Dissertations / Theses: 'Hypersonic Vehicles'

1

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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2

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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3

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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4

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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5

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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6

Kang, Bryan H. (Bryan Heejin). "Air data and surface pressure measurement for hypersonic vehicles." Thesis, Massachusetts Institute of Technology, 1989. http://hdl.handle.net/1721.1/40135.

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7

Fiorentini, Lisa. "Nonlinear Adaptive Controller Design For Air-breathing Hypersonic Vehicles." The Ohio State University, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=osu1274986563.

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8

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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9

Somanath, Amith. "Adaptive control of hypersonic vehicles in presence of actuation uncertainties." Thesis, Massachusetts Institute of Technology, 2010. http://hdl.handle.net/1721.1/59699.

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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2010.
Cataloged from PDF version of thesis.
Includes bibliographical references (p. 73-75).
The thesis develops a new class of adaptive controllers that guarantee global stability in presence of actuation uncertainties. Actuation uncertainties culminate to linear plants with a partially known input matrix B. Currently available multivariable adaptive controllers yield global stability only when the input matrix B is completely known. It is shown in this work that when additional information regarding the structure of B is available, this difficulty can be overcome using the proposed class of controllers. In addition, a nonlinear damping term is added to the adaptive law to further improve the stability characteristics. It is shown here that the adaptive controllers developed above are well suited for command tracking in hypersonic vehicles (HSV) in the presence of aerodynamic and center of gravity (CG) uncertainties. A model that accurately captures the effect of CG shifts on the longitudinal dynamics of the HSV is derived from first principles. Linearization of these nonlinear equations about an operating point indicate that a constant gain controller does not guarantee vehicle stability, thereby motivating the use of an adaptive controller. Performance improvements are shown using simulation studies carried out on a full scale nonlinear model of the HSV. It is shown that the tolerable CG shifts can be almost doubled by using an adaptive controller as compared to a linear controller while tracking reference commands in velocity and altitude.
by Amith Somanath.
S.M.

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10

Vick, Tyler J. "Geometry Modeling and Adaptive Control of Air-Breathing Hypersonic Vehicles." University of Cincinnati / OhioLINK, 2014. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1397468045.

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11

Mannava, Anusha. "Adaptive Control of Nonminimum Phase Aerospace Vehicles- A Case Study on Air-Breathing Hypersonic Vehicle Model." The Ohio State University, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=osu1503265018577074.

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12

Navarro-Martinez, Salvador. "Numerical simulation of laminar flow over hypersonic compression ramps." Thesis, University of Southampton, 2002. https://eprints.soton.ac.uk/47095/.

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13

Sudalagunta, Praneeth Reddy. "Control-oriented Modeling of an Air-breathing Hypersonic Vehicle." Diss., Virginia Tech, 2016. http://hdl.handle.net/10919/72872.

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Design and development of future high speed aircraft require the use of advanced modeling tools early on in the design phase to study and analyze complex aeroelastic, thermoelastic, and aerothermal interactions. This phase, commonly referred to as the conceptual design phase, involves using first principle based analytical models to obtain a practical starting point for the preliminary and detailed design phases. These analytical models are expected to, firstly, capture the effect of complex interactions between various subsystems using basic physics, and secondly, minimize computational costs. The size of a typical air-breathing hypersonic vehicle can vary anywhere between 12 ft, like the NASA X-43A, to 100 ft, like the NASP demonstrator vehicle. On the other hand, the performance expectations can vary anywhere between cruising at Mach 5 @ 85; 000 ft to Mach 10 @ 110; 000 ft. Reduction of computational costs is essential to efficiently sort through such a vast design space, while capturing the various complex interactions between subsystems has shown to improve accuracy of the design estimates. This motivates the need to develop modelling tools using first principle based analytical models with "needed" fidelity, where fidelity refers to the extent of interactions captured. With the advent of multidisciplinary design optimization tools, the need for an integrated modelling and analysis environment for high speed aircraft has increased substantially over the past two decades. The ever growing increase in performance expectations has made the traditional design approach of optimize first, integrate later obsolete. Designing a closed-loop control system for an aircraft might prove to be a difficult task with a geometry that yields an optimal (L/D) ratio, a structure with optimal material properties, and a propulsion system with maximum thrust-weight ratio. With all the subsystems already optimized, there is very little freedom for control designers to achieve their high performance goals. Integrated design methodologies focus on optimizing the overall design, as opposed to individual subsystems. Control-oriented modelling is an approach that involves making appropriate assumptions while modelling various subsystems in order to facilitate the inclusion of control design during the conceptual design phase. Due to their high lift-to-drag ratio and low operational costs, air-breathing hypersonic vehicles have spurred some interest in the field of high speed aircraft design over the last few decades. Modeling aeroelastic effects for such an aircraft is challenging due to its tightly integrated airframe and propulsion system that leads to significant deflections in the thrust vector caused by flexing of the airframe under extreme aerodynamic and thermal loads. These changes in the orientation of the thrust vector in turn introduce low frequency oscillations in the flight path angle, which make control system design a challenging task. Inclusion of such effects in the vehicle dynamics model to develop accurate control laws is an important part of control-oriented modeling. The air-breathing hypersonic vehicle considered here is assumed to be a thin-walled structure, where deformations due to axial, bending, shear, and torsion are modeled using the six independent displacements of a rigid cross section. Free vibration mode shapes are computed accurately using a novel scheme that uses estimates of natural frequency from the Ritz method as initial guesses to solve the governing equations using SUPORE, a two-point boundary value problem solver. A variational approach involving Hamilton's principle of least action is employed to derive the second order nonlinear equations of motion for the flexible aircraft. These nonlinear equations of motion are then linearized about a given cruise condition, modal analysis carried out on the linearized system, and the coupling between various significant modes studied. Further, open-loop stability analysis in time domain is conducted.
Ph. D.

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14

Brewer, Keith Merritt. "Exergy Methods for the Mission-Level Analysis and Optimization of Generic Hypersonic Vehicles." Thesis, Virginia Tech, 2006. http://hdl.handle.net/10919/32007.

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Though the field of hypersonic vehicle design is thriving again, few studies to date demonstrate the technology through a mission in which multiple flight conditions and constraints are encountered. This is likely due to the highly integrated and sensitive nature of hypersonic vehicle components. Consequently, a formal Mach 6 through Mach 10 flight envelope is explored which includes cruise, acceleration/climb, deceleration/descend and turn mission segments. An exergy approach to the vehicle synthesis/design, in which trade-offs between dissimilar technologies are observed, is proposed and measured against traditional methods of assessing highly integrated systems. A quasi one-dimensional hypersonic vehicle system simulation program was constructed. Composed of two sub-systems, propulsion and airframe, mechanisms for loss are computed from such irreversible processes as shocks, friction, heat transfer, mixing, and incomplete combustion. The propulsion sub-system consists of inlet, combustor, and nozzle, while the airframe provides trim and force accounting measures. An energy addition mechanism, based on the potential of MHD technology, is utilized to maintain a shock-on-lip inlet operating condition. Thirteen decision variables (seven design and six operational) were chosen to govern the vehicle geometry and performance. A genetic algorithm was used to evaluate the optimal vehicle synthesis/design for three separate objective functions, i.e the optimizations involved the maximization of thrust efficiency, the minimization of fuel mass consumption, and the minimization of exergy destruction plus fuel exergy loss. The principal results found the minimum fuel consumption and minimum exergy destruction measures equivalent, both meeting the constraints of the mission while using 11% less fuel than the thrust efficiency measure. Optimizing the vehicle for the single most constrained mission segment yielded a vehicle capable of flying the entire mission but with fuel consumption and exergy destruction plus fuel loss values greater than the above mentioned integrated vehicle solutions. In essence, the mission-level analysis provided much insight into the dynamics of mission-level hypersonic flight and demonstrated the usefulness of an exergy destruction minimization measure for highly integrated synthesis/design.
Master of Science

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15

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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16

Finlayson, Simon, and Allan Paull. "A 256 CHANNEL HIGH SPEED MODULAR FLIGHT COMPUTER FOR HYPERSONIC LAUNCH VEHICLES." International Foundation for Telemetering, 2005. http://hdl.handle.net/10150/604873.

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ITC/USA 2005 Conference Proceedings / The Forty-First Annual International Telemetering Conference and Technical Exhibition / October 24-27, 2005 / Riviera Hotel & Convention Center, Las Vegas, Nevada
Hypersonic test vehicles require extensive data acquisition in order to accurately determine and refine engine performance. The increasing speed of scramjet engines places new constraints on data manipulation and system control. A compact modular flight computer has been developed that has high speed analog data acquisition, a programmable high data rate PCM (Pulse Code Modulation) encoder, compact data storage, and high speed I/O (Input/Output) capabilities. Principle to the design is the thermal management required for space environments. A functional overview is presented together with a summary of the analog performance. The integration of future capability requirements is also discussed.

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17

Sigthorsson, David O. "Control-Oriented Modeling and Output Feedback Control of Hypersonic Air-Breathing Vehicles." The Ohio State University, 2008. http://rave.ohiolink.edu/etdc/view?acc_num=osu1228230786.

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18

Alsuwian, Turki Mohammed. "Comparative Analysis of Flight Control Designs for Hypersonic Vehicles at Subsonic Speeds." University of Dayton / OhioLINK, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1543828056218447.

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19

Sharifzadeh, Shayan. "Design Optimization and Analysis of Long-Range Hydrogen-Fuelled Hypersonic Cruise Vehicles." Thesis, The University of Sydney, 2017. http://hdl.handle.net/2123/19127.

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Aviation industry is continuously growing especially for very long distance flights due to the globalisation of local economies around the world and the explosive economic growth in Asia. Reducing the time of intercontinental flights from 16-20 hours to 4 hours or less would therefore make the, already booming, ultra-long distance aviation sector even more attractive. To accomplish this drastic travel time reduction for civil transport, hypersonic cruise aircraft are considered as a potential cost-effective solution. Such vehicles should also be fuelled by liquid hydrogen, which is identified as the only viable propellant to achieve antipodal hypersonic flight with low environmental impact. Despite considerable research on hypersonic aircraft and hydrogen fuel, several major challenges should still be addressed before such airliner becomes reality. The current thesis is therefore motivated by the potential benefit of hydrogen-fuelled hypersonic cruise vehicles associated with their limited state-of-the-art. Hypersonic cruise aircraft require innovative structural configurations and thermal management solutions due to the extremely harsh flight environment, while the uncommon physical properties of liquid hydrogen, combined with high and long-term heat fluxes, introduce complex design and technological storage issues. Achieving hypersonic cruise vehicles is also complicated by the multidisciplinary nature of their design. In the scope of the present research, appropriate methodologies are developed to assess, design and optimize the thermo-structural model and the cryogenic fuel tanks of long-range hydrogen-fuelled hypersonic civil aircraft. Two notional vehicles, cruising at Mach 5 and Mach 8, are then investigated with the implemented methodologies. The design analysis of light yet highly insulated liquid hydrogen tanks for hypersonic cruise vehicles indicates an optimal gravimetric efficiency of 70-75% depending on insulation system, tank wall material, tank diameter, and flight profile. A combination of foam and load-bearing aerogel blanket leads to the lightest cryogenic tank for both the Mach 5 and the Mach 8 aircraft. If the aerogel blanket cannot be strengthened sufficiently so that it can bear the full load, then a combination of foam and fibrous insulation materials gives the best solution for both vehicles. The aero-thermal and structural design analysis of the Mach 5 cruiser shows that the lightest hot-structure is a titanium alloy construction made of honeycomb sandwich panels. This concept leads to a wing-body weight of 143.9 t, of which 36% accounts for the wing, 32% for the fuselage, and 32% for the cryogenic tanks. As expected, hypersonic thermal loads lead to important weight penalties (of more than 35 %). The design of the insulated cold structure, however, demonstrates that the long-term high-speed flight of the airliner requires a substantial thermal protection system, such that the best configuration (obtained by load-bearing aerogel blanket) leads to a titanium cold design of only 4% lighter than the hot structure. Using aluminium 7075 rather than titanium offers a further weight saving of about 2 %, resulting in a 135.4 t wing-body weight (with a contribution of 23 %, 25 %, 18%and 34%from the TPS, the wing, the fuselage, and the cryogenic tanks respectively). Given the design hypotheses, the difference in weight is not significant enough to make a decisive choice between hot and cold concepts. This requires the current methodologies to be further elaborated by relaxing the simplifications. Investigation of the thermal protection must be extended from one single point to different regions of the vehicle, and the TPS thickness and weight should be considered in the structural sizing of the cold design. More generally, the design process should be matured by including additional (static, dynamic and transient) loads, special structural concepts, multi-material configurations and other parameters such as cost and safety aspects.

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20

Sharifzadeh, Shayan. "Design Optimization and Analysis of Long-Range Hydrogen-Fuelled Hypersonic Cruise Vehicles." Doctoral thesis, Universite Libre de Bruxelles, 2017. http://hdl.handle.net/2013/ULB-DIPOT:oai:dipot.ulb.ac.be:2013/255764.

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Aviation industry is continuously growing especially for very long distance flights due to the globalisation of local economies around the world and the explosive economic growth in Asia. Reducing the time of intercontinental flights from 16-20 hours to 4 hours or less would therefore make the, already booming, ultra-long distance aviation sector even more attractive. To accomplish this drastic travel time reduction for civil transport, hypersonic cruise aircraft are considered as a potential cost-effective solution. Such vehicles should also be fuelled by liquid hydrogen, which is identified as the only viable propellant to achieve antipodal hypersonic flight with low environmental impact. Despite considerable research on hypersonic aircraft and hydrogen fuel, several major challenges should still be addressed before such airliner becomes reality. The current thesis is therefore motivated by the potential benefit of hydrogen-fuelled hypersonic cruise vehicles associated with their limited state-of-the-art.Hypersonic cruise aircraft require innovative structural configurations and thermal management solutions due to the extremely harsh flight environment, while the uncommon physical properties of liquid hydrogen, combined with high and long-term heat fluxes, introduce complex design and technological storage issues. Achieving hypersonic cruise vehicles is also complicated by the multidisciplinary nature of their design. In the scope of the present research, appropriate methodologies are developed to assess, design and optimize the thermo-structural model and the cryogenic fuel tanks of long-range hydrogen-fuelled hypersonic civil aircraft. Two notional vehicles, cruising at Mach 5 and Mach 8, are then investigated with the implemented methodologies. The design analysis of light yet highly insulated liquid hydrogen tanks for hypersonic cruise vehicles indicates an optimal gravimetric efficiency of 70-75% depending on insulation system, tank wall material, tank diameter, and flight profile. A combination of foam and load-bearing aerogel blanket leads to the lightest cryogenic tank for both the Mach 5 and the Mach 8 aircraft. If the aerogel blanket cannot be strengthened sufficiently so that it can bear the full load, then a combination of foam and fibrous insulation materials gives the best solution for both vehicles. The aero-thermal and structural design analysis of the Mach 5 cruiser shows that the lightest hot-structure is a titanium alloy construction made of honeycomb sandwich panels. This concept leads to a wing-body weight of 143.9 t, of which 36% accounts for the wing, 32% for the fuselage, and 32% for the cryogenic tanks. As expected, hypersonic thermal loads lead to important weight penalties (of more than 35%). The design of the insulated cold structure, however, demonstrates that the long-term high-speed flight of the airliner requires a substantial thermal protection system, such that the best configuration (obtained by load-bearing aerogel blanket) leads to a titanium cold design of only 4% lighter than the hot structure. Using aluminium 7075 rather than titanium offers a further weight saving of about 2%, resulting in a 135.4 t wing-body weight (with a contribution of 23%, 25%, 18% and 34% from the TPS, the wing, the fuselage, and the cryogenic tanks respectively). Given the design hypotheses, the difference in weight is not significant enough to make a decisive choice between hot and cold concepts. This requires the current methodologies to be further elaborated by relaxing the simplifications. Investigation of the thermal protection must be extended from one single point to different regions of the vehicle, and the TPS thickness and weight should be considered in the structural sizing of the cold design. More generally, the design process should be matured by including additional (static, dynamic and transient) loads, special structural concepts, multi-material configurations and other parameters such as cost and safety aspects.
Doctorat en Sciences de l'ingénieur et technologie
This thesis was conducted in co-tutelle between University of Sydney and Université Libre de Bruxelles.Professor Dries Verstraete was my supervisor at the University of Sydney (so as a member of SydneyUni), but is automatically registered here as a member of ULB because he worked at ULB almost ten years ago.Ben Thornber is also a member of the University of Sydney but the application does not save it for an unknown reason.
info:eu-repo/semantics/nonPublished

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21

Bradford, John Edward. "A technique for rapid prediction of aftbody nozzle performance for hypersonic launch vehicle design." Diss., Georgia Institute of Technology, 2001. http://hdl.handle.net/1853/12896.

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22

Zaludin, Zairil A. "Flight dynamics and automatic flight control system of an hypersonic transport aircraft." Thesis, University of Southampton, 1999. https://eprints.soton.ac.uk/47120/.

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23

Vithana, Sameera J. "The numerical study of 3-dimensional laminar hypersonic blunt-fin interactions." Thesis, University of Southampton, 2007. https://eprints.soton.ac.uk/52003/.

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The three-dimensional numerical simulation of a Mach 6.7 perfect gas, with a unit Reynolds number of 7.6 x 106m-1, over several configurations of a blunt-fin attached to a flat plate are carried out. The resulting interference flowfield is reported in this thesis. The laminar Navier-Stokes code developed by Narvarro-Martinez [47] has been modified to solve any general three-dimensional problem, and the complete Navier-Stokes equations. The numerical scheme is operator split, allowing independent numerical schemes to be used on each of the individual contributions to the Navier-Stokes, which can be combined later to advance the entire solution in time. The inviscid part uses a first order Godunov method with a HLLC approximate Riemann solver; second order accuracy is achieved through the MUSCL approach. The viscous contribution is modeled by a centered difference scheme. An iterative matrix solver is used to advance the implicit solution in time. To handle large three-dimensional grids, the code is implicit and run on a parallel computer cluster. The three-dimensional results from the various blunt-fins simulated show a complex rich three-dimensional structure, with several horseshoe vortices formed within the separated flow. Extremely large heat transfer rates have been measured along the path of these vortices on the plate surface, and on the leading edge of the unswept blunt-fin. In particular cases heat transfer rates as high as (h/hu)60 were measured for the 5mm diameter fin. The 5mm fin results show remarkable similarity to the experimental results obtained by Schuricht [53]. The results obtained using a swept fin, and a fin of doubled fin diameter also show good agreement with the trends observed by Schuricht and others for a laminar interaction.

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24

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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25

Weeks, Carrell Elizabeth. "Evaluation of a Gamma Titanium Aluminide for Hypersonic Structural Applications." Thesis, Georgia Institute of Technology, 2005. http://hdl.handle.net/1853/6955.

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Titanium matrix composites have been extensively evaluated for their potential to replace conventional superalloys in high temperature structural applications, with significant weight-savings while maintaining comparable mechanical properties. The purpose of this investigation is the evaluation of a gamma titanium aluminide alloy with nominal composition Ti-46.5Al-4(Cr,Nb,Ta,B)at.% as a matrix material for use in intermediate temperature applications (400-800㩠in future aerospace transportation systems, as very light-weight structures are needed for cost and weight reduction goals. Mechanical characterization testing was performed over the potential usable temperature range (21-800㩮 Thermal expansion behavior was evaluated, as thermal mismatch of the constituents is an expected problem in composites employing this matrix material. Monotonic testing was conducted on rolled sheet material samples to obtain material properties. The alloy exhibited good strength and stiffness retention at elevated temperatures, as well as improved toughness. Monotonic testing was also conducted on specimens exposed to elevated temperatures to determine the degradation effects of high temperature exposure and oxidation. The exposure did not significantly degrade the alloy properties at elevated temperatures; however, room temperature ductility decreased. Analytical modeling using AGLPLY software was conducted to predict the residual stress state after composite consolidation as well as the potential mechanical behavior of [0]4 laminates with a 㭍ET matrix. Silicon carbide (Ultra-SCS) and alumina (Nextel 610) fibers were selected as potential reinforcing materials for the analysis. High residual stresses were predicted due to the thermal mismatch in the materials. Laminates with Nextel 610 fibers were found to offer the better potential for a composite in this comparison as they provide a better thermal match. Coupons of SCS-6/㭍ET were manufactured with different volume fractions (10% and 20%). Both manufacturing attempts resulted in transverse cracking in the matrix from the residual thermal stress.

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26

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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27

Sockalingam, Subramani. "Coupling of Fluid Thermal Simulation for Nonablating Hypersonic Reentry Vehicles Using Commercial Codes FLUENT and LS-DYNA." University of Cincinnati / OhioLINK, 2008. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1218801526.

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Sockalingam, Subramani. "Coupling of fuid thermal simulation for nonablating hypersonic reentry vehicles using commercial codes FLUENT and LS-DYNA." Cincinnati, Ohio : University of Cincinnati, 2008. http://rave.ohiolink.edu/etdc/view.cgi?acc_num=ucin1218801526.

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Thesis (M.S.)--University of Cincinnati, 2008.
Advisors: Ala Tabiei PhD (Committee Chair), David Thompson PhD (Committee Member), Prem Khosla PhD (Committee Member), Kumar Vemaganti PhD (Committee Member). Title from electronic thesis title page (viewed Sept. 27, 2008). Includes abstract. Keywords: "Hypersonic Reentry Vehicles;Thermal Protection System TPS;Ablation." Includes bibliographical references.

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29

Ordaz, Irian. "A probabilistic and multi-objective conceptual design methodology for the evaluation of thermal management systems on air-breathing hypersonic vehicles." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2008. http://hdl.handle.net/1853/26478.

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Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2009.
Committee Chair: Mavris, Dimitri N.; Committee Member: German, Brian J.; Committee Member: Osburg, Jan; Committee Member: Ruffin, Stephen M.; Committee Member: Schrage, Daniel P.. Part of the SMARTech Electronic Thesis and Dissertation Collection.

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30

Mackall, Dale A., and Robert D. Sakahara. "TECHNICAL CAPABILITIES AND RESOURCES OF THE EXTENDED TEST RANGE ALLIANCE." International Foundation for Telemetering, 1999. http://hdl.handle.net/10150/607322.

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International Telemetering Conference Proceedings / October 25-28, 1999 / Riviera Hotel and Convention Center, Las Vegas, Nevada
The Edwards Flight Test Range is a part of 20,000 square miles of DOD airspace (R-2508). A hypersonic air vehicle traveling above Mach 3 can easily exceed that airspace within seconds. An Unpiloted Autonomous Vehicle can exceed the airspace when flying long duration missions. To satisfy the flight-test requirements of Hypersonic Air Vehicles and Unpiloted Autonomous Vehicles, additional airspace and extended test ranges are required. The Air Force Flight Test Center and Dryden Flight Research Center at Edwards Air Force Base, California have mutual goals to support these flight test programs. To meet these goals, the Extended Test Range Alliance was formed as an engineering and operations team to satisfy program requirements in the areas of telemetry, flight termination, ground communications, uplink command, and differential global positioning systems. This paper will discuss the resources and technical capabilities available through the Extended Test Range.

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Morham, Brett G. "Numerical Examination of Flow Field Characteristics and Fabri Choking of 2D Supersonic Ejectors." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/340.

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An automated computer simulation of the two-dimensional planar Cal Poly Supersonic Ejector test rig is developed. The purpose of the simulation is to identify the operating conditions which produce the saturated, Fabri choke and Fabri block aerodynamic flow patterns. The effect of primary to secondary stagnation pressure ratio on the efficiency of the ejector operation is measured using the entrainment ratio which is the secondary to primary mass flow ratio. The primary flow of the ejector is supersonic and the secondary (entrained) stream enters the ejector at various velocities at or below Mach 1. The primary and secondary streams are both composed of air. The primary plume boundary and properties are solved using the Method of Characteristics. The properties within the secondary stream are found using isentropic relations along with stagnation conditions and the shape of the primary plume. The solutions of the primary and secondary streams iterate on a pressure distribution of the secondary stream until a converged solution is attained. Viscous forces and thermo-chemical reactions are not considered. For the given geometry the saturated flow pattern is found to occur below stagnation pressure ratios of 74. The secondary flow of the ejector becomes blocked by the primary plume above pressure ratios of 230. The Fabri choke case exists between pressure ratios of 74 and 230, achieving optimal operation at the transition from saturated to Fabri choked flow, near the pressure ratio of 74. The case of optimal expansion yields an entrainment ratio of 0.17. The entrainment ratio results of the Cal Poly Supersonic Ejector simulation have an average error of 3.67% relative to experimental data. The accuracy of this inviscid simulation suggests ejector operation in this regime is governed by pressure gradient rather than viscous effects.

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Fojtl, Michal. "Výpočet aerodynamických charakteristik nosiče pro nízkou oběžnou dráhu." Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2017. http://www.nusl.cz/ntk/nusl-316914.

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Master’s thesis deals with aerodynamic heating of launch vehicle during ascent phase by using CFD simulation. Ascent trajectory and payload fairing geometry is design using data of existing small launch vehicles. Critical flight regimes are identified using 2D calculations, and in these regimes analysis is performed by axially symmetric simulations. Simulation results are compared to values obtained from theoretical and semi-empirical calculations.

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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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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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35

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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36

Yu, Weiwei. "Contribution to study and implementation of intelligent adaptive control strategies : application to control of complex dynamic systems." Phd thesis, Université Paris-Est, 2011. http://tel.archives-ouvertes.fr/tel-00665586.

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The main limitation of the CMAC (Cerebellar Model Articulation Controller) network in realistic applications for complex automated systems (robots, automated vehicles, etc...) is related to the required memory size. It is pertinent to remind that the memory used by CMAC depends firstly on the input signal quantification step and secondly on the input space dimension. For real CMAC based control applications, on the one hand, in order to increase the accuracy of the control the chosen quantification step must be as small as possible; on the other hand, generally the input space dimension is greater than two. In order to overcome the problem relating the memory size, how both the generalization and step quantization parameters may influence the CMAC's approximation quality has been discussed. Our goal is to find an optimal CMAC structure for complex dynamic systems' control. Biped robots and Flight control design for airbreathing hypersonic vehicles are two actual areas of such systems. We have applied the investigated concepts on these two quite different areas. The presented simulation results show that an optimal or sub-optimal structure carrying out a minimal modeling error could be achieved. The choice of an optimal structure allows decreasing the memory size and reducing the computing time as well

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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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38

Janicki, William D. (William Daniel). "Asymptotic analysis of hypersonic vehicle dynamics along entry trajectory." Thesis, Massachusetts Institute of Technology, 1991. http://hdl.handle.net/1721.1/42502.

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39

DiGregorio, Nicholas J. "Characteristics of Turbulent Boundary Layers along a Hypersonic Vehicle." Thesis, State University of New York at Buffalo, 2018. http://pqdtopen.proquest.com/#viewpdf?dispub=10822170.

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The flight conditions of a hypersonic vehicle on an ascent trajectory are computed and Reynolds-averaged Navier-Stokes (RANS) simulations of the turbulent boundary layers are performed across a Mach number range of 0.3 up to 16 using the computational fluid dynamics (CFD) software, VULCAN. The boundary conditions and leading edge geometry are varied from the simple case of adiabatic and sharp to cooled and blunted to reveal the physics of how these effects impact the results of flat plate boundary layer methods as applied to practical aerospace systems. The law of the wall, the Van Driest transformation, and a shear stress preserving transformation's ability to collapse boundary layer velocity profiles under the conditions of variable wall boundary condition and leading edge geometry is explored.

Boundary layer related quantities examined include the boundary layer thickness, local skin friction coefficient, displacement thickness, momentum thickness, heat flux, and integrated loads. It is found that cooling the surface serves to increase the density of the boundary layer, making it thinner. This thinning of the boundary layer thickness increases the velocity gradients, thus increasing the shear stresses and the local skin friction coefficient. The effects on turbulent boundary layers of blunting the leading edge are explained by the difference in properties, particularly viscosity, caused by a detached bow shock instead of a Mach wave that comes off of a sharp nose plate. Heat flux into a vehicle is found to be insignificant at low speeds, but increases drastically as the Mach number rises into the supersonic and hypersonic regimes. It is observed that the integrated skin friction coefficient decreases as Mach number increases and the leading edge becomes blunted, however, it increases as more cooling is applied at the boundary. The integrated heat flux computed from a sharp leading edge geometry is greater compared to a blunted leading edge due to greater temperature gradients in the sharp nose case relative to the blunt nose case.

The shear stress preserving transformation, derived with the inclusion of a stress balance condition, is found to produce a better collapse of the velocity profile data than the Van Driest transformation and the incompressible law of the wall regardless of Mach number, boundary condition or leading edge geometry. The normalized untransformed velocity gradients are compared to the velocity gradients resulting from the Van Driest and shear stress preserving tranformation. It is shown that the velocity gradients from the shear stress preserving match the normalized untransformed velocity gradients more closely than the Van Driest velocity gradients do. The advantages, disadvantages, and limitations of each transformation are discussed.

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40

Choi, June. "Application of hypersonic vehicle flying qualities criteria and computational considerations." Thesis, Massachusetts Institute of Technology, 1994. http://hdl.handle.net/1721.1/47356.

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41

Araki, John Jun. "Reentry dynamics and handling qualities of a generic hypersonic vehicle." Thesis, Massachusetts Institute of Technology, 1992. http://hdl.handle.net/1721.1/42532.

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42

Camillo, Giannino Ponchio. "Longitudinal stability analysis and control of an airbreathing hypersonic vehicle." Instituto Tecnológico de Aeronáutica, 2014. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=3154.

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This work presents the open-loop stability analysis and an active control strategy for an airbreathing hypersonic vehicle. The 14-XB, a bidimensional flow airframe derived from the Brazilian 14-X Aerospace Hypersonic Vehicle, is adopted as study platform. In order to perform such analyses, a simulation mathematical model of the airframe longitudinal forces and moments is obtained using perfect gas equations, after considering the relevance of the real gas hypotheses for the expected Mach number range and verifying that the simpler formulation is sufficient. An all-moving horizontal tail is designed in order to enable the aircraft trimming. The horizontal tail design considered simple constraints based on static analysis, and the same gas equations as those used for the airframe study. In order to analyze the aircraft';s dynamic behavior, a Six-Degree-of-Freedom set of equations of motion considering a spherical, rotating Earth is presented in detail, and the necessary conditions to have adequate longitudinal trimming in this scenario are discussed. The open-loop stability of the 14-XB with the designed horizontal tail is assessed through eigenvalue analysis and numerical flight simulations with the horizontal tail fixed at a trim position. Having observed that the aircraft presents unstable long-term natural modes, an active control strategy is suggested in order to stabilize the vehicle and track a desired flight path angle, assuming that thrust is constant and the control surface is an all-moving horizontal tail. The suggested control structure presents pitch stability augmentation system and flight path angle compensator. Optimal gains are calculated using linear quadratic design, along with a gain-scheduling strategy based on simultaneous control design, and the resulting controller presents proper results.

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43

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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44

Shakiba-Herfeh, Mohammad. "Modeling and Nonlinear Control of a 6-DOF Hypersonic Vehicle." The Ohio State University, 2015. http://rave.ohiolink.edu/etdc/view?acc_num=osu1420666327.

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45

Costa, Felipe Jean da. "Thermo-structural analysis of the brazilian 14-x hypersonic aerospace vehicle." Instituto Tecnológico de Aeronáutica, 2014. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=3167.

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The use of computational tools in the aerospace industry has been an important ally in the prediction of physical phenomena. Investment in knowledge in computational simulation is essential for engineering professionals, who work in multidisciplinary design. However, paraphrasing the noble patron of aeronautics, Alberto Santos Dumont, "things are more beautiful when we have a wide vision of your work field", indicating, in the computational simulation context, that the knowledge of methods and theories allocated inside of the software black box allows a complete and lucid analysis of results. Evidently, by commercial reasons we cannot get the knowledge of how to give the total operation of the software, but the core concepts that guarantee the reliability of that can be studied. The Brazilian VHA 14-X is a technological demonstrator of a hypersonic airbreathing propulsion system based on supersonic combustion (scramjet) to fly at Earth';s atmosphere at 30 km altitude at Mach number 7 to 10, designed at the Laboratory of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu, at the Institute for Advanced Studies. Basically, scramjet is a fully integrated airbreathing aeronautical engine that uses the oblique/conical shock waves generated during the hypersonic flight, to promote compression and deceleration of freestream atmospheric air at the inlet of the scramjet. Scramjet is an aeronautical engine, without moving parts, therefore it is necessary another propulsion system to accelerate the scramjet to the operation conditions. The Brazilian two-stage rocket engines (S31 and S30) are able to boost the VHA 14-X to the predetermined conditions of the scramjet operation, 30 km altitude but at Mach number 7 and 10. Therefore, it is needed to design the structure of the VHA 14-X to support the aerodynamic loads during the atmospheric hypersonic flight at the same conditions. One-dimensional theoretical analysis, applied at 30 km altitude at Mach number 7 and 10, provide the pressure distribution on the VHA 14-X upper and lower surfaces. Structural materials for the stringers and ribs as well as coating materials for the thermal protection systems were specified based on preliminary studies. ANSYS Workbench software, which provides the Structural Numerical Analysis using Finite Element Method, has been applied to the structural analysis of the VHA 14-X waverider unpowered scramjet at 30 km altitude at Mach number 7 and 10. Stress field, strains and deformations are presented. Finally, the thermal response in terms of temperature field, total heat flux and directional heat flux in x, y and z axis are provided over the VHA 14-X waverider unpowered scramjet at 30 km altitude at Mach number 7 and 10.

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46

Groves, Kevin. "Modelling, simulation, and control design of an air-breathing hypersonic vehicle." The Ohio State University, 2005. http://rave.ohiolink.edu/etdc/view?acc_num=osu1302726196.

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47

Markell, Kyle Charles. "Exergy Methods for the Generic Analysis and Optimization of Hypersonic Vehicle Concepts." Thesis, Virginia Tech, 2005. http://hdl.handle.net/10919/31256.

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This thesis work presents detailed results of the application of exergy-based methods to highly dynamic, integrated aerospace systems such as hypersonic vehicle concepts. In particular, an exergy-based methodology is compared to a more traditional based measure by applying both to the synthesis/design and operational optimization of a hypersonic vehicle configuration comprised of an airframe sub-system and a propulsion sub-system consisting of inlet, combustor, and nozzle components. A number of key design and operational decision variables are identified as those which govern the hypersonic vehicle flow physics and thermodynamics and detailed one-dimensional models of each component and sub-system are developed. Rates of exergy loss as well as exergy destruction resulting from irreversible loss mechanisms are determined in each of the hypersonic vehicle sub-systems and their respective components. Multiple optimizations are performed for both the propulsion sub-system only and for the entire hypersonic vehicle system for single mission segments and for a partial, three-segment mission. Three different objective functions are utilized in these optimizations with the specific goal of comparing exergy methods to a standard vehicle performance measure, namely, the vehicle overall efficiency. Results of these optimizations show that the exergy method presented here performs well when compared to the standard performance measure and, in a number of cases, leads to more optimal syntheses/designs in terms of the fuel mass flow rate required for a given task (e.g., for a fixed-thrust requirement or a given mission). In addition to the various vehicle design optimizations, operational optimizations are conducted to examine the advantage if any of energy exchange to maintain shock-on-lip for both design and off-design conditions. Parametric studies of the hypersonic vehicle sub-systems and components are also conducted and provide further insights into the impacts that the design and operational decision variables and flow properties have on the rates of exergy destruction.
Master of Science

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48

Culler, Adam John. "Coupled Fluid-Thermal-Structural Modeling and Analysis of Hypersonic Flight Vehicle Structures." The Ohio State University, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=osu1280930589.

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49

Szpak, Benjamin R. "Aerothermoelastic considerations for a control surface on an air-breathing hypersonic vehicle." Connect to resource, 2009. http://hdl.handle.net/1811/37034.

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50

Gunbatar, Yakup. "Nonlinear Adaptive Control and Guidance for Unstart Recovery for a Generic Hypersonic Vehicle." The Ohio State University, 2014. http://rave.ohiolink.edu/etdc/view?acc_num=osu1406160002.

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Dissertations / Theses on the topic 'Hypersonic Vehicles'

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