Dissertations / Theses: 'Hypersonic'

1

Kumar, D. "Hypersonic control effectiveness." Thesis, Cranfield University, 1995. http://hdl.handle.net/1826/4252.

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The present study analyses the effects of a number of geometric parameters on the performance of a trailing edge control flap on a hypersonic body. The tests were conducted in a gun tunnel at Mach 8.2 and Mach 12.3. The study revealed that flap deflection promoted separation lengthscales and boundary layer transition. The latter significantly increased the local aerothermal loads on the flap. For well separated flows, flap heat transfer rates were successfully predicted by reference temperature theory. The promotion of transition caused a progressive reduction in the lengthscales of separated flows. In a free-flight environment, vehicle incidence varies considerably. Incidence was found to promote transition on both flat plates and control flaps. The latter resulted in a considerable increase in flap heat transfer. A modified version of reference temperature theory successfully predicted the aerothermal loads on the flap. For laminar and transitional interactions, the separated flow lengthscale was found to have a complex variation with incidence. A number of relevant flow parameters were identified. The intense heat loads on a vehicle in hypersonic flight dictates the blunting of the leading edge. This strengthens the leading edge shock structure and generates an entropy layer. Bluntness was found to significantly decrease the separation interaction scales on the flap. This was due to a reduction in the pressure recovered on the flap. The latter adverse affects control effectiveness. The aerothermal loads on the control flap was successfully predicted by reference temperature theory. An investigation into the efficiency of an under-expanded transverse jet controls was conducted on an axi-symmetric slender blunt cone. Force measurements found that the interaction augmented the jet reaction force by 70% at zero incidence. This increased to 110% at low incidence. The experiments found that the scale of the interaction region was determined by Poj/pes. Using this parameter, a closed loop algorithm for the shape of the separation front was developed. The latter can be used to predict jet reaction control effectiveness.

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Netterfield, Michael Phillip. "Hypersonic cavity flows." Thesis, Imperial College London, 1989. http://hdl.handle.net/10044/1/47586.

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3

Gorishnyy, Taras. "Hypersonic phononic crystals." Thesis, Massachusetts Institute of Technology, 2007. http://hdl.handle.net/1721.1/42133.

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Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Materials Science and Engineering, 2007.
Includes bibliographical references (p. 133-140).
Manipulation of the distribution of phonons inM a solid is important for both basic science and applications ranging from heat management to reduction of noise in electronic circuits and creating materials with superior acoustic and acousto-optical properties. This thesis explores hypersonic phononic crystals as means to achieve control over high frequency acoustic phonons. An integrated approach to fabrication, measurement and analysis of hypersonic phononic crystals with band gaps in the GHz range is presented. First, the phonon dispersion relation for one dimensional polymeric phononic crystals fabricated by coextrusion of a large number of poly(methylmethacrylate)/poly(carbonate) and poly(methylmethacrylate)/poly(ethylene terephthalate) bilayer pairs is investigated as a function of a lattice constant and composition using Brillouin light scattering and numerical simulations. This set of relatively simple multilayer structures represents an excellent platform to gain a basic understanding of phononic band gap phenomena. In addition, their in-plane phonon dispersion is used to extract information about the elastic constants and glass transition temperatures of individual nanolayers in a periodic multilayer arrangement. Next, two dimensional epoxy/air phononic crystals fabricated in a photoresist using interference lithography are studied. These structures are 2D single crystalline, enabling direction-resolved measurements of their phonon dispersion relation. As a result, the complete experimental phononic band diagram is obtained and correlated with numerical simulations. Finally, phononic properties of three dimensional elastomeric poly(dimethylsiloxane) crystals are investigated and the mechanical tunability of their dispersion relation is demonstrated.
(cont.) This set of structures forms the basis for understanding how to design and fabricate acoustic and acousto-optical devices with performance characteristics that can be adjusted dynamically during operation. The investigations described in this thesis demonstrate both theoretically and experimentally that 1D, 2D and 3D periodic submicron structures have complex phonon dispersion relations at GHz frequencies. As a result, these crystals can be used to manipulate the flow of random thermal phonons as well as externally generated acoustic waves resulting in novel acoustic and thermal properties.
by Taras Gorishnyy.
Ph.D.

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4

Haq, Z. U. "Hypersonic vehicle interference heating." Thesis, University of Southampton, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.336171.

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5

Sagerman, Denton Gregory. "Hypersonic Experimental Aero-thermal Capability Study Through Multilevel Fidelity Computational Fluid Dynamics." University of Dayton / OhioLINK, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1499433256220438.

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6

Gibson, Travis Eli. "Adaptive control of hypersonic vehicles." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/46635.

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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Mechanical Engineering, 2008.
Includes bibliographical references (p. 105-109).
The guidance, navigation and control of hypersonic vehicles are highly challenging tasks due to the fact that the dynamics of the airframe, propulsion system and structure are integrated and highly interactive. Such a coupling makes it difficult to model various components with a requisite degree of accuracy. This in turn makes various control tasks including altitude and velocity command tracking in the cruise phase of the flight extremely difficult. This work proposes an adaptive controller for a hypersonic cruise vehicle subject to: aerodynamic uncertainties, center-of-gravity movements, actuator saturation, failures, and time-delays. The adaptive control architecture is based on a linearized model of the underlying rigid body dynamics and explicitly accommodates for all uncertainties. Within the control structure is a baseline Proportional Integral Filter commonly used in optimal control designs. The control design is validated using a highfidelity HSV model that incorporates various effects including coupling between structural modes and aerodynamics, and thrust pitch coupling. Analysis of the Adaptive Robust Controller for Hypersonic Vehicles (ARCH) is carried out using a control verification methodology. This methodology illustrates the resilience of the controller to the uncertainties mentioned above for a set of closed-loop requirements that prevent excessive structural loading, poor tracking performance, and engine stalls. This analysis enables the quantification of the improvements that result from using and adaptive controller for a typical maneuver in the V-h space under cruise conditions.
by Travis Eli Gibson.
S.M.

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7

Starkey, Ryan P., Mark J. Lewis, and Charles H. Jones. "PLASMA TELEMETRY IN HYPERSONIC FLIGHT." International Foundation for Telemetering, 2002. http://hdl.handle.net/10150/607506.

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International Telemetering Conference Proceedings / October 21, 2002 / Town & Country Hotel and Conference Center, San Diego, California
Problems associated with telemetry blackout caused by the plasma sheath surrounding a hypersonic vehicle are addressed. In particular, the critical nature of overcoming this limitation for test and evaluation purposes is detailed. Since the telemetry blackout causes great concern for atmospheric cruise vehicles, ballistic missiles, and reentry vehicles, there have been many proposed approaches to solving the problem. This paper overviews aerodynamic design methodologies, for which the required technologies are only now being realized, which may allow for uninterrupted transmission through a plasma sheath. The severity of the signal attenuation is dependent on vehicle configuration, trajectory, flightpath, and mission.

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8

Ajmani, Kumud. "Turbulence modeling in hypersonic inlets." Thesis, Virginia Polytechnic Institute and State University, 1987. http://hdl.handle.net/10919/101365.

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A study is conducted to analyze the performance of different turbulence models when applied to flow through a Mach 7.4 hypersonic inlet. The analysis, which is two-dimensional, is done by comparing computational results from a Parabolized Navier Stokes code, with experimental data. The McDonald Camarata (MC) and Baldwin Lomax (BL) models were the two zero-equation models used in the study. The Turbulent Kinetic Energy (TKE) model was chosen as a representative higher order model. The MC model, when run with transition of flow, provides a solution which compares excellently with the data. Transition has a first order effect on the overall solution provided by the code. The BL model predicts separation of flow in the inlet, which contradicts experimental findings. The TKE model does not perform any better than the MC and BL models, despite the fact that it is a higher order turbulence model. The BL and TKE models predict transition in the inlet at a location which is much earlier than observed in the experiment. This may be attributed to the empirical constants used to determine the point of transition.
M.S.

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9

Laurence, Stuart Jon Hornung H. G. "Proximal bodies in hypersonic flow /." Diss., Pasadena, Calif. : Caltech, 2006. http://resolver.caltech.edu/CaltechETD:etd-04242006-172719.

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10

Louie, Ken. "Mathematical problems in inviscid hypersonic flow." Thesis, University of Oxford, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.291297.

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11

Bennett, J. S. "Hypersonic flow control using magneto-hydrodynamics." Thesis, Cranfield University, 2008. http://hdl.handle.net/1826/4258.

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The focus of the present work is on the use of magneto hydrodynamics as a flow control device for supersonic and hypersonic vehicles. A three dimensional parabolized Navier Stokes solver was developed to take into account the effects of magnetic fields , by incorporating two magneto-hydrodynamic models. The modified solver was then used to study the effects of magneto-hydrodynamics on a variety of configurations, one study of which involved surrogate model based optimisation procedures. The first component of research involved validation of the low magnetic Reynolds number model model against well documented test cases. Good agreement with the nu- merical test cases for flows past a blunt body and a flat plate boundary layer flow, both in the presence of a magnetic field, was found. A novel application of the method of man- ufactured solutions to the simplified mapeto-hydrodynamic model was made to ensure its Accuracy. Assessment of the procedures used for numerical optimisation, were made against known closed-form solutions, and a theoretical axisymmetric body of revolution. An investigation for an optimal magnetic field configuration, for an over-sped Ram- jet intake was made. It was found that for a suitable choice of magnetic field strength, shock on lip could be achieved. Furthermore, for a suitable choice for the position of the magnetic field source, the design condition can also be satisfied using a weaker mag- netic field. Finally a study examining the use of magnetic fields for flows past a slender body were was performed. Given a suitably orientated dipole source, it was shown that the magnetic field can introduce asymmetries, for an otherwise symmetric flowfield, and thereby introduce side form on the missile.

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12

Lewis, H. O. "Hypersonic free-flight dynamic stability studies." Thesis, University of Southampton, 1996. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.243193.

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13

Rolim, Tiago Cavalcanti. "Experimental analyisis of a hypersonic waverider." Instituto Tecnológico de Aeronáutica, 2009. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=796.

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This work presents the results of shock tunnel tests of a Mach 10 waverider with sharp leading edges. The waverider surface was generated from a conical flowfield with the volume and the viscous lift-to-drag ratio as optimization parameters. A compression and expansion ramps were added to the pure waverider surface in order to simulate the flow over a scramjet engine. The compression ramp was designed so as to provide the ideal conditions for the supersonic combustion of the Hydrogen while the expansion section was derived from an ideal minimum length supersonic nozzle. The experimental data included Schlieren photographs of the flow and the pressure distribution over the compression surface. These data were compared with the inviscid theory. During these investigations, the IEAv's T3 shock tunnel was used to simulate the hypersonic flow. The stagnation conditions as well as the free stream properties were estimated using numerical codes. The tunnel operated at Mach number ranges of 8.9 to 10, Reynolds number from 2.25 x 106 to 8.76 x 106 (m-1) and Knudsen number from 0.06 to 0.19. From the Schlieren photographs it was noted that the inlet flowfield behaves according to the predictions of the hypersonic viscous interaction models. Also, the pressure variation along the compression surface centerline was obtained using piezoelectric pressure sensors. The resulted profile presented the general trend of the flow described by these models.

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14

Roberts, Timothy Peter. "Dynamic effects of hypersonic separated flow." Thesis, University of Southampton, 1989. https://eprints.soton.ac.uk/52258/.

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15

Guarendi, Andrew N. "Numerical Investigations of Magnetohydrodynamic Hypersonic Flows." University of Akron / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=akron1365985409.

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16

Boon, Simon Edward. "Numerical analysis of hypersonic inlet flows." Thesis, Imperial College London, 2008. http://hdl.handle.net/10044/1/8713.

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The research uses CFD to investigate the internal flow of two hypersonic engine inlets: the Hypersonic Research Engine (HRE), a dual-mode ramjet/scramjet, and the Sustained Hypersonic Flight Experiment (SHyFE), a ramjet developed by QinetiQ. Various interactions are considered, namely shock-expansion, shock-shock and shock-boundary layer interactions. To isolate the different interactions, both inviscid and viscous turbulent computations are considered. For the HRE, axisymmetric computations are performed at Mach numbers of 5, 6 and 7, consistent with ground testing conditions used by NASA. The HRE was designed to cruise at a range of Mach numbers; for a given set of freestream flow conditions, dramatically different internal flow characteristics have been found depending on whether the engine arrived at the flow conditions through either acceleration or deceleration. CFD surface data and throat profiles have been compared to, and agree well with, experimental data obtained by NASA. Two flow conditions are investigated for the SHyFE inlet. Firstly, the self-starting characteristics of the SHyFE intake are examined, where the effect of increased internal compression is considered. The findings show undesirable wave interactions, which lead to flow non-uniformities, and decreased shock stabilization properties have adverse effects on the performance of the engine. Secondly, the effect of freestream incidence on the inlet is examined. The SHyFE engine is designed to cruise at a mean incidence of between 2° and 3°, however, it is conceivable that the engine will, at times, operate at 5°. Fully three dimensional computations are performed at an angle of attack of 5° where the resulting flows show that Mach reflections on the inner surface of the cowl can lead to shock-detachment, as well as showing that shock-boundary layer interactions on the centrebody can cause centrebody flow separation which can unstart the engine.

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17

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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18

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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19

Kumar, Sanjeev. "Numerical simulation of chemically reactive hypersonic flows." Aachen Shaker, 2005. http://deposit.d-nb.de/cgi-bin/dokserv?idn=980112028.

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20

Agbormbai, Adolf Akombi. "Gas surface interactions in rarefied hypersonic flows." Thesis, Imperial College London, 1988. http://hdl.handle.net/10044/1/46929.

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21

Wiese, Daniel Philip. "Adaptive control of a generic hypersonic vehicle." Thesis, Massachusetts Institute of Technology, 2013. http://hdl.handle.net/1721.1/81714.

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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Mechanical Engineering, 2013.
Cataloged from PDF version of thesis.
Includes bibliographical references (p. 111-115).
This thesis presents a an adaptive augmented, gain-scheduled baseline LQR-PI controller applied to the Road Runner six-degree-of-freedom generic hypersonic vehicle model. Uncertainty in control effectiveness, longitudinal center of gravity location, and aerodynamic coefficients are introduced in the model, as well as sensor bias and noise, and input time delays. The performance of the baseline controller is compared to the same design augmented with one of two different model-reference adaptive controllers: a classical open-loop reference model design, and modified closed-loop reference model design. Both adaptive controllers show improved command tracking and stability over the baseline controller when subject to these uncertainties. The closed-loop reference model controller offers the best performance, tolerating a reduced control effectiveness of 50%, rearward center of gravity shift of -0.9 to -1.6 feet (6-11% of vehicle length), aerodynamic coefficient uncertainty scaled 4x the nominal value, and sensor bias of +1.6 degrees on sideslip angle measurement. The closed-loop reference model adaptive controller maintains at least 73% of the delay margin provided by the robust baseline design, tolerating input time delays of between 18-46 ms during 3 degree angle of attack doublet, and 80 degree roll step commands.
by Daniel Philip Wiese.
S.M.

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22

Chamitoff, Gregory Errol. "Robust intelligent flight control for hypersonic vehicles." Thesis, Massachusetts Institute of Technology, 1992. http://hdl.handle.net/1721.1/44275.

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23

Ahmed, Mahmoud Y. M. "Aerothermodynamic design optimization of spiked hypersonic vehicles." Thesis, University of Sheffield, 2010. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.531198.

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24

Claus, Malcolm. "Jet interaction effects on a hypersonic interceptor." Thesis, Kingston University, 2001. http://eprints.kingston.ac.uk/20674/.

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A series of experiments were undertaken at the Defence Evaluation Research Agency (DERA) Farnborough, within the Aero Physics group into the phenomenon called Jet Interaction (JI). Jet Interaction (JI) is produced by the interaction of a jet with the external flow around a vehicle. This study focused on investigating the effects of a divert thruster employed to provide a vehicle with a rapid divert capability on the external flow-field and on the induced forces and moments exerted on the vehicle by the jet. The research was based on studying the effects on a hypersonic interceptor sometimes referred to as a KKV (Kinetic Kill Vehicle). There are many important parameters in JI. One of these is the jet Amplification Factor (AF). This is caused by the deflection of the free-stream around and over the jet and results in a pressure increase on the vehicle surface which adds to the divert thrust force. The experiments were carried out in the intermittent hypersonic gun tunnel, at a free-steam Mach number of 12.1. This produced a Reynolds number (based on diameter) Red of 300,000 with a one-tenth scale vehicle; these conditions correspond to a full-scale vehicle flying at an altitude of 41 km. To simulate the divertthruster, nitrogen was supplied to the model through a purpose-made force balance. Measured forces ineluded normal, axial and side as well as pitch and yawing moments. The experimental results have been compared with that of a full size vehicle featuring a 2kN divert thruster. The results have then been matched to the effective altitude as a function of the thrust coefficient (Cr). This allows the experimental data to be interpreted for a full-scale vehicle in order to answer design questions important to system engineers. The results from this investigation show that the effectiveness of a divert jet is influenced by the vehicle's altitude, achieving a negligible increase in AF with a Cr > 2.5 at an altitude > 50 km. The seeker will suffer from jet induced problems at low Cr levels for a. = 10°. An increase in Cr causes the separation region in front of the jet to extend to the nose of the tested configuration for M1 = 12.1 while complete separation is achieved at Cr > 1.2. Injection Mach number (MJ) has a small influence on AF. However it does not influence the separation region. Penetration height (h) of the jet is increased for higher Mach number injection. Both AF and the separation region are influenced by nozzle geometry. A series of different nozzle geometries were tested. These had the effect of reducing the measured amplification factor to a maximum of 1, except for the dual circular orifice combination, which doubled the measured AF achieved for a single circular orifice. The influence of nozzle geometry reduced the Cr levels required to produce a negligible increase in AF and the corresponding altitude. The angle of attack (œ) has a strong influence on AF at low Cr levels, however it becomes negligible when complete separation is achieved.

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25

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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26

Granciu, Veronica Maria <1982&gt. "Diagnostic Techniques for MHD in Hypersonic Flows." Doctoral thesis, Alma Mater Studiorum - Università di Bologna, 2010. http://amsdottorato.unibo.it/2312/1/Granciu_Veronica_Maria_Tesi.pdf.

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27

Granciu, Veronica Maria <1982&gt. "Diagnostic Techniques for MHD in Hypersonic Flows." Doctoral thesis, Alma Mater Studiorum - Università di Bologna, 2010. http://amsdottorato.unibo.it/2312/.

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28

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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29

Bosco, Arianna [Verfasser]. "Reynolds stress model for hypersonic flows / Arianna Bosco." Aachen : Hochschulbibliothek der Rheinisch-Westfälischen Technischen Hochschule Aachen, 2011. http://d-nb.info/1014297168/34.

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30

Dominy, Robert Gerald. "Rarefied hypersonic shock wave and blunt body flows." Thesis, Imperial College London, 1988. http://hdl.handle.net/10044/1/47034.

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31

Williams, Simon. "Three-dimensional separation of a hypersonic boundary layer." Thesis, Imperial College London, 2005. http://hdl.handle.net/10044/1/11450.

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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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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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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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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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36

Yang, Xiaobo. "Numerical study of film cooling in hypersonic flows." Thesis, University of Glasgow, 2002. http://theses.gla.ac.uk/2063/.

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In this thesis, a numerical study of film cooling in hypersonic laminar and turbulent flows has been performed using an in-house Navier-Stokes solver. The aim of this computational work is to investigate the mechanism and effectiveness of film cooling in hypersonic laminar and turbulent flows. Hypersonic flow over a flat plate without film cooling was first studied to provide a reference datum to check the effectiveness of film cooling. For laminar film cooling (M¥ = 9.9), three different primary flow conditions were first used for validation. The inclusion of the development of the flow in the plenum chamber upstream of the slot was found to provide better heat prediction than a uniform boundary condition at the slot exit. Detailed information of the flow field including velocity profile, Mach contour, temperature contour and heat transfer rate was presented. The mechanism of film cooling has been revealed according to the plots of calculated velocity profiles, Mach contours and temperature contours downstream of the slot. The coolant fluid was found to affect the primary boundary layer in two ways: 1) initially a separate layer established by the coolant fluid itself in the near slot area, 2) later a mixing layer between the primary and coolant flow streams. Then five coolant injection rates between 2.95 x 10-4 and 7.33 x 10-4kg/s and three slot heights 0.8382, 1.2192, 1.6002 mm, were examined in hypersonic laminar film cooling. For turbulent film cooling (M¥ = 8.2), for the geometry used in the experiment, the injection at an angle of 20° was found to be appropriate. Different turbulence models including Wilcox's k - w model. Menter's baseline and SST model have been tested. It is concluded that the Wilcox's k - w turbulence model with dilatation-dissipation correction provides the best heat prediction. Again, five coolant injection rates varies from 5.07 x 10-4 to 30.69 x 10-4 kg/s and three slot heights (the same as studied in the laminar film cooling) were studied to check the influence on film cooling effectiveness. Both the coolant and the primary flow were air. Film cooling was found to be an effective way to protect wall surfaces that are exposed under a high heat transfer environment especially in hypersonic laminar flow. Increasing the coolant injection rate can obviously increase the film cooling effectiveness. Again, this works better in laminar flow than in turbulent flow. The coolant injection rate in turbulent flow should be considered to be high enough to give good heat protection. Slot height in both laminar and turbulent flows under the flow conditions in this study was found to be less important, which means other factors can be considered in priority when constructing film cooling systems. With the application of curve fitting, the cooling length was described using power laws according to curve fitting results. A two-equation film coating model has been presented to illustrate the relation between the film cooling effectiveness and the parameter x/(h/m). For film cooling effectiveness in log-log coordinates, a second-order polynomial curve can be used to fit the laminar flows, whilst a straight line is suitable for the turbulent flows.

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37

Kang, Bryan H. (Bryan Heejin). "Air-data estimation for air-breathing hypersonic vehicles." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/47394.

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38

Fico, Vincenzo. "A high-order method for computational hypersonic aerothermodynamics." Thesis, University of Strathclyde, 2011. http://oleg.lib.strath.ac.uk:80/R/?func=dbin-jump-full&object_id=16946.

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39

Starkey, Ryan P., Mark J. Lewis, and Charles H. Jones. "PLASMA SHEATH CHARACTERIZATION FOR TELEMETRY IN HYPERSONIC FLIGHT." International Foundation for Telemetering, 2003. http://hdl.handle.net/10150/606733.

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International Telemetering Conference Proceedings / October 20-23, 2003 / Riviera Hotel and Convention Center, Las Vegas, Nevada
During certain hypersonic flight regimes, shock heating of air creates a plasma sheath resulting in telemetry attenuation or blackout. The severity of the signal attenuation is dependent on vehicle configuration, flight trajectory, and transmission frequency. This phenomenon is investigated with a focus placed on the nonequilibrium plasma sheath properties (electron concentration, plasma frequency, collision frequency, and temperature) for a range of flight conditions and vehicle design considerations. Trajectory and transmission frequency requirements for air-breathing hypersonic vehicle design are then addressed, with comparisons made to both shuttle orbiter and RAM-C II reentry flights.

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40

Vanstone, Leon. "Shock-induced separation of transitional hypersonic boundary layers." Thesis, Imperial College London, 2014. http://hdl.handle.net/10044/1/24803.

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This thesis presents a joint experimental/CFD investigation of shock-induced boundary layer separations in hypersonic transitional boundary layers with an emphasis on collapse and re-establishment times of the separation bubble. This study also provides high fidelity measurements and excellent characterisation of the flow field in order to provide benchmark data of a challenging flow configuration with which to benchmark next generation CFD solvers. The experiments were conducted in the Imperial College Aeronautics Department Number Two Gun Tunnel, a Mach 8.9 axisymmetric facility with a freestream unit Reynolds number of 47 million An axisymmetric blunt-nosed cylinder fitted with an 8 degree flare forms the primary vehicle for this study, although a 1.3 degree cowl geometry was also used to impinge a shock onto the blunt-nosed cylinder. The shock boundary layer interaction was designed such that it was separated for a laminar boundary layer and collapsed for a turbulent one. Carefully controlled turbulent spots were generated upstream of the interaction region which passed through the separation causing its collapse and subsequent re-establishment. Two intermittency cases are considered, one where turbulent spot spacing is large and collapse/re-establishment pairs can be considered independent of each other and one where they can not. Experimental surface quantities through the interaction region are measured using either heat-transfer or pressure measurements and schlieren video is used to diagnose the larger shock structure. Further a non-intrusive toluene PLIF method is assessed for use in this facility and shows promise. CFD simulations are done using an in-house operator split Godunov solver with a Baldwin-Lomax turbulence model. CFD simulations show good agreement with experiment and provides information on flow quantities that would be extremely difficult to measure otherwise. Collapse times of the separation bubble were found to be fast in relation to characteristic spot passage times. The collapse process is also fast in relation to the surrounding flows ability to adjust, with collapse associated with significant shock curvature of the immediate outboard shock structures. This leads to unsteadiness, with surface pressure measurements exceeding the range bounded by the laminar separated and turbulent collapsed cases. The severity of the unsteadiness appears to be driven by turbulent spot spacing. Re-establishment is considerably slower, showing asymptotic recovery that is likely driven by viscous diffusion rates, taking many characteristic spot passage times to recover.

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41

Zanchetta, Marcantony. "Kinetic heating and transition studies at hypersonic speeds." Thesis, Imperial College London, 1996. http://hdl.handle.net/10044/1/37124.

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The thesis reports on an experimental and computational study of kinetic heating at hypersonic speeds. Of particular interest is the transition of the laminar boundary layer to a state of turbulent motion. The experiments are performed in a Mach 9 Gun tunnel with a 5° semi-angle cone geometry. Twelve hemispherically blunted nose radii are tested at three unit Reynolds numbers. Testing has indicated that as the nose is progressively blunted, the transition region moves downstream. Further amounts of bluntness enhance other instability mechanisms and transition events are witnessed in the near nose regions. There are clearly two transitional regimes, denoted the "small bluntness" and "transition reversal" regime, respectively. This study investigates the structure of the transitional boundary layer in both regimes using thin film heat transfer rate gauges and liquid crystal surface thermography. The heat transfer measurements indicate that the small bluntness transition regime is governed by the rapid formation, growth and merging of turbulent events. Transition occurs over hundreds of boundary layer lengths. The reversal regime transition process is characterised by the birth of turbulent events in the nose and near nose regions. The temporal formation rate of the events is governed by roughness. In a low roughness environment, transition occurs over many model lengths. Increasing the roughness level, increases the spot formation rate, and transition is witnessed immediately downstream of the spherical nose region. The role of roughness is further explored using boundary layer trips. The trip causes a laminar wake which rapidly undergoes transition and forms a turbulent wedge. Event circumferential spreading angles are found for a variety of trip geometries and locations. The heat transfer distribution in the wedge is mapped using the thin film gauges. Computational work is used to perform laminar flow field predictions. Of interest is the entropy layer caused by the presence of the bow shock, and its interaction with the boundary layer. Heat transfer predictions in the transitional region are also performed aided with the experimentally obtained intermittency information.

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42

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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43

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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44

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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45

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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46

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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47

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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48

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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49

Kumar, Sanjeev [Verfasser]. "Numerical simulation of chemically reactive hypersonic flows / Sanjeev Kumar." Aachen : Shaker, 2006. http://d-nb.info/980112028/34.

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50

Denman, Paul Ashley. "Experimental study of hypersonic boundary layers and base flows." Thesis, Imperial College London, 1996. http://hdl.handle.net/10044/1/45466.

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This experimental study documents the development and separation of a hypersonic boundary layer produced naturally on the cold surface of a sharp slender cone. At the base of the conical forebody, the equilibrium turbulent boundary layer was allowed to separate over an axisymmetric rearward facing step to form a compressible base flow. The investigation was conducted in the Imperial College No.2 gun tunnel at a freestream Mach number of 9 and unit Reynolds numbers of 15 and 55 million. The compressible boundary layer study was carried out at both of the available freestream unit Reynolds numbers and the measured data include distributions of wall static pressure and heat transfer rate, together with profiles of pitot pressure through the boundary layer. Using the chordwise distribution of surface heat flux as a means of transition detection, the cone transition Reynolds number was found to be 5.4x10^. This result, together with that obtained from flat plate studies conducted in the same test facility, provided a ratio of cone to flat plate transition Reynolds number of 0.8. Boundary layer integral quantities and shape factors are derived from velocity profiles and in most cases the measured data extended close enough to the wall to detect the peak values of the integrands. The separated flow region formed at the base of the cone was documented only at the higher unit Reynolds number, a condition under which the approaching turbulent boundary layer was found to be close to equilibrium. The data include pitot pressure profiles recorded normal to the surface downstream of reattachment, together with wall static pressure and heat transfer rate distributions measured throughout the base flow region. Reattachment occurred approximately two step heights downstream of separation and a surface flow visualisation study indicated the existence of Taylor-Goertler type vortices, emanating from the reattachment line in the downstream direction. A simple shear layer expansion model is developed and shown to provide a favourable prediction of the measured pitot pressure profiles recorded downstream of the reattachment line. The success of this second order model implies that the dynamics of the corner expansion process, except in the immediate vicinity of the wall, is governed largely by inviscid pressure mechanisms and that the supersonic region of the boundary layer expansion is essentially isentropic.

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Dissertations / Theses on the topic 'Hypersonic'

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