MU campus, to others for purchase, 1996. http://wwwlib.umi.com/cr/mo/fullcit?p9717164.

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28

Shek, H. H.-W. "A study of moderately underexpanded single and twinjet rocket exhaust plumes in quiescent and in a mach 7 hypersonic freestream." Thesis, University of Oxford, 1997. http://ora.ox.ac.uk/objects/uuid:b735b5db-8d12-46b1-86aa-ccc4996e5e3e.

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Rocket plume flowfields have an importance due to their influence on the signature of the rocket and also on the distribution of the plume gases around the vehicle. Little information on the co-flowing situation exists other than a previous study at Oxford. This thesis thus represents a significant database for co-flowing rocket plumes of this form. The work presented deals with two new aspects of co- flowing rocket plumes in that detailed flowfield measurements have been made and plumes from twin nozzle have been investigated for the first time in this thesis. This study on twinjet rocket plumes was carried out using the University of Oxford Gun Tunnel. Twinjet rockets with nozzle exit Mach numbers of 3 and 5 were tested in quiescent and in co-flow at Mach 7 using nitrogen and hydrogen injections. A major feature of the twinjet case was the so-called impingement shock between the flows from the two nozzles. It was discovered that this shock was insensitive to the freestream and scaling parameters are suggested for its geometry. Comparisons with single equivalent thrust nozzles are made at downstream locations and similar Pitot pressure profiles were observed for nitrogen injection in a nitrogen freestream after approximately 3 nozzle diameters downstream. Shear layers were studied and fluctuations in this region were measured by fast-response Pitot pressure and heat transfer probes sampled at 1.1 MHz. The extent of the shear layer was deduced using a new Oxford Total Temperature Probe. With the freestream stagnation temperature at approximately 650 K and injected gas at 350 K, a linear variation for the deduced total temperature across the shear layer was obtained. This was consistent with the Pitot pressure variations across this region. Convective heat transfer coefficient fluctuations and flow total temperature fluctuations across rocket flowiields were obtained using three thin-film heat transfer probes and found to be closely correlated. Experimental results for the twinjet and the single jet were compared with CFD simulations and good overall agreements were achieved. Instrumentation for the hypersonic experiments was investigated and a fast-response (~ 20 kHz) Pitot probe suited for flows heavily contaminated with particulate was developed and tested.

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29

Horvath, Istvan. "Development of a cantilever beam, capacitive sensing, skin friction gage and supporting instrumentation for measurements /." This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-06162009-063052/.

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30

Lee, Jaewoo. "Efficient inverse methods for supersonic and hypersonic body design, with low wave drag analysis." Diss., Virginia Tech, 1991. http://hdl.handle.net/10919/37406.

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With the renewed interest in the supersonic and hypersonic flight vehicles, new inverse Euler methods are developed in these flow regimes where a space marching numerical technique is valid. In order to get a general understanding for the specification of target pressure distributions, a study of minimum drag body shapes was conducted over a Mach number range from 3 to 12. Numerical results show that the power law bodies result in low drag shapes, where the n=.69 (l/d = 3) or n=.70 (l/d = 5) shapes have lower drag than the previous theoretical results (n=.75 or n=.66 depending on the particular form of the theory). To validate the results, a numerical analysis was made including viscous effects and the effect of gas model. From a detailed numerical examination for the nose regions of the minimum drag bodies, aerodynamic bluntness and sharpness are newly defined. Numerous surface pressure-body geometry rules are examined to obtain an inverse procedure which is robust, yet demonstrates fast convergence. Each rule is analyzed and examined numerically within the inverse calculation routine for supersonic (M_∞= 3) and hypersonic (M_∞ = 6.28) speeds. Based on this analysis, an inverse method for fully three dimensional supersonic and hypersonic bodies is developed using the Euler equations. The method is designed to be easily incorporated into existing analysis codes, and provides the aerodynamic designer with a powerful tool for design of aerodynamic shapes of arbitrary cross section. These shapes can correspond to either "wing like" pressure distributions or to "body like" pressure distributions. Examples are presented illustrating the method for a non-axisymmetric fuselage type pressure distribution and a cambered wing type application. The method performs equally well for both nonlifting and lifting cases. For the three dimensional inverse procedure, the inverse solution existence and uniqueness problem are discussed. Sample calculations demonstrating this problem are also presented.
Ph. D.

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31

Comstock, Robert. "Hypersonic Heat Transfer Load Analysis in STAR-CCM+." DigitalCommons@CalPoly, 2020. https://digitalcommons.calpoly.edu/theses/2226.

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This thesis investigates the capabilities of STAR-CCM+, a Computational Fluid Dynamics (CFD) software owned by Siemens, in predicting hypersonic heat transfer loads on forward-facing surfaces. Results show that STAR-CCM+ predicted peak heat transfer loads within +/- 20% of experimental data on the leading edge of a delta wing design from the X-20 Dyna-Soar program with 73o of sweep. Steady-state laminar simulations were run as replications of wind tunnel tests documented in NASA CR-535, a NASA technical report that measured and studied the hypersonic pressure and heat transfer loads on preliminary X- 20 wing designs across a wide range of Reynolds numbers and Mach numbers in different wind tunnel and shock tunnel facilities. One of the Mach 8.08 test cases that was run at NASA Arnold Engineering Development Center Wind Tunnel B was selected as the case of comparison for this thesis, which was designated as test AD462M-1 in the original report. The CFD simulations assumed an ideal gas in laminar flow with temperature-dependent viscosity, thermal conductivity, and isobaric specific heat across an angle of attack range from 0o to 30o. A separate CFD study of heat transfer loads of a hemisphere-cylinder at Mach 6.74 was used as a simpler and less computationally-expensive validation case compared against wind tunnel data from NASA Langley Research Center to help select the appropriate CFD solver and mesh settings for this thesis. For the hemisphere-cylinder, the heat transfer load at the stagnation point was overpredicted in STAR-CCM+ by 21.8%. Peak heat transfer loads on the delta wing leading edge were all within +/- 20% of the wind tunnel data, which was published for angles of attack between 15o to 30o. A more adverse heat transfer gradient along the leading edge of the delta wing was also observed in the direction from the front of the wing to the outer wing tip when compared to wind tunnel data. The pressure loads on the delta wing leading edge in CFD were within +/-10% of wind tunnel measurements.

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32

Hayne, Michael J. "Hypervelocity flow over rearward-facing steps /." [St. Lucia, Qld], 2004. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe18242.pdf.

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33

Gupta, Anurag. "The artificially blunted leading edge concept for aerothermodynamic performance enhancement." Diss., Georgia Institute of Technology, 1999. http://hdl.handle.net/1853/12442.

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34

Fort, James A. "A numerical study of attached oblique detonation /." Thesis, Connect to this title online; UW restricted, 1993. http://hdl.handle.net/1773/7087.

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35

Zoebelein, Till. "Development of an LU-scheme for the solution of hypersonic non-equilibrium flow." Thesis, Georgia Institute of Technology, 1998. http://hdl.handle.net/1853/12509.

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36

Cinnella, Pasquale. "Flux-split algorithms for flows with non-equilibrium chemistry and thermodynamics." Diss., Virginia Polytechnic Institute and State University, 1989. http://hdl.handle.net/10919/54506.

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New flux-split algorithms are developed for high velocity, high-temperature flow situations, when finite-rate chemistry and non-equilibrium thermodynamics greatly affect the physics of the problem. Two flux-vector-split algorithms, of the Steger-Warming and of the Van Leer type, and one flux-difference-split algorithm of the Roe type are established and utilized for the accurate numerical simulation of flows with dissociation, ionization, and combustion phenomena. Several thermodynamic models are used, including a simplified vibrational non-equilibrium model and an equilibrium model based upon refined statistical mechanics properties. The framework provided is flexible enough to accommodate virtually any chemical model and a wide range of non-equilibrium, multi-temperature thermodynamic models. A theoretical study of the main features of flows with free electrons, for conditions that require the use of two translational temperatures in the thermal model, is developed. Interesting and unexpected results are obtained, because acoustic wave speeds of the symmetric form u±α no longer appear. A simple but powerful asymptotic analysis is developed which allows the establishment of the fundamental gas-dynamic properties of flows with multiple translational temperatures. The new algorithms developed demonstrate their accuracy and robustness for challenging flow problems. The influence of several assumptions on the chemical and thermal behavior of the flows is investigated, and a comparison with results obtained using different numerical approaches, in particular spectral methods, is provided, and proves to be favorable to the present techniques. Other calculations in one and two space dimensions indicate large sensitivities with respect to chemical and thermodynamic modeling. The algorithms developed are of sufficient generality to begin to examine these effects in detail. Preliminary numerical simulations are performed using elementary modeling of transport phenomena.
Ph. D.

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37

Rohweder, Matthew Flynn. "A numerical investigation of flowfield modification in high-speed airbreathing inlets using energy deposition." Diss., Rolla, Mo. : Missouri University of Science and Technology, 2010. http://scholarsmine.mst.edu/thesis/pdf/Rohweder_09007dcc80722a47.pdf.

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Thesis (M.S.)--Missouri University of Science and Technology, 2010.
Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed Jan. 5, 2010). Includes bibliographical references (p. 52-53).

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38

Sekhar, Susheel Kumar. "Viscous hypersonic flow physics predictions using unstructured Cartesian grid techniques." Diss., Georgia Institute of Technology, 2012. http://hdl.handle.net/1853/45857.

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Aerothermodynamics is an integral component in the design and implementation of hypersonic transport systems. Accurate estimates of the aerodynamic forces and heat transfer rates are critical in trajectory analysis and for payload weight considerations. The present work seeks to investigate the ability of an unstructured Cartesian grid framework in modeling hypersonic viscous flows. The effectiveness of modeling viscous phenomena in hypersonic flows using the immersed boundary ghost cell methodology of this solver is analyzed. The capacity of this framework to predict the surface physics in a hypersonic non-reacting environment is investigated. High velocity argon gas flows past a 2-D cylinder are simulated for a set of freestream conditions (Reynolds numbers), and impact of the grid cell sizes on the quality of the solution is evaluated. Additionally, the formulation is verified over a series of hypersonic Mach numbers for the flow past a hemisphere, and compared to experimental results and empirical estimates. Next, a test case that involves flow separation and the interaction between a hypersonic shock wave and a boundary layer, and a separation bubble is investigated using various adaptive mesh refinement strategies. The immersed boundary ghost cell approach is tested with two temperature clipping strategies, and their impact on the overall solution accuracy and smoothness of the surface property predictions are compared. Finally, species diffusion terms in the conservation equations, and collision cross-section based transport coefficients are installed, and hypersonic flows in thermochemical nonequilibrium environments are studied, and comparisons of the off-surface flow properties and the surface physics predictions are evaluated. First, a 2-D cylinder in a hypersonic reacting air flow is tested with an adiabatic wall boundary condition. Next, the same geometry is tested to evaluate the viscous chemistry prediction capability of the solver with an isothermal wall boundary condition, and to identify the strengths and weaknesses of the immersed boundary ghost cell methodology in computing convective heating rates in such an environment.

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39

Stewart, Benjamin S. "Predicted scramjet testing capabilities of the proposed RHYFL-X expansion tube /." [St. Lucia, Qld], 2004. http://www.library.uq.edu.au/pdfserve.php?image=thesisabs/absthe18241.pdf.

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40

Chuck, Chen. "Numerical simulation of oblique detonation and shock-deflagration waves with a laminar boundary-layer /." Thesis, Connect to this title online; UW restricted, 1990. http://hdl.handle.net/1773/9966.

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41

Tirtey, Sandy C. "Characterization of a transitional hypersonic boundary layer in wind tunnel and flight conditions." Doctoral thesis, Universite Libre de Bruxelles, 2009. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/210367.

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Laminar turbulent transition is known for a long time as a critical phenomenon influencing the thermal load encountered by hypersonic vehicle during their planetary re-entry trajectory. Despite the efforts made by several research laboratories all over the world, the prediction of transition remains inaccurate, leading to oversized thermal protection system and dramatic limitations of hypersonic vehicles performances. One of the reasons explaining the difficulties encountered in predicting transition is the wide variety of parameters playing a role in the phenomenon. Among these parameters, surface roughness is known to play a major role and has been investigated in the present thesis.

A wide bibliographic review describing the main parameters affecting transition and their coupling is proposed. The most popular roughness-induced transition predictions correlations are presented, insisting on the lack of physics included in these methods and the difficulties encountered in performing ground hypersonic transition experiments representative of real flight characteristics. This bibliographic review shows the importance of a better understanding of the physical phenomenon and of a wider experimental database, including real flight data, for the development of accurate prediction methods.

Based on the above conclusions, a hypersonic experimental test campaign is realized for the characterization of the flow field structure in the vicinity and in the wake of 3D roughness elements. This fundamental flat plate study is associated with numerical simulations for supporting the interpretation of experimental results and thus a better understanding of transition physics. Finally, a model is proposed in agreement with the wind tunnel observations and the bibliographic survey.

The second principal axis of the present study is the development of a hypersonic in-flight roughness-induced transition experiment in the frame of the European EXPERT program. These flight data, together with various wind tunnel measurements are very important for the development of a wide experimental database supporting the elaboration of future transition prediction methods.
Doctorat en Sciences de l'ingénieur
info:eu-repo/semantics/nonPublished

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42

Bhutta, Bilal A. "A new parabolized Navier-Stokes scheme for hypersonic reentry flows." Diss., Virginia Polytechnic Institute and State University, 1985. http://hdl.handle.net/10919/52287.

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High Mach number, low-Reynolds number (high-altitude) reentry flowfield predictions are an important problem area in computational aerothermodynamics. Available numerical tools for handling such flows are very few and significantly limited in their applicability. A new implicit fully-iterative Parabolized Navier-Stokes (PNS) scheme is developed to accurately predict such low-Reynolds number flows. In this new approach the differential equations governing the conservation of mass, momentum and energy, and the algebraic equation of state for a perfect gas are solved simultaneously in a coupled manner. The idea is presented that by treating the governing equations in this manner (rather than eliminating the pressure terms in the governing equations by using appropriate differentiated forms of the equation of state) it may be possible to have an unconditionally time-like numerical scheme. The stability of a simplified version of this new PNS scheme is also studied, and it is demonstrated that these simplified equations are unconditionally time-like in the subsonic as well as the supersonic flow regions. A pseudo-time integration approach is used in addition to a new second-order accurate fully-implicit smoothing, to improve the efficiency of the solution algorithm. The new PNS scheme is used to predict the flowfield around a seven-deg sphere-cone vehicle under high- and low-Reynolds number conditions. Two test case, Case A and Case B, are chosen such that Case A has a large freestream Reynolds number (2.92x10⁵), whereas Case B has a freestream Reynolds number of 1.72x10³, which is smaller than the usual limit of applicability of the non-iterative PNS schemes (Re~10⁴ or larger). Comparisons are made with other available numerical schemes, and the results substantiate the stability, accuracy and efficiency claims of the new Parabolized Navier-Stokes scheme.
Ph. D.

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43

Horvath, Istvan. "Development of a cantilever beam, capacitive sensing, skin friction gage and suppporting instrumentation for measurements." Thesis, Virginia Tech, 1993. http://hdl.handle.net/10919/43315.

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A cantilever beam type, capacitive sensing, skin friction gage has been developed. A prototype along with supporting electronics has been constructed. The cantilever beam gage is a change of area variable capacitive transducer. It is designed to measure the wall shear stress in a short duration, supersonic flow. The supporting electronics consists of an electrical oscillator for frequency modulation, and a frequency demodulator. The change in capacitance due to the shear stress in the flow modulates the output signal of the oscillator, which is then demodulated to extract a voltage signal which corresponds to the change in capacitance of the gage. The gage and the electronics were constructed from simple, inexpensive components for the purpose of proving the concept of a capacitive sensing transducer. static calibrations have been completed and statistical analysis has been done to test the performance of the gage. A 0.12 mV response due to the expected 98.1 g m/s2 force input of the skin friction of the Mach 2.9 design flow, over the 0.49 in2 (316.1 mm) area of the gage's sensing head, was measured as the average output of the skin friction gage instrumented with stainless steel strips.
Master of Science

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44

Malo-Molina, Faure Joel. "Numerical study of innovative scramjet inlets coupled to combustors using hydrocarbon-air mixture." Diss., Georgia Institute of Technology, 2010. http://hdl.handle.net/1853/33906.

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To advance the design of hypersonic vehicles, high-fidelity multi-physics CFD is used to characterize 3-D scramjet flow-fields in two novel streamline traced configurations. The two inlets, Jaws and Scoop, are analyzed and compared to a traditional rectangular inlet used as a baseline for on/off-design conditions. The flight trajectory conditions selected are Mach 6 and a dynamic pressure of 1,500 psf (71.82 kPa). Analysis of these hypersonic inlets is performed to investigate distortion effects downstream with multiple single cavity combustors acting as flame holders, and several fuel injection strategies. The best integrated scramjet inlet/combustor design is identified. The flow physics is investigated and the integrated performance impact of the two innovative scramjet inlet designs is quantified. Frozen and finite rate chemistry is simulated with 13 gaseous species and 20 reactions for an Ethylene/air finite-rate chemical model. In addition, URANS and LES modeling are compared to explore overall flow structure and to contrast individual numerical methods. The flow distortion in Jaws and Scoop is similar to some of the distortion in the traditional rectangular inlet, despite design differences. The baseline and Jaws performance attributes are stronger than Scoop, but Jaws accomplishes this while eradicating the cowl lip interaction, and lessening the total drag and spillage penalties. The innovative inlets work best on-design, whereas for off-design, the traditional inlet is best. Early pressure losses and flow distortions in the isolator aid the mixing of air and fuel, and improve the overall efficiency of the system. Although the trends observed with and without chemical reactions are similar, the former yields roughly 10% higher mixing efficiency and upstream reactions are present. These show a significant impact on downstream development. Unsteadiness in the combustor increases the mixing efficiency, varying the flame anchoring and combustion pressure effects upstream of the step.

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45

Benton, George Lynn. "Effects of a vibrationally excited gas on viscous shock-layer flows." Thesis, Virginia Polytechnic Institute and State University, 1985. http://hdl.handle.net/10919/104294.

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46

Dreyer, Emily Rose. "Assessment of Reduced Fidelity Modeling of a Maneuvering Hypersonic Vehicle." The Ohio State University, 2021. http://rave.ohiolink.edu/etdc/view?acc_num=osu1610018486409227.

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47

Eppard, William M. "Kinetic algorithms for non-equilibrium gas dynamics." Diss., This resource online, 1993. http://scholar.lib.vt.edu/theses/available/etd-06062008-165605/.

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48

(6624017), Joshua B. Edelman. "Nonlinear Growth and Breakdown of the Hypersonic Crossflow Instability." Thesis, 2019.

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A sharp, circular 7° half-angle cone was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel

at 6° angle of attack, extending several previous experiments on the growth and breakdown of

stationary crossflow instabilities in the boundary layer.

Measurements were made using infrared

imaging and surface pressure sensors. Detailed measurements of the stationary and traveling

crossflow vortices, as well as various secondary instability modes, were collected over a large

region of the cone.

The Rod Insertion Method (RIM) roughness, first developed for use on a flared cone, was

adapted for application to crossflow work. It was demonstrated that the roughness elements were

the primary factor responsible for the appearance of the specific pattern of stationary streaks

downstream, which are the footprints of the stationary crossflow vortices. In addition, a roughness

insert was created with a high RMS level of normally-distributed roughness to excite the naturally

most-amplified stationary mode.

The nonlinear breakdown mechanism induced by each type of roughness appears to be

different. When using the discrete RIM roughness, the dominant mechanism seems to be the

modulated second mode, which is significantly destabilized by the large stationary vortices. This

is consistent with recent computations. There is no evidence of the presence of traveling crossflow

when using the RIM roughness, though surface measurements cannot provide a complete picture.

The modulated second mode shows strong nonlinearity and harmonic development just prior

to breakdown. In addition, pairs of hot streaks merge together within a constant azimuthal

band, leading to a peak in the heating simultaneously with the peak amplitude of the measured

secondary instability. The heating then decays before rising again to turbulent levels. This nonmonotonic

heating pattern is reminiscent of experiments on a flared cone and earlier computations

of crossflow on an elliptic cone.

When using the distributed roughness there are several differences in the nonlinear breakdown

behavior. The hot streaks appear to be much more uniform and form at a higher wavenumber,

which is expected given computational results. Furthermore, the traveling crossflow waves become

very prominent in the surface pressure fluctuations and weakly nonlinear. In addition there

appears in the spectra a higher-frequency peak which is hypothesized to be a type-I secondary instability

under the upwelling of the stationary vortices. The traveling crossflow and the secondary

instability interact nonlinearly prior to breakdown.

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49

Srinath, S. "Experimental Study Of Large Angle Blunt Cone With Telescopic Aerospike Flying At Hypersonic Mach Numbers." Thesis, 2009. https://etd.iisc.ac.in/handle/2005/1116.

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The emerging and competitive environment in the space technology requires the improvements in the capability of aerodynamic vehicles. This leads to the analysis in drag reduction of the vehicle along with the minimized heat transfer rate. Using forward facing solid aerospike is the simplest way among the existing drag reduction methodologies for hypersonic blunt cone bodies. But the flow oscillations associated with this aerospike makes it difficult to implement. When analyzing this flow, it can be understood that this oscillating flow can be compared to conical cavity flow. Therefore in the spiked flows, it is decided to implement the technique used in reducing the flow oscillation of the cavities. Based on this method the shallow conical cavity flow generated by the aerospike fixed ahead of a 120o blunt cone body is fissured as multiple cavities by so many disks formed from 10o cone. Now the deep conical cavities had the length to mean depth ratio of unity; this suppresses the unnecessary oscillations of the shallow cavity. The total length of the telescopic aerospike is fixed as 100mm. And one another conical tip plain aerospike of same length is designed for comparing the telescopic spike’s performance at hypersonic flow Mach numbers of 5.75 and 7.9. A three component force balance system capable of measuring drag, lift and pitching moment is designed and mounted internally into the skirt of the model. Drag measurement is done for without spike, conical tip plain spiked and telescopic spiked blunt cone body. The three configurations are tested at different angles of attack from 0 to 10 degree with a step of 2. A discrete iterative deconvolution methodology is implemented in this research work for obtaining the clean drag history from the noisy drag accelerometer signal. The drag results showed the drag reduction when compared to the without spike blunt cone body. When comparing to the plain spiked, the telescopic spiked blunt cone body has lesser drag at higher angles of attack. Heat transfer measurements are done over the blunt cone surface using the Platinum thin film gauges formed over the Macor substrate. These results and the flow visualization give better understanding of the flow and the heat flux rate caused by the flow. The enhancement in the heat flux rate over the blunt cone surface is due to the shock interaction. And in recirculation region the heat flux rate is very much lesser when compared to without spike blunt cone body. It is observed that the shock interaction in the windward side is coming closer towards the nose of the blunt cone as the angle of attack increases and the oscillation of the oblique shock also decreases. Schlieren visualization showed that there is dispersion in the oblique shock, particularly in the leeward side. In the telescopic spike there are multiple shocks generated from each and every disk which coalesces together to form a single oblique shock. And the effect of the shock generated by the telescopic spike is stronger than the effect of the shock generated by the conical tip plain spike.

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50

Srinath, S. "Experimental Study Of Large Angle Blunt Cone With Telescopic Aerospike Flying At Hypersonic Mach Numbers." Thesis, 2009. http://hdl.handle.net/2005/1116.

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Abstract:

The emerging and competitive environment in the space technology requires the improvements in the capability of aerodynamic vehicles. This leads to the analysis in drag reduction of the vehicle along with the minimized heat transfer rate. Using forward facing solid aerospike is the simplest way among the existing drag reduction methodologies for hypersonic blunt cone bodies. But the flow oscillations associated with this aerospike makes it difficult to implement. When analyzing this flow, it can be understood that this oscillating flow can be compared to conical cavity flow. Therefore in the spiked flows, it is decided to implement the technique used in reducing the flow oscillation of the cavities. Based on this method the shallow conical cavity flow generated by the aerospike fixed ahead of a 120o blunt cone body is fissured as multiple cavities by so many disks formed from 10o cone. Now the deep conical cavities had the length to mean depth ratio of unity; this suppresses the unnecessary oscillations of the shallow cavity. The total length of the telescopic aerospike is fixed as 100mm. And one another conical tip plain aerospike of same length is designed for comparing the telescopic spike’s performance at hypersonic flow Mach numbers of 5.75 and 7.9. A three component force balance system capable of measuring drag, lift and pitching moment is designed and mounted internally into the skirt of the model. Drag measurement is done for without spike, conical tip plain spiked and telescopic spiked blunt cone body. The three configurations are tested at different angles of attack from 0 to 10 degree with a step of 2. A discrete iterative deconvolution methodology is implemented in this research work for obtaining the clean drag history from the noisy drag accelerometer signal. The drag results showed the drag reduction when compared to the without spike blunt cone body. When comparing to the plain spiked, the telescopic spiked blunt cone body has lesser drag at higher angles of attack. Heat transfer measurements are done over the blunt cone surface using the Platinum thin film gauges formed over the Macor substrate. These results and the flow visualization give better understanding of the flow and the heat flux rate caused by the flow. The enhancement in the heat flux rate over the blunt cone surface is due to the shock interaction. And in recirculation region the heat flux rate is very much lesser when compared to without spike blunt cone body. It is observed that the shock interaction in the windward side is coming closer towards the nose of the blunt cone as the angle of attack increases and the oscillation of the oblique shock also decreases. Schlieren visualization showed that there is dispersion in the oblique shock, particularly in the leeward side. In the telescopic spike there are multiple shocks generated from each and every disk which coalesces together to form a single oblique shock. And the effect of the shock generated by the telescopic spike is stronger than the effect of the shock generated by the conical tip plain spike.

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Dissertations / Theses on the topic 'Hypersonic Aerodynamics'

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