You searched for subject:(Hypersonic).
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Penn State University
1.
Feldick, Andrew Mark.
Coupled Nonequilibrium Flow, Energy And Radiation Transport For Hypersonic Planetary Entry.
Degree: 2013, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/19764 
► Hypersonic atmospheric entry is an important problem for space exploration. Vehicles which travel through space and enter into planetary atmospheres often travel at extreme speeds,…
(more)
▼ Hypersonic atmospheric entry is an important problem for space exploration. Vehicles which travel through space and enter into planetary atmospheres often travel at extreme speeds, up to approximately 15 km/s. At these speeds, the flow around such blunt bodies has a strong bow shock, where temperatures rise higher than 10,000 K. At these tem- peratures air gas species tend to dissociate, and ionize. Chemical and thermal nonequi- librium may also exist, as the low densities result in fewer particle collisions. In these extreme conditions, radiative heat transfer may become important, both with respect to heat transfer to the vehicle, and to the composition of the surrounding fluid. The descrip- tion of radiative transfer in such flows requires accurate, and computationally efficient models for the solution of the radiative transfer equation (RTE). The present work is aimed at developing a series of methods which can accurately and efficiently solve the RTE, and which can be closely coupled to state-of-the-art
hypersonic computational fluid dynamics solvers, to study the effects of radiation–flow coupling. The methods developed here include a line-by-line (LBL) accurate tangent slab solver, which is the current state-of-the-art. In addition a two dimensional axisymmetric ray tracing photon Monte Carlo solver is adapted to include LBL accurate spectral modeling. This RTE solver has been incorporated within Data-Parallel Line Relaxation (DPLR), a hyper- sonic Computational Fluid Dynamics (CFD) solver, to allow for close radiation–flow coupling. A three dimensional ray tracing Monte Carlo solver has also been developed, to enable radiation modeling in three dimensional flowfields, such as a blunt body flying at angle of attack. A spherical harmonics solver has been included, using LBL accurate spectral modeling to provide an assessment of a common commercially available RTE solver. The accuracy of the tangent slab and spherical harmonic solvers, and the effect of close radiation–flow coupling has been studied in a series of representative flows.
Advisors/Committee Members: Michael F Modest, Dissertation Advisor/Co-Advisor, Michael F Modest, Committee Chair/Co-Chair, Deborah Levin, Committee Member, Daniel Connell Haworth, Committee Member, Gary Stuart Settles, Committee Chair/Co-Chair.
Subjects/Keywords: Radiation; Hypersonic Flow; Hypersonic; RTE; Monte Carlo
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APA (6th Edition):
Feldick, A. M. (2013). Coupled Nonequilibrium Flow, Energy And Radiation Transport For Hypersonic Planetary Entry. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/19764
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Feldick, Andrew Mark. “Coupled Nonequilibrium Flow, Energy And Radiation Transport For Hypersonic Planetary Entry.” 2013. Thesis, Penn State University. Accessed March 01, 2021.
https://submit-etda.libraries.psu.edu/catalog/19764.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Feldick, Andrew Mark. “Coupled Nonequilibrium Flow, Energy And Radiation Transport For Hypersonic Planetary Entry.” 2013. Web. 01 Mar 2021.
Vancouver:
Feldick AM. Coupled Nonequilibrium Flow, Energy And Radiation Transport For Hypersonic Planetary Entry. [Internet] [Thesis]. Penn State University; 2013. [cited 2021 Mar 01].
Available from: https://submit-etda.libraries.psu.edu/catalog/19764.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Feldick AM. Coupled Nonequilibrium Flow, Energy And Radiation Transport For Hypersonic Planetary Entry. [Thesis]. Penn State University; 2013. Available from: https://submit-etda.libraries.psu.edu/catalog/19764
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Virginia Tech
2.
DellaFera, Andrew Brian.
Optimization of Hypersonic Airbreathing Propulsion Systems through Mixed Analysis Methods.
Degree: MS, Aerospace Engineering, 2019, Virginia Tech
URL: http://hdl.handle.net/10919/95512
► Ramjets and scramjets are the key to sustained flight at speeds above five times the speed of sound. These propulsion systems pose a challenging simulation…
(more)
▼ Ramjets and scramjets are the key to sustained flight at speeds above five times the speed
of sound. These propulsion systems pose a challenging simulation environment due to the
wide range of flow seen by the system structure. A scramjet simulation model is formulated
using a series of combustion models with the goal of accurately modelling the combustion
processes throughout the engine. The combustor model is paired with an isolator model and
the engine model is compared against previous studies. A structural analysis model is then
paired with the engine simulation, and the combined model is used within an optimizer to
find an optimum design.
Advisors/Committee Members: Black, Jonathan T. (committeechair), Lowe, Kevin T. (committee member), Cummings, Russell Mark (committee member), Doyle, Daniel Drayson (committee member), Adams, Colin (committee member).
Subjects/Keywords: Hypersonic; Propulsion; Optimization
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APA (6th Edition):
DellaFera, A. B. (2019). Optimization of Hypersonic Airbreathing Propulsion Systems through Mixed Analysis Methods. (Masters Thesis). Virginia Tech. Retrieved from http://hdl.handle.net/10919/95512
Chicago Manual of Style (16th Edition):
DellaFera, Andrew Brian. “Optimization of Hypersonic Airbreathing Propulsion Systems through Mixed Analysis Methods.” 2019. Masters Thesis, Virginia Tech. Accessed March 01, 2021.
http://hdl.handle.net/10919/95512.
MLA Handbook (7th Edition):
DellaFera, Andrew Brian. “Optimization of Hypersonic Airbreathing Propulsion Systems through Mixed Analysis Methods.” 2019. Web. 01 Mar 2021.
Vancouver:
DellaFera AB. Optimization of Hypersonic Airbreathing Propulsion Systems through Mixed Analysis Methods. [Internet] [Masters thesis]. Virginia Tech; 2019. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/10919/95512.
Council of Science Editors:
DellaFera AB. Optimization of Hypersonic Airbreathing Propulsion Systems through Mixed Analysis Methods. [Masters Thesis]. Virginia Tech; 2019. Available from: http://hdl.handle.net/10919/95512
Ryerson University
3.
Miri, Seyed Hossein.
Shock-less Hypersonic Intakes.
Degree: 2012, Ryerson University
URL: https://digital.library.ryerson.ca/islandora/object/RULA%3A1565
► The accuracy of CFD for simulating hypersonic air intake flow is verified by calculating the flow inside a Busemann type intake. The CFD results are…
(more)
▼ The accuracy of CFD for simulating
hypersonic air intake flow is verified by calculating the flow inside a Busemann type intake. The CFD results are then compared against the “exact” solution for the Busemann intake as calculated from the Taylor-McColl equations for conical flow. The method proposed by G. Emanuel (the Lens Analogy) for generating an intake shape that transforms parallel and uniform
hypersonic (freestream) flow isentropically to another parallel and uniform, less
hypersonic, flow has been verified by CFD (SOLVER II) simulation, based on Finite Volume Method (FVM). The shock-less (isentropic) nature of the Lens Analogy (LA) flow shapes has been explored at both on and off-design Mach numbers. The Lens Analogy (LA) method exhibits a limit line (singularity) for low Mach number flows, where the streamlines perform an unrealistic reversal in direction. CFD calculations show no corresponding anomalies.
Advisors/Committee Members: Ryerson University (Degree grantor).
Subjects/Keywords: Aerodynamics; Hypersonic; Aerothermodynamics; Aerodynamics; Hypersonic – Mathematics; Aerodynamics; Hypersonic – Mathematical models
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APA (6th Edition):
Miri, S. H. (2012). Shock-less Hypersonic Intakes. (Thesis). Ryerson University. Retrieved from https://digital.library.ryerson.ca/islandora/object/RULA%3A1565
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Miri, Seyed Hossein. “Shock-less Hypersonic Intakes.” 2012. Thesis, Ryerson University. Accessed March 01, 2021.
https://digital.library.ryerson.ca/islandora/object/RULA%3A1565.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Miri, Seyed Hossein. “Shock-less Hypersonic Intakes.” 2012. Web. 01 Mar 2021.
Vancouver:
Miri SH. Shock-less Hypersonic Intakes. [Internet] [Thesis]. Ryerson University; 2012. [cited 2021 Mar 01].
Available from: https://digital.library.ryerson.ca/islandora/object/RULA%3A1565.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Miri SH. Shock-less Hypersonic Intakes. [Thesis]. Ryerson University; 2012. Available from: https://digital.library.ryerson.ca/islandora/object/RULA%3A1565
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
San Jose State University
4.
Carlozzi, Alexander Anthony.
Estimating Stagnation-Point Radiative Heating Using Maslen's Inverse Method And High-Temperature Equilibrium Air Properties.
Degree: MS, Aerospace Engineering, 2015, San Jose State University
URL: https://doi.org/10.31979/etd.95nk-nuqz
;
https://scholarworks.sjsu.edu/etd_theses/4533
► This thesis attempts to estimate stagnation point radiative heat flux for hypervelocity reentry vehicles. To begin, a shock shape is assumed. Using the inverse…
(more)
▼ This thesis attempts to estimate stagnation point radiative heat flux for hypervelocity reentry vehicles. To begin, a shock shape is assumed. Using the inverse method developed by Maslen, the body that supports such a shock is calculated. High temperature equilibrium theory that accounts for chemical reactions in air was incorporated into Maslen’s Method. Using the calculated air temperature and density behind the bow-shock, radiative heat flux is then calculated at the stagnation point. When a non-grey transparent gas was assumed in the shock-layer, a radiative heat flux at the stagnation point on the vehicle with a nose radius equal to that of the Apollo Command Module was calculated to be 2.9 〖×10〗4 BTU⁄(〖ft〗2 s), using a free-stream velocity of 36,000 feet per second at an altitude of 170,000 feet. This was approximately 18% less than the flux calculated for an infinite slab of comparable gas conditions and shock-layer thickness. When an emitting and absorbing shock-layer were modeled with radiative cooling, agreement with rigorous analysis from previously published data were in agreement on the order of a 1.2 to 13.7% difference.
Subjects/Keywords: Heating; Hypersonic; Radiative; Reentry
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APA (6th Edition):
Carlozzi, A. A. (2015). Estimating Stagnation-Point Radiative Heating Using Maslen's Inverse Method And High-Temperature Equilibrium Air Properties. (Masters Thesis). San Jose State University. Retrieved from https://doi.org/10.31979/etd.95nk-nuqz ; https://scholarworks.sjsu.edu/etd_theses/4533
Chicago Manual of Style (16th Edition):
Carlozzi, Alexander Anthony. “Estimating Stagnation-Point Radiative Heating Using Maslen's Inverse Method And High-Temperature Equilibrium Air Properties.” 2015. Masters Thesis, San Jose State University. Accessed March 01, 2021.
https://doi.org/10.31979/etd.95nk-nuqz ; https://scholarworks.sjsu.edu/etd_theses/4533.
MLA Handbook (7th Edition):
Carlozzi, Alexander Anthony. “Estimating Stagnation-Point Radiative Heating Using Maslen's Inverse Method And High-Temperature Equilibrium Air Properties.” 2015. Web. 01 Mar 2021.
Vancouver:
Carlozzi AA. Estimating Stagnation-Point Radiative Heating Using Maslen's Inverse Method And High-Temperature Equilibrium Air Properties. [Internet] [Masters thesis]. San Jose State University; 2015. [cited 2021 Mar 01].
Available from: https://doi.org/10.31979/etd.95nk-nuqz ; https://scholarworks.sjsu.edu/etd_theses/4533.
Council of Science Editors:
Carlozzi AA. Estimating Stagnation-Point Radiative Heating Using Maslen's Inverse Method And High-Temperature Equilibrium Air Properties. [Masters Thesis]. San Jose State University; 2015. Available from: https://doi.org/10.31979/etd.95nk-nuqz ; https://scholarworks.sjsu.edu/etd_theses/4533
Clemson University
5.
Wilson, Daniel Mason.
Rapid, Low-Fidelity Analysis Methods Using Historical Aircraft Data for Initializing the Conceptual Design of Hypersonic Aircraft.
Degree: MS, Mechanical Engineering, 2020, Clemson University
URL: https://tigerprints.clemson.edu/all_theses/3336
► Hypersonic flight is generally considered greater than Mach 5 and represents an important frontier in aircraft capabilities. Difficulties arise because hypersonic flow differs from…
(more)
▼ Hypersonic flight is generally considered greater than Mach 5 and represents an important frontier in aircraft capabilities. Difficulties arise because
hypersonic flow differs from supersonic and subsonic flow, driving a need for different
hypersonic aircraft configurations, as evidenced by the history of
hypersonic flight. To assist in beginning the conceptual design of
hypersonic aircraft, a process is outlined for
hypersonic design initialization based on historical aircraft data and primarily composed of design data and regression models. It is a rapid, low-fidelity analysis to provide a starting point for the conceptual design process by assessing vehicle type, providing initial estimates for weights and geometry with uncertainty, and exploring changes in estimates with design spaces. Using this framework, an initial set of reasonable aircraft configurations is obtained based on speed, altitude, and payload requirements. The initialization process is outlined for rocket-powered, air-launched
hypersonic vehicles, but additional design data and regression models are provided for other types of
hypersonic vehicles. An example is provided to demonstrate how the outlined process would be used to begin the conceptual design of a new
hypersonic aircraft with a given set of mission requirements.
Advisors/Committee Members: Richard Figliola, Gregory Mocko, Yiqiang Han.
Subjects/Keywords: Design; Hypersonic; Initialization; Regression
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APA (6th Edition):
Wilson, D. M. (2020). Rapid, Low-Fidelity Analysis Methods Using Historical Aircraft Data for Initializing the Conceptual Design of Hypersonic Aircraft. (Masters Thesis). Clemson University. Retrieved from https://tigerprints.clemson.edu/all_theses/3336
Chicago Manual of Style (16th Edition):
Wilson, Daniel Mason. “Rapid, Low-Fidelity Analysis Methods Using Historical Aircraft Data for Initializing the Conceptual Design of Hypersonic Aircraft.” 2020. Masters Thesis, Clemson University. Accessed March 01, 2021.
https://tigerprints.clemson.edu/all_theses/3336.
MLA Handbook (7th Edition):
Wilson, Daniel Mason. “Rapid, Low-Fidelity Analysis Methods Using Historical Aircraft Data for Initializing the Conceptual Design of Hypersonic Aircraft.” 2020. Web. 01 Mar 2021.
Vancouver:
Wilson DM. Rapid, Low-Fidelity Analysis Methods Using Historical Aircraft Data for Initializing the Conceptual Design of Hypersonic Aircraft. [Internet] [Masters thesis]. Clemson University; 2020. [cited 2021 Mar 01].
Available from: https://tigerprints.clemson.edu/all_theses/3336.
Council of Science Editors:
Wilson DM. Rapid, Low-Fidelity Analysis Methods Using Historical Aircraft Data for Initializing the Conceptual Design of Hypersonic Aircraft. [Masters Thesis]. Clemson University; 2020. Available from: https://tigerprints.clemson.edu/all_theses/3336
Delft University of Technology
6.
Viavattene, Giulia (author).
Flying Qualities and Controllability of Hypersonic Spaceplanes.
Degree: 2018, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:c731b2d3-91b5-4617-a844-0a1472abcaf6
► Spaceplanes represent a new promising concept for space flight. A spaceplane is a reusable, safe, efficient and economical space transportation system that can generate lift…
(more)
▼ Spaceplanes represent a new promising concept for space flight. A spaceplane is a reusable, safe, efficient and economical space transportation system that can generate lift during its atmospheric flight, analogously to an aircraft, and it is able to travel in space as a spacecraft. What makes spaceplanes so attractive is the possibility of reusing the system for more than one mission, and the flexibility that they allow in mission. The growing interest in
hypersonic spaceplanes requires that these vehicles have adequate properties of safety of flight and ease of controllability in nominal and off-nominal conditions. From this it follows the need to study their flying qualities and controllability characteristics. The current thesis addresses the flying-quality and controllability analyses, together with the development of a control system capable of improving these properties. The analyses are conducted along the ascent and re-entry trajectory of a representative single-stage-to-orbit spaceplane, the Festip System Study concept 1. The stability, trim capabilities and flying qualities of the open-loop system are analysed by studying the eigenvalues and eigenvectors of the state matrix of the state-space model. Comparing the obtained results with the requirements specified in the military documents MIL-F-8785C and MIL-HDBK-1797 for subsonic vehicles, it is possible to conclude that the reference vehicle is dynamically unstable. Thus, the need for an advanced control system arises. One concept seems particularly interesting for this application: the adaptive control system, which is characterised by a low sensitivity to disturbances thanks to its adaptive gains. Not only the control system is design to be optimal in terms of integrated control error and effort, but also a robust design methodology is applied to identify a control design that is as insensitive as possible to uncertainties of the input and design parameters. The responses for both longitudinal and lateral control in nominal and off-nominal conditions are simulated and evaluated. It results that the system behaviour is strongly related to the control system performance. The robust and advanced control system is able to stabilise the vehicle with relatively low control effort and minimise the effect of disturbances, guaranteeing safety of flight and mission success.
Advisors/Committee Members: Mooij, Erwin (mentor), Delft University of Technology (degree granting institution).
Subjects/Keywords: control; spaceplanes; hypersonic; robust control
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APA (6th Edition):
Viavattene, G. (. (2018). Flying Qualities and Controllability of Hypersonic Spaceplanes. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:c731b2d3-91b5-4617-a844-0a1472abcaf6
Chicago Manual of Style (16th Edition):
Viavattene, Giulia (author). “Flying Qualities and Controllability of Hypersonic Spaceplanes.” 2018. Masters Thesis, Delft University of Technology. Accessed March 01, 2021.
http://resolver.tudelft.nl/uuid:c731b2d3-91b5-4617-a844-0a1472abcaf6.
MLA Handbook (7th Edition):
Viavattene, Giulia (author). “Flying Qualities and Controllability of Hypersonic Spaceplanes.” 2018. Web. 01 Mar 2021.
Vancouver:
Viavattene G(. Flying Qualities and Controllability of Hypersonic Spaceplanes. [Internet] [Masters thesis]. Delft University of Technology; 2018. [cited 2021 Mar 01].
Available from: http://resolver.tudelft.nl/uuid:c731b2d3-91b5-4617-a844-0a1472abcaf6.
Council of Science Editors:
Viavattene G(. Flying Qualities and Controllability of Hypersonic Spaceplanes. [Masters Thesis]. Delft University of Technology; 2018. Available from: http://resolver.tudelft.nl/uuid:c731b2d3-91b5-4617-a844-0a1472abcaf6
University of Notre Dame
7.
Brock E. Hedlund.
Measurement of Flow Perturbations Occurring Both Naturally
and by Plasma Induction over a Mach 4.5 Corner Separation
Zone</h1>.
Degree: Aerospace and Mechanical Engineering, 2018, University of Notre Dame
URL: https://curate.nd.edu/show/d791sf2933j
► This thesis documents the efforts to characterize the Mach 4.5 flow over a flat plate model with a 30° compression ramp over a range…
(more)
▼ This thesis documents the efforts to
characterize the Mach 4.5 flow over a flat plate model with a 30°
compression ramp over a range of unit Reynolds numbers (Rel = 4‧105
– 1‧107 m-1) in low enthalpy (T0 = 293 K) and high enthalpy (T0 =
800 – 1250 K) flow conditions. The compression ramp model was
designed, fabricated, and implemented in a
hypersonic test facility
(ACT-1) to simulate small-scale scramjet inlet operation. Flow
visualization was provided by a schlieren imaging system and planar
laser Rayleigh scattering (PLRS) to characterize the flow structure
with various Rel. Three measurement methods were used to determine
the frequency spectrum of flow perturbations: a high-frequency
Shack-Hartmann wavefront sensor, a laser differential
interferometer, and high-frequency pressure sensors. All three
methods provide complementary results to indicate the existence of
dominant frequencies of perturbations in the flow. It is determined
that of these three measurement tools, Shack-Hartmann wavefront
sensing is the most beneficial for this flow analysis. These
measurements detect modification to the flow perturbation spectra
during plasma actuation and across varied Rel. Damping of the
perturbations in certain frequency bands occurred in the spectra of
measurements taken in the corner separation bubble. In this way,
the spectra indicate dominant frequency bands of perturbations
occurring in regions of the boundary layer and separation bubble.
It is shown, that in the highest tested Rel cases, initial
disturbances in the freestream develop into strong perturbations,
which significantly affect the boundary layer and separation zone
size over the compression ramp. Both amplification and damping
effects from the natural state are seen within different bands of
the spectra of perturbations measured within the flow occurring
over the separation region of the compression ramp collected during
high-frequency plasma actuation (100 kHz). The amplification
observed leads to a fast laminar to turbulent transition
immediately after flow reattachment and potential stabilization of
the shock wave structure. This effect was not observed at F = 50
kHz plasma pulsing frequency.
Advisors/Committee Members: Seong-Kyun Im, Committee Member, Stanislav Gordeyev, Committee Member, Sergey Leonov, Research Director.
Subjects/Keywords: Flow Measurement; Scramjet; Hypersonic
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APA ·
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Export
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Manager
APA (6th Edition):
Hedlund, B. E. (2018). Measurement of Flow Perturbations Occurring Both Naturally
and by Plasma Induction over a Mach 4.5 Corner Separation
Zone</h1>. (Thesis). University of Notre Dame. Retrieved from https://curate.nd.edu/show/d791sf2933j
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Hedlund, Brock E.. “Measurement of Flow Perturbations Occurring Both Naturally
and by Plasma Induction over a Mach 4.5 Corner Separation
Zone</h1>.” 2018. Thesis, University of Notre Dame. Accessed March 01, 2021.
https://curate.nd.edu/show/d791sf2933j.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Hedlund, Brock E.. “Measurement of Flow Perturbations Occurring Both Naturally
and by Plasma Induction over a Mach 4.5 Corner Separation
Zone</h1>.” 2018. Web. 01 Mar 2021.
Vancouver:
Hedlund BE. Measurement of Flow Perturbations Occurring Both Naturally
and by Plasma Induction over a Mach 4.5 Corner Separation
Zone</h1>. [Internet] [Thesis]. University of Notre Dame; 2018. [cited 2021 Mar 01].
Available from: https://curate.nd.edu/show/d791sf2933j.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Hedlund BE. Measurement of Flow Perturbations Occurring Both Naturally
and by Plasma Induction over a Mach 4.5 Corner Separation
Zone</h1>. [Thesis]. University of Notre Dame; 2018. Available from: https://curate.nd.edu/show/d791sf2933j
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of Florida
8.
Nivison, Scott A.
Sparse and Deep Learning-Based Nonlinear Control Design with Hypersonic Flight Applications.
Degree: PhD, Electrical and Computer Engineering, 2017, University of Florida
URL: https://ufdc.ufl.edu/UFE0051837
The task of hypersonic vehicle (HSV) flight control is both intriguing and complicated. HSV control requires dealing with interactions between structural, aerodynamic
Advisors/Committee Members: KHARGONEKAR,PRAMOD P (committee chair), DIXON,WARREN E (committee member), FITZ-COY,NORMAN G (committee member).Subjects/Keywords: adaptive – hypersonic – learning – lyapunov
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APA (6th Edition):
Nivison, S. A. (2017). Sparse and Deep Learning-Based Nonlinear Control Design with Hypersonic Flight Applications. (Doctoral Dissertation). University of Florida. Retrieved from https://ufdc.ufl.edu/UFE0051837
Chicago Manual of Style (16th Edition):
Nivison, Scott A. “Sparse and Deep Learning-Based Nonlinear Control Design with Hypersonic Flight Applications.” 2017. Doctoral Dissertation, University of Florida. Accessed March 01, 2021.
https://ufdc.ufl.edu/UFE0051837.
MLA Handbook (7th Edition):
Nivison, Scott A. “Sparse and Deep Learning-Based Nonlinear Control Design with Hypersonic Flight Applications.” 2017. Web. 01 Mar 2021.
Vancouver:
Nivison SA. Sparse and Deep Learning-Based Nonlinear Control Design with Hypersonic Flight Applications. [Internet] [Doctoral dissertation]. University of Florida; 2017. [cited 2021 Mar 01].
Available from: https://ufdc.ufl.edu/UFE0051837.
Council of Science Editors:
Nivison SA. Sparse and Deep Learning-Based Nonlinear Control Design with Hypersonic Flight Applications. [Doctoral Dissertation]. University of Florida; 2017. Available from: https://ufdc.ufl.edu/UFE0051837
University of Illinois – Urbana-Champaign
9.
Morgan, Jonathan Deming.
Prediction of flight measurements of high enthalpy nonequilibrium flow from a cubesat-class atmospheric probe.
Degree: MS, Aerospace Engineering, 2018, University of Illinois – Urbana-Champaign
URL: http://hdl.handle.net/2142/101076
► Spectral radiation is sensitive to many physical chemical aspects of high-enthalpy, non-equilibrium flows affect radiation and heat transfer estimates at low-earth altitudes, which is important…
(more)
▼ Spectral radiation is sensitive to many physical chemical aspects of high-enthalpy, non-equilibrium flows affect radiation and heat transfer estimates at low-earth altitudes, which is important as a flow diagnostic. For diatomic molecules, high energy collisions in the bow shock region result in ro-vibrionic transitions and release of photons in different wavelength bands. In this work, the Direct Simulations Monte Carlo (DSMC) method is used for the SASSI2 mission to determine the external flow field around a 3U CubeSat. The Nonequilibrium Radiative Transport and Spectra Program (NEQAIR) code is used to determine ultraviolet radiation from nitric oxide and the tangent slab approximation is used to estimate spacecraft visible glow radiance from nitrogen dioxide. Additional calculations are performed to provide a sensitivity analysis of radiance estimates based on the DSMC code utilized, chemical reaction rates, and CubeSat orientation.
Advisors/Committee Members: Levin, Deborah A (advisor).
Subjects/Keywords: hypersonic; nonequilibrium; cubesat; DSMC
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Manager
APA (6th Edition):
Morgan, J. D. (2018). Prediction of flight measurements of high enthalpy nonequilibrium flow from a cubesat-class atmospheric probe. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/101076
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Morgan, Jonathan Deming. “Prediction of flight measurements of high enthalpy nonequilibrium flow from a cubesat-class atmospheric probe.” 2018. Thesis, University of Illinois – Urbana-Champaign. Accessed March 01, 2021.
http://hdl.handle.net/2142/101076.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Morgan, Jonathan Deming. “Prediction of flight measurements of high enthalpy nonequilibrium flow from a cubesat-class atmospheric probe.” 2018. Web. 01 Mar 2021.
Vancouver:
Morgan JD. Prediction of flight measurements of high enthalpy nonequilibrium flow from a cubesat-class atmospheric probe. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2018. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/2142/101076.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Morgan JD. Prediction of flight measurements of high enthalpy nonequilibrium flow from a cubesat-class atmospheric probe. [Thesis]. University of Illinois – Urbana-Champaign; 2018. Available from: http://hdl.handle.net/2142/101076
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of Arizona
10.
Tolle, Frederick Francis, 1921-.
AN INVESTIGATION OF THE INFLUENCE OF A FORWARD EJECTED GAS STREAM IN HYPERSONIC FLOW ABOUT BLUNT-BODIES
.
Degree: 1973, University of Arizona
URL: http://hdl.handle.net/10150/288101
Subjects/Keywords: Aerodynamics; Hypersonic.
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APA (6th Edition):
Tolle, Frederick Francis, 1. (1973). AN INVESTIGATION OF THE INFLUENCE OF A FORWARD EJECTED GAS STREAM IN HYPERSONIC FLOW ABOUT BLUNT-BODIES
. (Doctoral Dissertation). University of Arizona. Retrieved from http://hdl.handle.net/10150/288101
Chicago Manual of Style (16th Edition):
Tolle, Frederick Francis, 1921-. “AN INVESTIGATION OF THE INFLUENCE OF A FORWARD EJECTED GAS STREAM IN HYPERSONIC FLOW ABOUT BLUNT-BODIES
.” 1973. Doctoral Dissertation, University of Arizona. Accessed March 01, 2021.
http://hdl.handle.net/10150/288101.
MLA Handbook (7th Edition):
Tolle, Frederick Francis, 1921-. “AN INVESTIGATION OF THE INFLUENCE OF A FORWARD EJECTED GAS STREAM IN HYPERSONIC FLOW ABOUT BLUNT-BODIES
.” 1973. Web. 01 Mar 2021.
Vancouver:
Tolle, Frederick Francis 1. AN INVESTIGATION OF THE INFLUENCE OF A FORWARD EJECTED GAS STREAM IN HYPERSONIC FLOW ABOUT BLUNT-BODIES
. [Internet] [Doctoral dissertation]. University of Arizona; 1973. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/10150/288101.
Council of Science Editors:
Tolle, Frederick Francis 1. AN INVESTIGATION OF THE INFLUENCE OF A FORWARD EJECTED GAS STREAM IN HYPERSONIC FLOW ABOUT BLUNT-BODIES
. [Doctoral Dissertation]. University of Arizona; 1973. Available from: http://hdl.handle.net/10150/288101
University of Sydney
11.
Sharifzadeh, Shayan.
Design Optimization and Analysis of Long-Range Hydrogen-Fuelled Hypersonic Cruise Vehicles
.
Degree: 2017, University of Sydney
URL: http://hdl.handle.net/2123/19127
► 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…
(more)
▼ 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…
Subjects/Keywords: Hypersonic;
aircraft;
design;
cryogenic;
hydrogen
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APA (6th Edition):
Sharifzadeh, S. (2017). Design Optimization and Analysis of Long-Range Hydrogen-Fuelled Hypersonic Cruise Vehicles
. (Thesis). University of Sydney. Retrieved from http://hdl.handle.net/2123/19127
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Sharifzadeh, Shayan. “Design Optimization and Analysis of Long-Range Hydrogen-Fuelled Hypersonic Cruise Vehicles
.” 2017. Thesis, University of Sydney. Accessed March 01, 2021.
http://hdl.handle.net/2123/19127.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Sharifzadeh, Shayan. “Design Optimization and Analysis of Long-Range Hydrogen-Fuelled Hypersonic Cruise Vehicles
.” 2017. Web. 01 Mar 2021.
Vancouver:
Sharifzadeh S. Design Optimization and Analysis of Long-Range Hydrogen-Fuelled Hypersonic Cruise Vehicles
. [Internet] [Thesis]. University of Sydney; 2017. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/2123/19127.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Sharifzadeh S. Design Optimization and Analysis of Long-Range Hydrogen-Fuelled Hypersonic Cruise Vehicles
. [Thesis]. University of Sydney; 2017. Available from: http://hdl.handle.net/2123/19127
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of New South Wales
12.
Khraibut, Amna.
Laminar Hypersonic Leading Edge Separation.
Degree: Engineering & Information Technology, 2018, University of New South Wales
URL: http://handle.unsw.edu.au/1959.4/59808
;
https://unsworks.unsw.edu.au/fapi/datastream/unsworks:49842/SOURCE01?view=true
► A numerical and analytical investigation of hypersonic leading edge separation has been conducted. This particular configuration was first analyzed by Chapman in the 1950s for…
(more)
▼ A numerical and analytical investigation of
hypersonic leading edge separation has been conducted. This particular configuration was first analyzed by Chapman in the 1950s for supersonic flows. This study will look at this configuration under
hypersonic flow conditions, numerically, analytically, and also with basic surface pressure measurements taken in T-ADFA. Effects of wall temperature and leading edge bluntness have been considered in the study, including perfect and real gas simulations to examine their effects on separation. Slip corrections have also been implemented. A compressible Navier-Stokes solver, US3D, specifically developed for
hypersonic flows has been used to carry out the numerical simulations. The simulations revealed many interesting features caused by wall temperature and bluntness effects. Firstly, while the size of separation has been shown to increase with the increase in wall temperature, wall cooling seems to give rise to secondary vortices within the separated region. An explanation for the existence of these secondary vortices has been presented in terms of gradients, characteristic angle, and the Reynolds number on the dividing streamline. The numerical results have also been analyzed in the context of the triple-deck theory, Chapman's isentropic recompression theory, and analytical work of Oswatitisch and Inger with particular emphasis on the streamline angles and curvatures at separation and reattachment. The most interesting finding has been in characterizing the separation in terms of the triple-deck scaled angle, which delineates stable and unstable secondary vortices. It is shown that a unique relationship exists between this angle and the size of separation and that it is independent of the agency provoking separation. Bluntness studies, on the other hand, showed a very complex interplay between bluntness which delays separation, viscous interactions which promote separation, and the incidence angle. In bluntness studies, map of pressure gradients revealed a complex mechanism in which disturbances can provoke secondary vortices. Analysis of streamline angles and curvatures revealed a very complex separation process on the microscale as a result of separation. Three-dimensional nozzle simulations using the same CFD solver have been carried-out to verify the freestream conditions. These simulations have been validated by performing a pitot pressure survey in T-ADFA. Both laminar and turbulent flow simulations were considered. Comparison with the experimental measurements showed that the laminar flow assumption overpredicts the size of the core flow, while the turbulent boundary layer assumption showed very good agreement with experiments. Limited surface pressure data were also obtained and overall showed good agreement with the numerical results.
Advisors/Committee Members: gai, sudhir, Engineering & Information Technology, UNSW Canberra, UNSW, neely, andrew, Engineering & Information Technology, UNSW Canberra, UNSW.
Subjects/Keywords: flow separation; hypersonic flows; aerothermodynamics
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APA (6th Edition):
Khraibut, A. (2018). Laminar Hypersonic Leading Edge Separation. (Doctoral Dissertation). University of New South Wales. Retrieved from http://handle.unsw.edu.au/1959.4/59808 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:49842/SOURCE01?view=true
Chicago Manual of Style (16th Edition):
Khraibut, Amna. “Laminar Hypersonic Leading Edge Separation.” 2018. Doctoral Dissertation, University of New South Wales. Accessed March 01, 2021.
http://handle.unsw.edu.au/1959.4/59808 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:49842/SOURCE01?view=true.
MLA Handbook (7th Edition):
Khraibut, Amna. “Laminar Hypersonic Leading Edge Separation.” 2018. Web. 01 Mar 2021.
Vancouver:
Khraibut A. Laminar Hypersonic Leading Edge Separation. [Internet] [Doctoral dissertation]. University of New South Wales; 2018. [cited 2021 Mar 01].
Available from: http://handle.unsw.edu.au/1959.4/59808 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:49842/SOURCE01?view=true.
Council of Science Editors:
Khraibut A. Laminar Hypersonic Leading Edge Separation. [Doctoral Dissertation]. University of New South Wales; 2018. Available from: http://handle.unsw.edu.au/1959.4/59808 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:49842/SOURCE01?view=true
University of Texas – Austin
13.
Dannemiller, David Paul.
Periodic cruise of a hypersonic cruiser for fuel minimization.
Degree: MS, Aerospace Engineering, 1981, University of Texas – Austin
URL: http://hdl.handle.net/2152/41765
Subjects/Keywords: Aerodynamics; Hypersonic
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Export
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Manager
APA (6th Edition):
Dannemiller, D. P. (1981). Periodic cruise of a hypersonic cruiser for fuel minimization. (Masters Thesis). University of Texas – Austin. Retrieved from http://hdl.handle.net/2152/41765
Chicago Manual of Style (16th Edition):
Dannemiller, David Paul. “Periodic cruise of a hypersonic cruiser for fuel minimization.” 1981. Masters Thesis, University of Texas – Austin. Accessed March 01, 2021.
http://hdl.handle.net/2152/41765.
MLA Handbook (7th Edition):
Dannemiller, David Paul. “Periodic cruise of a hypersonic cruiser for fuel minimization.” 1981. Web. 01 Mar 2021.
Vancouver:
Dannemiller DP. Periodic cruise of a hypersonic cruiser for fuel minimization. [Internet] [Masters thesis]. University of Texas – Austin; 1981. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/2152/41765.
Council of Science Editors:
Dannemiller DP. Periodic cruise of a hypersonic cruiser for fuel minimization. [Masters Thesis]. University of Texas – Austin; 1981. Available from: http://hdl.handle.net/2152/41765
University of Adelaide
14.
Harrland, Alan.
Hypersonic inlet for a laser powered propulsion system.
Degree: 2012, University of Adelaide
URL: http://hdl.handle.net/2440/79072
► The idea of laser powered lightcraft was first conceptualised in the early 1970's as a means of launching small scale satellite payloads into orbit at…
(more)
▼ The idea of laser powered lightcraft was first conceptualised in the early 1970's as a means of launching small scale satellite payloads into orbit at a much lower cost in comparison to conventional techniques. Propulsion in the lightcraft is produced via laser induced detonation of the incoming air stream, which results in the energy source for propulsion being decoupled from the vehicle. In air breathing mode the lightcraft carries no onboard fuel or oxidiser, allowing theoretically infinite specific impulses to be achieved. Recently interest has been renewed in this innovative
technology through cross-continent and industry research programs aimed at making laser propulsion a reality. In a ground launched satellite, the vehicle must travel through the atmosphere at speeds greatly in excess of the speed of sound in order to achieve the required orbital velocities. Supersonic, and in particular
hypersonic, flight regimes exhibit complicated physics that render traditional subsonic inlet design techniques inadequate. The laser induced detonation propulsion system requires a suitable engine configuration that offers good performance over all flight speeds and angles of attack to ensure the required thrust is maintained throughout the mission. Currently a
hypersonic inlet has not been developed for the laser powered lightcraft vehicle. Stream traced
hypersonic inlets have demonstrated the required performance in conventional hydrocarbon fuelled scramjet engines. This design technique is applied to the laser powered lightcraft vehicle, with its performance evaluated against the traditional lightcraft inlet design. Four different
hypersonic lightcraft inlets have been produced employing both the stream traced inlet design methodology, and traditional axi-symmetric inlet techniques. This thesis outlines the inlet design methodologies employed, with a detailed analysis of the performance of the lightcraft inlet at angles of attack and off-design conditions. Fully three-dimensional turbulent computational fluid dynamics simulations have been performed on a variety of inlet configurations. The performance of the lightcraft inlets have been evaluated at differing angles of attack. An idealised laser detonation simulation has also been performed to verify that the lightcraft inlet does not unstart during the laser powered propulsion cycle.
Advisors/Committee Members: Doolan, Cornelius Joseph (advisor), School of Mathematical Sciences (school).
Subjects/Keywords: hypersonic; stream traced inlet; lightcraft; hypersonic inlet design; computational fluid dynamics
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Manager
APA (6th Edition):
Harrland, A. (2012). Hypersonic inlet for a laser powered propulsion system. (Thesis). University of Adelaide. Retrieved from http://hdl.handle.net/2440/79072
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Harrland, Alan. “Hypersonic inlet for a laser powered propulsion system.” 2012. Thesis, University of Adelaide. Accessed March 01, 2021.
http://hdl.handle.net/2440/79072.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Harrland, Alan. “Hypersonic inlet for a laser powered propulsion system.” 2012. Web. 01 Mar 2021.
Vancouver:
Harrland A. Hypersonic inlet for a laser powered propulsion system. [Internet] [Thesis]. University of Adelaide; 2012. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/2440/79072.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Harrland A. Hypersonic inlet for a laser powered propulsion system. [Thesis]. University of Adelaide; 2012. Available from: http://hdl.handle.net/2440/79072
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Indian Institute of Science
15.
Nagashetty, K.
Experimental Investigations on Hypersonic Waverider.
Degree: MSc Engg, Faculty of Engineering, 2018, Indian Institute of Science
URL: http://etd.iisc.ac.in/handle/2005/3195
► In the flying field of space transportation domain, the increased efforts involving design and development of hypersonic flight for space missions is on toe to…
(more)
▼ In the flying field of space transportation domain, the increased efforts involving design and development of
hypersonic flight for space missions is on toe to provide the optimum aerothermodynamic design data to satisfy mission requirements. Aerothermodynamics is the basis for designing and development of
hypersonic space transportation flight vehicles such as X 51 a, and other programmes like planetary probes for Moon and Mars, and Earth re-entry vehicles such as SRE and space shuttle. It enables safe flying of aerospace vehicles, keeping other parameters optimum for structural and materials with thermal protection systems. In this context, the experimental investigations on
hypersonic waverider are carried out at design Mach 6.
The
hypersonic waverider has high lift to drag ratio at design Mach number even at zero degree angle of incidence, and this seems to be one of the special characteristics for its shape at
hypersonic flight regime. The heat transfer rates are measured using 30 thin film platinum gauges sputtered on a Macor material that are embedded on the test model. The waverider has 16 sensors on top surface and 14 on bottom surface of a model. The surface temperature history is directly converted to heat transfer rates. The heat transfer data are measured for design (Mach 6) and off-design Mach numbers (8) in the
hypersonic shock tunnel, HST2. The results are obtained at stagnation enthalpy of ~ 2 MJ/kg, and Reynolds number range from 0.578 x 106 m-1 to 1.461 x 106 m-1. In addition, flow visualization is carried out by using Schlieren technique to obtain the shock structures and flow evolution around the Waverider. Some preliminary computational analyses are conducted using FLUENT 6.3 and HiFUN, which gave quantitative results. Experimentally measured surface heat flux data are compared with the computed one and both the data agree well. These detailed results are presented in the thesis.
Advisors/Committee Members: Reddy, K P J (advisor).
Subjects/Keywords: Hypersonic Waverider; Aerothermodynamics; Hypersonic Aerodynamics; Aerodynamics; Hypersonic Shock Tunnel - Heat Transfer; Hypersonic Shock Tunnel HST2; Flow Visualization; Shock Wave Loading; Hypersonic Mach Numbers; Aerospace Engineering
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APA (6th Edition):
Nagashetty, K. (2018). Experimental Investigations on Hypersonic Waverider. (Masters Thesis). Indian Institute of Science. Retrieved from http://etd.iisc.ac.in/handle/2005/3195
Chicago Manual of Style (16th Edition):
Nagashetty, K. “Experimental Investigations on Hypersonic Waverider.” 2018. Masters Thesis, Indian Institute of Science. Accessed March 01, 2021.
http://etd.iisc.ac.in/handle/2005/3195.
MLA Handbook (7th Edition):
Nagashetty, K. “Experimental Investigations on Hypersonic Waverider.” 2018. Web. 01 Mar 2021.
Vancouver:
Nagashetty K. Experimental Investigations on Hypersonic Waverider. [Internet] [Masters thesis]. Indian Institute of Science; 2018. [cited 2021 Mar 01].
Available from: http://etd.iisc.ac.in/handle/2005/3195.
Council of Science Editors:
Nagashetty K. Experimental Investigations on Hypersonic Waverider. [Masters Thesis]. Indian Institute of Science; 2018. Available from: http://etd.iisc.ac.in/handle/2005/3195
Indian Institute of Science
16.
Philip, Sarah Jobin.
Investigation of Heat Transfer Rates Around the Aerodynamic Cavities on a Flat Plate at Hypersonic Mach Numbers.
Degree: MSc Engg, Faculty of Engineering, 2018, Indian Institute of Science
URL: http://etd.iisc.ac.in/handle/2005/3900
► Aerodynamic cavities are common features on hypersonic vehicles which are caused in both large and small scale features like surface defects, pitting, gap in joints…
(more)
▼ Aerodynamic cavities are common features on
hypersonic vehicles which are caused in both large and small scale features like surface defects, pitting, gap in joints etc. In the
hypersonic regime, the presence of such cavities alters the flow phenomenon considerably and heating rates adjacent to the discontinuities can be greatly enhanced due to the diversion of flow. Since the 1960s, a great deal of theoretical and experimental research has been carried out on cavity flow physics and heating. However, most of the studies have been done to characterize the effect downstream and within the cavity. In the present study, a series of were carried out in the shock tunnel to investigate the heating characteristics, upstream and on the lateral side of the cavity. Heat flux measurement has been done using indigenously developed high resistance platinum thin film gauges. High resistance gauges, as contrary to the conventionally used low resistance gauges were showing good response to the extremely low heat flux values on a flat plate with sharp leading edge. The experimental measurements of heat done on a flat plate with sharp leading edge using these gauges show good match with theoretical relation by Crabtree et al. Flow visualization using high speed camera with the cavity model and shock structures visualized were similar to reported in supersonic cavity flow. This also goes to state that in spite of the fluctuating shear layer-the main feature of
hypersonic flow over a cavity ,reasonable studies can be done within the short test time of shock tunnel.
Numerical Simulations by solving the Navier-Stokes equation, using the commercially available CFD package FLUENT 13.0.0 has been done to complement the experimental studies.
Advisors/Committee Members: Reddy, K P J (advisor).
Subjects/Keywords: Aerodynamics; Hypersonic Cavity Flow Physics; Hypersonic Shock Tunnel HST2; Convective Heat Transfer; Hypersonic Flow - Numerical Simulation; Flow Visualization; Hypersonic Mach Numbers; Hypersonic Flow; Hypersonic Shock Tunnels - Aerodynamic Cavities; Aerodynmaic Cavities - Heat Transfer; Hypersonic Shock Tunnels - Heat Transfer; Flat Plates; Uncertainty Analysis; Aerospace Engineering
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APA (6th Edition):
Philip, S. J. (2018). Investigation of Heat Transfer Rates Around the Aerodynamic Cavities on a Flat Plate at Hypersonic Mach Numbers. (Masters Thesis). Indian Institute of Science. Retrieved from http://etd.iisc.ac.in/handle/2005/3900
Chicago Manual of Style (16th Edition):
Philip, Sarah Jobin. “Investigation of Heat Transfer Rates Around the Aerodynamic Cavities on a Flat Plate at Hypersonic Mach Numbers.” 2018. Masters Thesis, Indian Institute of Science. Accessed March 01, 2021.
http://etd.iisc.ac.in/handle/2005/3900.
MLA Handbook (7th Edition):
Philip, Sarah Jobin. “Investigation of Heat Transfer Rates Around the Aerodynamic Cavities on a Flat Plate at Hypersonic Mach Numbers.” 2018. Web. 01 Mar 2021.
Vancouver:
Philip SJ. Investigation of Heat Transfer Rates Around the Aerodynamic Cavities on a Flat Plate at Hypersonic Mach Numbers. [Internet] [Masters thesis]. Indian Institute of Science; 2018. [cited 2021 Mar 01].
Available from: http://etd.iisc.ac.in/handle/2005/3900.
Council of Science Editors:
Philip SJ. Investigation of Heat Transfer Rates Around the Aerodynamic Cavities on a Flat Plate at Hypersonic Mach Numbers. [Masters Thesis]. Indian Institute of Science; 2018. Available from: http://etd.iisc.ac.in/handle/2005/3900
Texas A&M University
17.
Brown II, Robert Lee.
Numerical Simulations of Hypersonic Aerothermoelastic Phenomena.
Degree: PhD, Aerospace Engineering, 2016, Texas A&M University
URL: http://hdl.handle.net/1969.1/158046
► This dissertation examines the following subjects important to hypersonic aerothermoelastic flows: aerodynamic heating on a double-wedge airfoil, active cooling on a double-wedge airfoil, and aerothermoelastic…
(more)
▼ This dissertation examines the following subjects important to
hypersonic aerothermoelastic flows: aerodynamic heating on a double-wedge airfoil, active cooling on a double-wedge airfoil, and aerothermoelastic panel utter with deformable supports. To facilitate this examination, an aerothermoelastic solver was created by coupling solvers for the structural elasticity and thermal diffusion equations to a Reynolds averaged Navier-Stokes (RANS) solver. Validation studies were performed on the aerodynamic, aerothermal, and aerothermoelastic configurations of the solver, which were then used to investigate the aforementioned subjects.
The aerodynamic solver was validated for
hypersonic heating simulations by comparison to the Fay-Riddell equation for the peak heat flux on a circular cylinder and by comparison to compressible boundary layer ow for heat flux into a at plate. A novel investigation was then performed of the heat flux on a double-wedge airfoil, looking at variations of heat flux due to angle of attack, wall temperature, Mach number, and altitude.
The aerothermal solver was validated using experimental and computational data from the
hypersonic heating of a spherical protuberance on a flat plate. The solver was then used in a novel study to analyze the effect of active cooling on the steady-state skin temperature of a double-wedge airfoil in
hypersonic flow. Active cooling using a piecewise continuous cooling distribution resulted in sufficient temperature reduction, but also results in significant chordwise temperature gradients.
The aerothermoelastic solver was validated using computational data from the analysis of panel utter. The solver was then used to examine the effects of deformable structural supports on the utter dynamics of the panel for the first time. Deformable supports along existing simply-supported boundaries were shown to be ineffective at improving resistance to utter or buckling. Deformable supports within a panel were shown to increase the utter resistance threefold and buckling resistance almost fivefold.
Advisors/Committee Members: Cizmas, Paul GA (advisor), Whitcomb, John D (committee member), Reddy, Junuthula N (committee member), Strganac, Thomas A (committee member).
Subjects/Keywords: hypersonic; panel flutter; aerothermal; aeroelastic; aerothermoelastic; simulation
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APA (6th Edition):
Brown II, R. L. (2016). Numerical Simulations of Hypersonic Aerothermoelastic Phenomena. (Doctoral Dissertation). Texas A&M University. Retrieved from http://hdl.handle.net/1969.1/158046
Chicago Manual of Style (16th Edition):
Brown II, Robert Lee. “Numerical Simulations of Hypersonic Aerothermoelastic Phenomena.” 2016. Doctoral Dissertation, Texas A&M University. Accessed March 01, 2021.
http://hdl.handle.net/1969.1/158046.
MLA Handbook (7th Edition):
Brown II, Robert Lee. “Numerical Simulations of Hypersonic Aerothermoelastic Phenomena.” 2016. Web. 01 Mar 2021.
Vancouver:
Brown II RL. Numerical Simulations of Hypersonic Aerothermoelastic Phenomena. [Internet] [Doctoral dissertation]. Texas A&M University; 2016. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/1969.1/158046.
Council of Science Editors:
Brown II RL. Numerical Simulations of Hypersonic Aerothermoelastic Phenomena. [Doctoral Dissertation]. Texas A&M University; 2016. Available from: http://hdl.handle.net/1969.1/158046
Penn State University
18.
Patil, Varun Nandan.
Study of Shock-Shock Interactions for the HET Facility Double Wedge Configuration using the DSMC Approach.
Degree: 2013, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/19072
► A set of numerical simulations designed to study the laminar, shock-shock interactions from hypersonic flows about a double wedge configuration for the Hypervelocity Expansion Tube…
(more)
▼ A set of numerical simulations designed to study the laminar, shock-shock interactions from
hypersonic flows about a double wedge configuration for the Hypervelocity Expansion Tube (HET) facility are presented.
Computations are made using the Direct Simulation Monte Carlo (DSMC) method, an approach for modeling finite-Knudsen number flows.
The current study focuses on the investigation of Mach 7 nitrogen flows about a 30-/55-deg double wedge model for stagnation enthalpies varying from 2.0-8.0 MJ/kg.
The simulation results of the double wedge flows are compared with the data obtained from experiments at HET.
Numerical Schlierens are generated to visualize the shock structure and shock-shock interactions present in these flows and are compared with the experimental Schlieren images.
The computed heat transfer values from the simulations match the experiment along the first surface, but on the second wedge the computed heat transfer distribution overpredicts the measured peak values.
The influence of different models for nonequilibrium nitrogen dissociation, rotational and vibrational relaxation rates, and gas-surface interactions on the shock interaction region are analyzed for high enthalpy flow features and heat transfer rates.
Overall good agreement is observed in the experimental and computational results. Unsteadiness of the flow and time-averaging of the experimental measurements are likely reasons for the inability of the DSMC simulations to exactly reproduce the experimental data.
Advisors/Committee Members: Dr Deborah Ann Levin, Thesis Advisor/Co-Advisor.
Subjects/Keywords: Hypersonic Flows; Shock-shock interactions; DSMC
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APA (6th Edition):
Patil, V. N. (2013). Study of Shock-Shock Interactions for the HET Facility Double Wedge Configuration using the DSMC Approach. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/19072
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Patil, Varun Nandan. “Study of Shock-Shock Interactions for the HET Facility Double Wedge Configuration using the DSMC Approach.” 2013. Thesis, Penn State University. Accessed March 01, 2021.
https://submit-etda.libraries.psu.edu/catalog/19072.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Patil, Varun Nandan. “Study of Shock-Shock Interactions for the HET Facility Double Wedge Configuration using the DSMC Approach.” 2013. Web. 01 Mar 2021.
Vancouver:
Patil VN. Study of Shock-Shock Interactions for the HET Facility Double Wedge Configuration using the DSMC Approach. [Internet] [Thesis]. Penn State University; 2013. [cited 2021 Mar 01].
Available from: https://submit-etda.libraries.psu.edu/catalog/19072.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Patil VN. Study of Shock-Shock Interactions for the HET Facility Double Wedge Configuration using the DSMC Approach. [Thesis]. Penn State University; 2013. Available from: https://submit-etda.libraries.psu.edu/catalog/19072
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Penn State University
19.
Sohn, Ilyoup.
Modeling and Simulation of Radiation from Hypersonic Flows using Monte Carlo Methods.
Degree: 2011, Penn State University
URL: https://submit-etda.libraries.psu.edu/catalog/11737
► During extreme-Mach number reentry into Earth's atmosphere, spacecraft experience hypersonic non-equilibrium flow conditions that dissociate molecules and ionize atoms. Such situations occur behind a shock…
(more)
▼ During extreme-Mach number reentry into Earth's atmosphere, spacecraft experience
hypersonic
non-equilibrium flow conditions that dissociate molecules and ionize atoms.
Such situations occur behind a shock wave leading to high temperatures, which have an adverse effect
on the thermal protection system and radar communications.
Since the electronic energy levels of gaseous species are strongly excited for high Mach number
conditions, the radiative contribution to the total heat load can be significant.
In addition, radiative heat source within the shock layer may affect the internal energy distribution of dissociated
and weakly ionized gas species and the number density of ablative species released from the surface of vehicles.
Due to the radiation total heat load to the heat shield surface of the vehicle may be altered beyond mission tolerances.
Therefore, in the design process of spacecrafts
the effect of radiation must be considered and radiation analyses coupled with flow solvers
have to be implemented to improve the reliability during the vehicle design stage.
To perform the first stage for radiation analyses coupled with gas-dynamics,
efficient databasing schemes for emission and absorption coefficients were
developed to
model radiation from
hypersonic, non-equilibrium flows.
For bound-bound transitions, spectral information including the line-center
wavelength and assembled parameters for efficient calculations of emission and absorption coefficients
are stored for typical air plasma species. Since the flow is non-equilibrium, a rate
equation approach including both collisional and
radiatively induced transitions was used to calculate
the electronic state populations,
assuming quasi-steady-state (QSS). The Voigt line shape function was
assumed for modeling the line broadening effect.
The accuracy and efficiency of the databasing scheme
was examined by comparing results of the databasing scheme with those of NEQAIR
for the Stardust flowfield.
An accuracy of approximately 1 % was achieved with an efficiency
about three times faster than the NEQAIR code.
To perform accurate and efficient analyses of chemically reacting flowfield - radiation interactions,
the direct simulation Monte Carlo (DSMC) and the photon Monte Carlo (PMC) radiative transport methods are used to
simulate flowfield - radiation coupling from transitional to peak heating freestream conditions.
The non-catalytic and fully catalytic surface conditions were modeled and good agreement of the stagnation-point
convective heating between DSMC and continuum fluid dynamics (CFD) calculation under the assumption of fully catalytic surface
was achieved. Stagnation-point radiative heating, however, was found to be very different.
To simulate three-dimensional radiative transport, the finite-volume based PMC (FV-PMC) method was employed.
DSMC - FV-PMC simulations with the goal of understanding
the effect of radiation on the flow structure for different degrees of
hypersonic non-equilibrium are presented.
It is found that except for the highest altitudes,…
Advisors/Committee Members: Deborah A Levin, Dissertation Advisor/Co-Advisor, Deborah A Levin, Committee Chair/Co-Chair, Cengiz Camci, Committee Member, Michael Matthew Micci, Committee Member, Michael F Modest, Committee Member.
Subjects/Keywords: Radiation; Hypersonic; Monte Carlo Methods; Nonequilibrium
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APA (6th Edition):
Sohn, I. (2011). Modeling and Simulation of Radiation from Hypersonic Flows using Monte Carlo Methods. (Thesis). Penn State University. Retrieved from https://submit-etda.libraries.psu.edu/catalog/11737
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Sohn, Ilyoup. “Modeling and Simulation of Radiation from Hypersonic Flows using Monte Carlo Methods.” 2011. Thesis, Penn State University. Accessed March 01, 2021.
https://submit-etda.libraries.psu.edu/catalog/11737.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Sohn, Ilyoup. “Modeling and Simulation of Radiation from Hypersonic Flows using Monte Carlo Methods.” 2011. Web. 01 Mar 2021.
Vancouver:
Sohn I. Modeling and Simulation of Radiation from Hypersonic Flows using Monte Carlo Methods. [Internet] [Thesis]. Penn State University; 2011. [cited 2021 Mar 01].
Available from: https://submit-etda.libraries.psu.edu/catalog/11737.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Sohn I. Modeling and Simulation of Radiation from Hypersonic Flows using Monte Carlo Methods. [Thesis]. Penn State University; 2011. Available from: https://submit-etda.libraries.psu.edu/catalog/11737
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Purdue University
20.
Sparapany, Michael.
Towards the Real-Time Application of Indirect Methods for Hypersonic Missions.
Degree: MSin Aeronautics and Astronautics, Aeronautics and Astronautics, 2015, Purdue University
URL: https://docs.lib.purdue.edu/open_access_theses/1161
► Conceptual hypersonic mission design has typically been performed in a computationally intensive, iterative manner using direct optimization methods. The introduction of modern computing has resulted…
(more)
▼ Conceptual
hypersonic mission design has typically been performed in a computationally intensive, iterative manner using direct optimization methods. The introduction of modern computing has resulted in the widespread adoption of direct methods, and useful information associated with optimal solutions has been lost. Optimization through indirect methods leverages this information, yielding high quality trajectories while reducing the dimensionality of the overall problem.
Advisors/Committee Members: Michael J Grant, James M Longuski, William A Crossley.
Subjects/Keywords: Hypersonic; Indirect; Methods; Optimization; Real; Time
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APA (6th Edition):
Sparapany, M. (2015). Towards the Real-Time Application of Indirect Methods for Hypersonic Missions. (Thesis). Purdue University. Retrieved from https://docs.lib.purdue.edu/open_access_theses/1161
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Chicago Manual of Style (16th Edition):
Sparapany, Michael. “Towards the Real-Time Application of Indirect Methods for Hypersonic Missions.” 2015. Thesis, Purdue University. Accessed March 01, 2021.
https://docs.lib.purdue.edu/open_access_theses/1161.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Sparapany, Michael. “Towards the Real-Time Application of Indirect Methods for Hypersonic Missions.” 2015. Web. 01 Mar 2021.
Vancouver:
Sparapany M. Towards the Real-Time Application of Indirect Methods for Hypersonic Missions. [Internet] [Thesis]. Purdue University; 2015. [cited 2021 Mar 01].
Available from: https://docs.lib.purdue.edu/open_access_theses/1161.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Sparapany M. Towards the Real-Time Application of Indirect Methods for Hypersonic Missions. [Thesis]. Purdue University; 2015. Available from: https://docs.lib.purdue.edu/open_access_theses/1161
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Indian Institute of Science
21.
Hima Bindu, V.
Experimental Investigations Of Aerothermodynamics Of A Scramjet Engine Configuration.
Degree: MSc Engg, Faculty of Engineering, 2011, Indian Institute of Science
URL: http://etd.iisc.ac.in/handle/2005/1120
► The recent resurgence in hypersonics is centered around the development of SCRAMJET engine technology to power future hypersonic vehicles. Successful flight trials by Australian and…
(more)
▼ The recent resurgence in hypersonics is centered around the development of SCRAMJET engine technology to power future
hypersonic vehicles. Successful flight trials by Australian and American scientists have created interest in the scramjet engine research across the globe. To develop scramjet engine, it is important to study heat transfer effects on the engine performance and aerodynamic forces acting on the body.
Hence, the main aim of present investigation is the design of scramjet engine configuration and measurement of aerodynamic forces acting on the model and heat transfer rates along the length of the combustor. The model is a two-dimensional single ramp model and is designed based on shock-on-lip (SOL) condition. Experiments are performed in IISc
hypersonic shock tunnel HST2 at two different Mach numbers of 8 and 7 for different angles of attack. Aerodynamic forces measurements using three-component accelerometer force balance and heat transfer rates measurements using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study.
Advisors/Committee Members: Reddy, K P J (advisor), Saravanan, S (advisor).
Subjects/Keywords: Aerothermodynamics; Scramjet Engines; Hypersonic Flows; Convective Heat Transfer; Aerodynamic Heating; Aerodynamic Forces; Hypersonic Shock Tunnel; Aerodynamic Measurements; Combustion Chambers - Heat Transfer; Accelerometer; Hypersonic Flow; Hypersonic Vehicles; Aeronautics
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APA (6th Edition):
Hima Bindu, V. (2011). Experimental Investigations Of Aerothermodynamics Of A Scramjet Engine Configuration. (Masters Thesis). Indian Institute of Science. Retrieved from http://etd.iisc.ac.in/handle/2005/1120
Chicago Manual of Style (16th Edition):
Hima Bindu, V. “Experimental Investigations Of Aerothermodynamics Of A Scramjet Engine Configuration.” 2011. Masters Thesis, Indian Institute of Science. Accessed March 01, 2021.
http://etd.iisc.ac.in/handle/2005/1120.
MLA Handbook (7th Edition):
Hima Bindu, V. “Experimental Investigations Of Aerothermodynamics Of A Scramjet Engine Configuration.” 2011. Web. 01 Mar 2021.
Vancouver:
Hima Bindu V. Experimental Investigations Of Aerothermodynamics Of A Scramjet Engine Configuration. [Internet] [Masters thesis]. Indian Institute of Science; 2011. [cited 2021 Mar 01].
Available from: http://etd.iisc.ac.in/handle/2005/1120.
Council of Science Editors:
Hima Bindu V. Experimental Investigations Of Aerothermodynamics Of A Scramjet Engine Configuration. [Masters Thesis]. Indian Institute of Science; 2011. Available from: http://etd.iisc.ac.in/handle/2005/1120
University of Toronto
22.
Chan, Jonathan.
Numerically Simulated Comparative Performance of a Scramjet and Shcramjet at Mach 11.
Degree: 2010, University of Toronto
URL: http://hdl.handle.net/1807/25446
► This study investigates the design and aeropropulsive performance of a complete, hydrogen powered, shock-induced combustion ramjet (shcramjet) at a flight Mach number of 11 and…
(more)
▼ This study investigates the design and aeropropulsive performance of a complete, hydrogen powered, shock-induced combustion ramjet (shcramjet) at a flight Mach number of 11 and altitude of 34.5 km. The design includes a Prandtl-Meyer compression inlet, cantilevered ramp fuel injectors, a shock-inducing wedge and a divergent nozzle. Numerical studies are undertaken using the WARP code that solves the three-dimensional Favre-averaged Navier-Stokes equations closed by the Wilcox k-ω turbulence model and the Jachimowski H2/air chemical kinetics model. Studies of fuel injection properties, mixing duct length, combustor wedge and nozzle geometry are completed to maximize the overall performance of the vehicle. The final shcramjet configuration generates a specific impulse of 1110 s. A comparison is undertaken with a scramjet vehicle at identical flight conditions and using many of the same components. The comparable scramjet generates a higher specific impulse of 1450 s although it is significantly larger and therefore heavier.
MAST
Advisors/Committee Members: Sislian, Jean Pascal, Aerospace Science and Engineering.
Subjects/Keywords: hypersonic; propulsion; scramjet; shcramjet; performance; simulation; 0538
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Manager
APA (6th Edition):
Chan, J. (2010). Numerically Simulated Comparative Performance of a Scramjet and Shcramjet at Mach 11. (Masters Thesis). University of Toronto. Retrieved from http://hdl.handle.net/1807/25446
Chicago Manual of Style (16th Edition):
Chan, Jonathan. “Numerically Simulated Comparative Performance of a Scramjet and Shcramjet at Mach 11.” 2010. Masters Thesis, University of Toronto. Accessed March 01, 2021.
http://hdl.handle.net/1807/25446.
MLA Handbook (7th Edition):
Chan, Jonathan. “Numerically Simulated Comparative Performance of a Scramjet and Shcramjet at Mach 11.” 2010. Web. 01 Mar 2021.
Vancouver:
Chan J. Numerically Simulated Comparative Performance of a Scramjet and Shcramjet at Mach 11. [Internet] [Masters thesis]. University of Toronto; 2010. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/1807/25446.
Council of Science Editors:
Chan J. Numerically Simulated Comparative Performance of a Scramjet and Shcramjet at Mach 11. [Masters Thesis]. University of Toronto; 2010. Available from: http://hdl.handle.net/1807/25446
University of Manchester
23.
Yang, Leichao.
Flow control using energy deposition at Mach 5.
Degree: PhD, 2012, University of Manchester
URL: https://www.research.manchester.ac.uk/portal/en/theses/flow-control-using-energy-deposition-at-mach-5(09bb331e-c3d3-4365-bdfc-6001a54c5702).html
;
http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.564303
► Flow control has always been an intense research subject with the pursuit of favourable control effects like drag reduction, transition delay, and separation prevention. In…
(more)
▼ Flow control has always been an intense research subject with the pursuit of favourable control effects like drag reduction, transition delay, and separation prevention. In practice, these flow control effects are achieved using mechanical actuators such as deflectors, vortex generators, transverse jets and so on. However, such mechanical actuators may face the drag penalty and limitation of actuation response time. In recent years, energy deposition has been suggested as a novel flow control technique in high-speed flow with preferable characteristics like non-intrusive, easy arrangement and high actuation frequency. The motivation of this work is to experimentally explore the flow behaviour after the certain amount of energy is deposited in Mach 5 flow. The energy deposition is implemented using a thermal bump (surface energy deposition) and laser beam focusing (volumetric energy deposition).This work starts with the development of a measurement technique of luminescent paint for the present challenging hypersonic testing environment, which is used for the further energy deposition experiment. The successes of the luminescent paint development is demonstrated both on two-dimensional and axisymmetric models. The luminescent paint shows high spatial resolution and the accuracy comparing to the pressure transducer reading. The surface energy deposition is performed using an embedded heating element (thermal bump) on a flat plate. Qualitative and quantitative measurement techniques are utilised to study the modification to the flow structure and the alteration to the distribution of pressure and heat transfer rate after thermal bump is activated. The results reveal the appearance of induced shock wave and suspicious vortices traces due to the activated thermal bump as reported in other literatures. Re-distribution of surface pressure and heat transfer rate are also found.For the volumetric energy deposition, the laser beam is firstly focused in quiescent air in order to understand the induced flow pattern and the impingement to a solid plate. High-speed schlieren photography is utilised to provide an insight to the dynamic evolution of the induced shock wave propagation and plasma kernel development after laser-induced air breakdown. Then, the laser energy deposition is conducted over a flat plate with the presence of Mach 5 flow. The outward motion of the induced shock wave significantly distorts the boundary layer and changes the surface pressure distribution. The results show the different pattern of boundary distortion when laser beam energy is deposited at different positions downstream of the leading edge of flat plate. The entire induced flow pattern is similar to those induced by a pulsed micro-jet. In spite of the laser pulse width of 4 ns, the entire dynamic process lasts about 100 μs.
Subjects/Keywords: 629.132; Hypersonic Flow; Energy Deposition;
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to Zotero / EndNote / Reference
Manager
APA (6th Edition):
Yang, L. (2012). Flow control using energy deposition at Mach 5. (Doctoral Dissertation). University of Manchester. Retrieved from https://www.research.manchester.ac.uk/portal/en/theses/flow-control-using-energy-deposition-at-mach-5(09bb331e-c3d3-4365-bdfc-6001a54c5702).html ; http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.564303
Chicago Manual of Style (16th Edition):
Yang, Leichao. “Flow control using energy deposition at Mach 5.” 2012. Doctoral Dissertation, University of Manchester. Accessed March 01, 2021.
https://www.research.manchester.ac.uk/portal/en/theses/flow-control-using-energy-deposition-at-mach-5(09bb331e-c3d3-4365-bdfc-6001a54c5702).html ; http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.564303.
MLA Handbook (7th Edition):
Yang, Leichao. “Flow control using energy deposition at Mach 5.” 2012. Web. 01 Mar 2021.
Vancouver:
Yang L. Flow control using energy deposition at Mach 5. [Internet] [Doctoral dissertation]. University of Manchester; 2012. [cited 2021 Mar 01].
Available from: https://www.research.manchester.ac.uk/portal/en/theses/flow-control-using-energy-deposition-at-mach-5(09bb331e-c3d3-4365-bdfc-6001a54c5702).html ; http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.564303.
Council of Science Editors:
Yang L. Flow control using energy deposition at Mach 5. [Doctoral Dissertation]. University of Manchester; 2012. Available from: https://www.research.manchester.ac.uk/portal/en/theses/flow-control-using-energy-deposition-at-mach-5(09bb331e-c3d3-4365-bdfc-6001a54c5702).html ; http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.564303
Delft University of Technology
24.
Van Dooren, R.M. (author).
The controllability of a winged hypersonic vehicle under aeroelastic effects.
Degree: 2016, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:e8527460-df6c-4a08-838d-247566fa4c44
Aerospace Engineering
Space Engineering
Space Flight
Advisors/Committee Members: Mooij, E. (mentor).Subjects/Keywords: Aeroelasticity; Controllability; hypersonic
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APA (6th Edition):
Van Dooren, R. M. (. (2016). The controllability of a winged hypersonic vehicle under aeroelastic effects. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:e8527460-df6c-4a08-838d-247566fa4c44
Chicago Manual of Style (16th Edition):
Van Dooren, R M (author). “The controllability of a winged hypersonic vehicle under aeroelastic effects.” 2016. Masters Thesis, Delft University of Technology. Accessed March 01, 2021.
http://resolver.tudelft.nl/uuid:e8527460-df6c-4a08-838d-247566fa4c44.
MLA Handbook (7th Edition):
Van Dooren, R M (author). “The controllability of a winged hypersonic vehicle under aeroelastic effects.” 2016. Web. 01 Mar 2021.
Vancouver:
Van Dooren RM(. The controllability of a winged hypersonic vehicle under aeroelastic effects. [Internet] [Masters thesis]. Delft University of Technology; 2016. [cited 2021 Mar 01].
Available from: http://resolver.tudelft.nl/uuid:e8527460-df6c-4a08-838d-247566fa4c44.
Council of Science Editors:
Van Dooren RM(. The controllability of a winged hypersonic vehicle under aeroelastic effects. [Masters Thesis]. Delft University of Technology; 2016. Available from: http://resolver.tudelft.nl/uuid:e8527460-df6c-4a08-838d-247566fa4c44
University of Minnesota
25.
Reinert, John.
Conjugate Heat Transfer Simulations for Hypersonic Vehicles.
Degree: PhD, Aerospace Engineering and Mechanics, 2020, University of Minnesota
URL: http://hdl.handle.net/11299/216841
► The accurate prediction of thermal responses is important for optimizing the design and operability for hypersonic flight vehicles. In order to efficiently simulate this process,…
(more)
▼ The accurate prediction of thermal responses is important for optimizing the design and operability for hypersonic flight vehicles. In order to efficiently simulate this process, a loosely coupled conjugate heat transfer solver was developed. Conjugate heat transfer simulations involve fluid and solid solvers. The fluid solver computes the flow field over the vehicle, and the solid solver calculates the transient heat conduction into the vehicle body. The two solvers are ``loosely'' coupled because both solvers exchange information at the surface of the vehicle, but operate on different time scales. The present work details the derivation of the conjugate heat transfer solver. The simulations were performed with US3D, an implicit finite volume unstructured compressible flow solver, with a newly developed implicit finite element transient heat conduction solver. The finite element solver is verified by comparing with analytical solutions for a bar, cylinder, and sphere. Validation cases for two geometries are shown: a fin-cone and HIFiRE-1. Both cases were shown to match well with the experimental data and flight test data. Additionally, the finite element method is compared to a finite volume method for solving the transient heat conduction equation. The comparison showed the benefits of the finite element method, such as refined temperature distribution and improved grid independence. Finally, the boundary layer transition (BoLT) vehicle is simulated for a segment of the trajectory. Results show the heating of the leading edge through time and the three-dimensional heating of the vehicle. At a specific time in the trajectory, the boundary layer and flow field are investigated. A comparative study is performed for the variable wall temperature and isothermal wall flow fields. The variable wall temperature was found to affect the wall heat flux and flow field structures. These results highlight the importance of performing conjugate heat transfer simulations when comparing to flight tests and experimental data.
Subjects/Keywords: aerothermal; CFD; conjugate heat transfer; hypersonic
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APA (6th Edition):
Reinert, J. (2020). Conjugate Heat Transfer Simulations for Hypersonic Vehicles. (Doctoral Dissertation). University of Minnesota. Retrieved from http://hdl.handle.net/11299/216841
Chicago Manual of Style (16th Edition):
Reinert, John. “Conjugate Heat Transfer Simulations for Hypersonic Vehicles.” 2020. Doctoral Dissertation, University of Minnesota. Accessed March 01, 2021.
http://hdl.handle.net/11299/216841.
MLA Handbook (7th Edition):
Reinert, John. “Conjugate Heat Transfer Simulations for Hypersonic Vehicles.” 2020. Web. 01 Mar 2021.
Vancouver:
Reinert J. Conjugate Heat Transfer Simulations for Hypersonic Vehicles. [Internet] [Doctoral dissertation]. University of Minnesota; 2020. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/11299/216841.
Council of Science Editors:
Reinert J. Conjugate Heat Transfer Simulations for Hypersonic Vehicles. [Doctoral Dissertation]. University of Minnesota; 2020. Available from: http://hdl.handle.net/11299/216841
Princeton University
26.
Grube, Nathan Elias.
Shock Wave – Turbulence Interactions
.
Degree: PhD, 2020, Princeton University
URL: http://arks.princeton.edu/ark:/88435/dsp01j6731673z
► Canonical shock / isotropic turbulence interactions (SITIs) are studied using direct numerical simulation (DNS) and perturbation analysis. Flow parameters include unprecedentedly high turbulence Mach numbers…
(more)
▼ Canonical shock / isotropic turbulence interactions (SITIs) are studied using direct numerical simulation (DNS) and perturbation analysis. Flow parameters include unprecedentedly high turbulence Mach numbers (Mt <= 0.7) and up to 15% dilatational turbulent kinetic energy (TKE).
These extreme conditions necessitate shock-capturing throughout the entire domain and the use of DNS inflow data from auxiliary forced isotropic turbulence simulations.
Three aspects of the DNS results are of particular interest: unprecedentedly high streamwise Reynolds stress amplification; mean flows that differ from classical solutions; and Reynolds stress anisotropy opposite to the predictions of linear theory.
These DNS results are elucidated by perturbation analyses. In high-Mt flows, both solenoidal and dilatational incident modes are important, but no existing work handles these modes in a convenient, unified way. Therefore, existing inviscid linear interaction analyses (LIAs) are reformulated in a general framework allowing any incident mode type, any inclination angle, and any shock obliquity. Integrated results are presented for isotropic incident fields of turbulence, sound, and entropy spots. LIA remains remarkably accurate (within 10% for TKE amplification) even for the strong turbulence considered here.
The new LIA is used as a starting point for second-order “quadratic interaction analysis” (QIA) and viscous LIA. QIA improves the mean predictions of classical theory; viscous LIA offers a possible explanation for a reported failure of single-wave LIA near so-called critical angles.
The anomalous post-shock Reynolds stress anisotropy is explained by the lengthscales of emitted vortical waves as a function of angle. The post-shock waves carrying the majority of the streamwise Reynolds stress are of longer wavelength than those carrying the transverse stress. This implies different timescales, with smaller scales losing the “memory” of their initial anisotropy faster than larger scales lose theirs. A model based on this understanding predicts anisotropies close to those observed.
Finally, the smoothing effects of shock motion (SM) are combined with QIA to give “SM-QIA” theory. The resulting mean pressure profiles closely match DNS data through the shock and downstream until the effects of nonlinear terms become important; SM-QIA thus provides an analog to the classical Rankine–Hugoniot shock-jump conditions applicable even in turbulent flows.
Advisors/Committee Members: Martin, Maria P (advisor).
Subjects/Keywords: Hypersonic;
Interaction;
Shock;
Shockwave;
Supersonic;
Turbulence
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APA (6th Edition):
Grube, N. E. (2020). Shock Wave – Turbulence Interactions
. (Doctoral Dissertation). Princeton University. Retrieved from http://arks.princeton.edu/ark:/88435/dsp01j6731673z
Chicago Manual of Style (16th Edition):
Grube, Nathan Elias. “Shock Wave – Turbulence Interactions
.” 2020. Doctoral Dissertation, Princeton University. Accessed March 01, 2021.
http://arks.princeton.edu/ark:/88435/dsp01j6731673z.
MLA Handbook (7th Edition):
Grube, Nathan Elias. “Shock Wave – Turbulence Interactions
.” 2020. Web. 01 Mar 2021.
Vancouver:
Grube NE. Shock Wave – Turbulence Interactions
. [Internet] [Doctoral dissertation]. Princeton University; 2020. [cited 2021 Mar 01].
Available from: http://arks.princeton.edu/ark:/88435/dsp01j6731673z.
Council of Science Editors:
Grube NE. Shock Wave – Turbulence Interactions
. [Doctoral Dissertation]. Princeton University; 2020. Available from: http://arks.princeton.edu/ark:/88435/dsp01j6731673z
Indian Institute of Science
27.
Srinath, S.
Experimental Study Of Large Angle Blunt Cone With Telescopic Aerospike Flying At Hypersonic Mach Numbers.
Degree: MSc Engg, Faculty of Engineering, 2011, Indian Institute of Science
URL: http://etd.iisc.ac.in/handle/2005/1116
► The emerging and competitive environment in the space technology requires the improvements in the capability of aerodynamic vehicles. This leads to the analysis in drag…
(more)
▼ The emerging and competitive environment in the space technology requires the improvements in the capability of aerodynamic vehicles. This leads to the analysis in drag reduction of the vehicle along with the minimized heat transfer rate. Using forward facing solid aerospike is the simplest way among the existing drag reduction methodologies for
hypersonic blunt cone bodies. But the flow oscillations associated with this aerospike makes it difficult to implement. When analyzing this flow, it can be understood that this oscillating flow can be compared to conical cavity flow. Therefore in the spiked flows, it is decided to implement the technique used in reducing the flow oscillation of the cavities. Based on this method the shallow conical cavity flow generated by the aerospike fixed ahead of a 120o blunt cone body is fissured as multiple cavities by so many disks formed from 10o cone. Now the deep conical cavities had the length to mean depth ratio of unity; this suppresses the unnecessary oscillations of the shallow cavity. The total length of the telescopic aerospike is fixed as 100mm. And one another conical tip plain aerospike of same length is designed for comparing the telescopic spike’s performance at
hypersonic flow Mach numbers of 5.75 and 7.9.
A three component force balance system capable of measuring drag, lift and pitching moment is designed and mounted internally into the skirt of the model. Drag measurement is done for without spike, conical tip plain spiked and telescopic spiked blunt cone body. The three configurations are tested at different angles of attack from 0 to 10 degree with a step of 2. A discrete iterative deconvolution methodology is implemented in this research work for obtaining the clean drag history from the noisy drag accelerometer signal. The drag results showed the drag reduction when compared to the without spike blunt cone body. When comparing to the plain spiked, the telescopic spiked blunt cone body has lesser drag at higher angles of attack.
Heat transfer measurements are done over the blunt cone surface using the Platinum thin film gauges formed over the Macor substrate. These results and the flow visualization give better understanding of the flow and the heat flux rate caused by the flow. The enhancement in the heat flux rate over the blunt cone surface is due to the shock interaction. And in recirculation region the heat flux rate is very much lesser when compared to without spike blunt cone body. It is observed that the shock interaction in the windward side is coming closer towards the nose of the blunt cone as the angle of attack increases and the oscillation of the oblique shock also decreases.
Schlieren visualization showed that there is dispersion in the oblique shock, particularly in the leeward side. In the telescopic spike there are multiple shocks generated from each and every disk which coalesces together to form a single oblique shock. And the effect of the shock generated by the telescopic spike is stronger than the effect of the shock generated by the…
Advisors/Committee Members: Jagadeesh, G (advisor).
Subjects/Keywords: Hypersonic Aerodynamics; Aerospike; Hypersonic Mach Numbers; Hypersonic Flow; Drag (Aerodynamics); Aerodynamic Heating; Cone-Aerodynamics; Telescopic Aerospike; Hypersonic Shock Tunnel HST2; Blunt Cone; Aeronautics
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APA (6th Edition):
Srinath, S. (2011). Experimental Study Of Large Angle Blunt Cone With Telescopic Aerospike Flying At Hypersonic Mach Numbers. (Masters Thesis). Indian Institute of Science. Retrieved from http://etd.iisc.ac.in/handle/2005/1116
Chicago Manual of Style (16th Edition):
Srinath, S. “Experimental Study Of Large Angle Blunt Cone With Telescopic Aerospike Flying At Hypersonic Mach Numbers.” 2011. Masters Thesis, Indian Institute of Science. Accessed March 01, 2021.
http://etd.iisc.ac.in/handle/2005/1116.
MLA Handbook (7th Edition):
Srinath, S. “Experimental Study Of Large Angle Blunt Cone With Telescopic Aerospike Flying At Hypersonic Mach Numbers.” 2011. Web. 01 Mar 2021.
Vancouver:
Srinath S. Experimental Study Of Large Angle Blunt Cone With Telescopic Aerospike Flying At Hypersonic Mach Numbers. [Internet] [Masters thesis]. Indian Institute of Science; 2011. [cited 2021 Mar 01].
Available from: http://etd.iisc.ac.in/handle/2005/1116.
Council of Science Editors:
Srinath S. Experimental Study Of Large Angle Blunt Cone With Telescopic Aerospike Flying At Hypersonic Mach Numbers. [Masters Thesis]. Indian Institute of Science; 2011. Available from: http://etd.iisc.ac.in/handle/2005/1116
University of New South Wales
28.
Bright, Courtney.
Secondary Injection Thrust Vector Control for Spacecraft Propulsion.
Degree: Engineering & Information Technology, 2019, University of New South Wales
URL: http://handle.unsw.edu.au/1959.4/64934
;
https://unsworks.unsw.edu.au/fapi/datastream/unsworks:63059/SOURCE02?view=true
► This thesis explores the application of secondary injection thrust vector control (SITVC) to spacecraft propulsion, where propellant consumption and system complexity are often increased by…
(more)
▼ This thesis explores the application of secondary injection thrust vector control (SITVC) to spacecraft propulsion, where propellant consumption and system complexity are often increased by the need to maintain alignment of the thrust vector with the spacecraft centre of mass. SITVC involves transverse injection of a secondary fluid into the main flow of a supersonic nozzle to produce a side force. While the technique has previously been studied for atmospheric vehicles with low area-ratio nozzles, this research is focussed on the performance of SITVC in high area-ratio nozzles expanding into vacuum.A two-dimensional numerical study was conducted to determine the influence of ambient pressure on SITVC performance. Performance was strongly affected by ambient pressure when the nozzle was overexpanded, but weakly affected when underexpanded. SITVC was modelled numerically for a three-dimensional, high area-ratio contour nozzle, where a maximum thrust vector angle of 16.4 degrees was achieved at a mass flow ratio of 0.343. The nozzle area-ratio, injector location, and injector angle were then varied to determine the effect on SITVC performance. A minor decrease in performance was observed with increasing nozzle area-ratio. SITVC efficiency was increased by moving the injector closer to the nozzle throat and by inclining it upstream, at the expense of maximum thrust vector angle.An experimental apparatus was designed to enable validation of the numerical results. The experimental results closely matched the numerical results for mass flow ratios up to 0.17, but the maximum thrust vector angle was reached at a lower mass flow ratio. These results were used to analytically model the performance of SITVC in the context of an in-orbit servicing mission, for comparison with standard reaction thrusters and mechanical thrust vector control (MTVC). SITVC was found to be more fuel efficient than reaction thrusters for the configuration studied, and was able to compensate for a much larger range of thrust misalignment. SITVC had a lower system mass than MTVC for thrust misalignments up to 9 degrees. MTVC became competitive at larger thrust misalignments, indicating a trade-off between system mass and complexity.
Advisors/Committee Members: Neely, Andrew, Engineering & Information Technology, UNSW Canberra, UNSW, Tuttle, Sean, Engineering & Information Technology, UNSW Canberra, UNSW.
Subjects/Keywords: Hypersonic; Propulsion; Injection pressure; Spacecraft; Vectoring
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APA (6th Edition):
Bright, C. (2019). Secondary Injection Thrust Vector Control for Spacecraft Propulsion. (Doctoral Dissertation). University of New South Wales. Retrieved from http://handle.unsw.edu.au/1959.4/64934 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:63059/SOURCE02?view=true
Chicago Manual of Style (16th Edition):
Bright, Courtney. “Secondary Injection Thrust Vector Control for Spacecraft Propulsion.” 2019. Doctoral Dissertation, University of New South Wales. Accessed March 01, 2021.
http://handle.unsw.edu.au/1959.4/64934 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:63059/SOURCE02?view=true.
MLA Handbook (7th Edition):
Bright, Courtney. “Secondary Injection Thrust Vector Control for Spacecraft Propulsion.” 2019. Web. 01 Mar 2021.
Vancouver:
Bright C. Secondary Injection Thrust Vector Control for Spacecraft Propulsion. [Internet] [Doctoral dissertation]. University of New South Wales; 2019. [cited 2021 Mar 01].
Available from: http://handle.unsw.edu.au/1959.4/64934 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:63059/SOURCE02?view=true.
Council of Science Editors:
Bright C. Secondary Injection Thrust Vector Control for Spacecraft Propulsion. [Doctoral Dissertation]. University of New South Wales; 2019. Available from: http://handle.unsw.edu.au/1959.4/64934 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:63059/SOURCE02?view=true
Delft University of Technology
29.
Carvalho Barreira Lages Ribeiro, André (author).
Optimal Aeroassisted Maneuvers for Orbital Transfer.
Degree: 2019, Delft University of Technology
URL: http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398
► Orbital plane changes require a considerable amount of propellant to be completed if the traditional fully propulsive maneuver is used. The cost associated with this…
(more)
▼ Orbital plane changes require a considerable amount of propellant to be completed if the traditional fully propulsive maneuver is used. The cost associated with this type of maneuvers is one of its major drawbacks, and performing an inclination change in orbit would require a large part of the total mass budget of the mission. Moreover, launching such vehicles or satellites would demand the usage of a larger launcher, further contributing to the increase of the mission cost. An alternative to the fully propulsive maneuver is the aeroassisted maneuver, proposed London (1961), where the vehicle performs one or more passes through the atmosphere of the central body. This allows for the aerodynamic forces to steer and brake the vehicle, reducing its orbital speed and changing the inclination of the orbit with less propellant required. To investigate the impact of the aeroassisted maneuver in the propellant requirements, the following research question was proposed: What is the impact of using an aeroassisted maneuver in reducing the amount of propellant needed to achieve a certain orbital inclination change? A simulation environment was developed using the TU Delft Astrodynamics Toolbox, capable of simulating an aeroassisted trajectory from a Geostationary Earth Orbit to a Low Earth Orbit, with an associated inclination change. A simple lateral guidance algorithm, capable of tracking the orbital plane, was also developed, and a node control algorithm was implemented to chose the guidance nodes during each individual trajectory. The performance of three vehicle configurations, with different lift-to-drag ratios, were investigated in terms of achievable inclination change. The HORUS-2B, a vehicle with a moderate to high lift-to-drag ratio, was the most suitable to perform the maneuver, as it allows for a larger inclination change while reducing the energy dissipation rate in the atmosphere. The simulator was integrated with an optimization algorithm, such that optimal trajectories could be investigated. The performance of the aeroassisted maneuver is measured in terms of three different objectives: the offset in inclination between the final and the target orbital planes, the velocity impulse applied at atmospheric exit to target the desired apogee and the maximum heat load experienced by the vehicle during any given atmospheric pass. The NSGA-II optimizer is selected to minimize the three objectives, and several parameters are tuned to increase the probability of finding the global optimum, namely the tuning parameters of the optimizer itself, the values of the constraint and the number of guidance nodes. It was found that an orbital plane of 20 degrees could be targeted with an error of only 0.28%, with the total maneuver requiring 87.32% less Delta V needed from thrust impulses. The heat load obtained is 1315.6 kJ/m2, which is well within the allowable range for Thermal Protection Systems currently available, and the total number of atmospheric passes is five. Moreover, it is possible to…
Advisors/Committee Members: Mooij, Erwin (mentor), Delft University of Technology (degree granting institution).
Subjects/Keywords: Optimization; Re-entry; Aeroassisted; Astrodynamics; Hypersonic
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APA (6th Edition):
Carvalho Barreira Lages Ribeiro, A. (. (2019). Optimal Aeroassisted Maneuvers for Orbital Transfer. (Masters Thesis). Delft University of Technology. Retrieved from http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398
Chicago Manual of Style (16th Edition):
Carvalho Barreira Lages Ribeiro, André (author). “Optimal Aeroassisted Maneuvers for Orbital Transfer.” 2019. Masters Thesis, Delft University of Technology. Accessed March 01, 2021.
http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398.
MLA Handbook (7th Edition):
Carvalho Barreira Lages Ribeiro, André (author). “Optimal Aeroassisted Maneuvers for Orbital Transfer.” 2019. Web. 01 Mar 2021.
Vancouver:
Carvalho Barreira Lages Ribeiro A(. Optimal Aeroassisted Maneuvers for Orbital Transfer. [Internet] [Masters thesis]. Delft University of Technology; 2019. [cited 2021 Mar 01].
Available from: http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398.
Council of Science Editors:
Carvalho Barreira Lages Ribeiro A(. Optimal Aeroassisted Maneuvers for Orbital Transfer. [Masters Thesis]. Delft University of Technology; 2019. Available from: http://resolver.tudelft.nl/uuid:89acba76-65ed-438d-810e-f3a863546398
University of Minnesota
30.
Houston, Mary.
Hypersonic Boundary Layer Stability Analysis Using Momentum Potential Theory.
Degree: MS, Aerospace Engineering and Mechanics, 2020, University of Minnesota
URL: http://hdl.handle.net/11299/217112
► Linear Stability Theory (LST) and the Parabolized Stability Equations (PSE) have provided valuable tools for analysis and prediction of laminar to turbulent transition for plates,…
(more)
▼ Linear Stability Theory (LST) and the Parabolized Stability Equations (PSE) have provided valuable tools for analysis and prediction of laminar to turbulent transition for plates, sharp cones, and geometries for which parallel-flow or a slowly-varying boundary layer can be assumed. However, these techniques struggle to capture the complex flow-physics present near the tip of blunt-cones. Input-output analysis has been used in conjunction with direct numerical simulation to capture the effects of nose bluntness on downstream stability. Using the results of the input-output analysis we apply momentum potential theory (MPT) to preform fluid-thermodynamic (FT) decomposition, separating disturbances into their vortical, thermal and acoustic components. A reference case of Mach 6 flow over a flat-plate is computed and output responses are compared to the results for Mach 6 flow over a blunt-cone of 7o half angle. Perturbation eigenfunctions and structures are examined in the areas of second-mode amplification. For both the flat-plate and blunt-cone the vortical components are the largest, followed by the thermal then acoustic components. Fluid-thermodynamic structures in the second-mode amplification region of blunt-cone show wall-normal stretching above the critical layer. Fluid-thermodynamic decomposition of full-domain input and output results for the blunt-cone geometry are considered. It is found that input sensitivity is highest at the top of the entropy layer and along the boundary layer edge for the fore-half of the cone. Output response in the streamwise direction is highest in the regions between the generalized inflection point (GIP) and the boundary layer edge and dissipates near the surface, whereas wall-normal response extends to the surface and shows a local minimum between the GIP and boundary layer edge. To compliment existing studies on hypersonic boundary layer response to surface roughness/ vibration we look at input sensitivity and output response at the surface. It is found that there is greater sensitivity to wall-normal forcing than streamwise forcing at the surface and among the three FT components in this direction the vortical had the highest relative output amplitude. Finally, total fluctuating enthalpy (TFE) is considered for both the flat-plate and blunt-cone, in both cases the thermal terms provides the strongest source of TFE.
Subjects/Keywords: CDF; Fluid dynamics; High speed; Hypersonic; Transition
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APA ·
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APA (6th Edition):
Houston, M. (2020). Hypersonic Boundary Layer Stability Analysis Using Momentum Potential Theory. (Masters Thesis). University of Minnesota. Retrieved from http://hdl.handle.net/11299/217112
Chicago Manual of Style (16th Edition):
Houston, Mary. “Hypersonic Boundary Layer Stability Analysis Using Momentum Potential Theory.” 2020. Masters Thesis, University of Minnesota. Accessed March 01, 2021.
http://hdl.handle.net/11299/217112.
MLA Handbook (7th Edition):
Houston, Mary. “Hypersonic Boundary Layer Stability Analysis Using Momentum Potential Theory.” 2020. Web. 01 Mar 2021.
Vancouver:
Houston M. Hypersonic Boundary Layer Stability Analysis Using Momentum Potential Theory. [Internet] [Masters thesis]. University of Minnesota; 2020. [cited 2021 Mar 01].
Available from: http://hdl.handle.net/11299/217112.
Council of Science Editors:
Houston M. Hypersonic Boundary Layer Stability Analysis Using Momentum Potential Theory. [Masters Thesis]. University of Minnesota; 2020. Available from: http://hdl.handle.net/11299/217112
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