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University of Florida
1.
Hanon, David S.
Scramjet Energetic Anlaysis concerning Upstream Fuel Injection.
Degree: MS, Aerospace Engineering - Mechanical and Aerospace Engineering, 2018, University of Florida
URL: https://ufdc.ufl.edu/UFE0052361 
► Injecting fuel in the inlet of a scramjet is a technique aimed at elevating the propulsive efficiency of this technology. This paper has sought to…
(more)
▼ Injecting fuel in the inlet of a
scramjet is a technique aimed at elevating the propulsive efficiency of this technology. This paper has sought to understand the physics and influence of this act on the
scramjet engine from a fluid dynamic and thermodynamic perspective.
Advisors/Committee Members: SEGAL,CORIN (committee chair).
Subjects/Keywords: scramjet
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APA (6th Edition):
Hanon, D. S. (2018). Scramjet Energetic Anlaysis concerning Upstream Fuel Injection. (Masters Thesis). University of Florida. Retrieved from https://ufdc.ufl.edu/UFE0052361
Chicago Manual of Style (16th Edition):
Hanon, David S. “Scramjet Energetic Anlaysis concerning Upstream Fuel Injection.” 2018. Masters Thesis, University of Florida. Accessed January 24, 2021.
https://ufdc.ufl.edu/UFE0052361.
MLA Handbook (7th Edition):
Hanon, David S. “Scramjet Energetic Anlaysis concerning Upstream Fuel Injection.” 2018. Web. 24 Jan 2021.
Vancouver:
Hanon DS. Scramjet Energetic Anlaysis concerning Upstream Fuel Injection. [Internet] [Masters thesis]. University of Florida; 2018. [cited 2021 Jan 24].
Available from: https://ufdc.ufl.edu/UFE0052361.
Council of Science Editors:
Hanon DS. Scramjet Energetic Anlaysis concerning Upstream Fuel Injection. [Masters Thesis]. University of Florida; 2018. Available from: https://ufdc.ufl.edu/UFE0052361
University of Notre Dame
2.
Nozomu Kato.
Flow Phenomena under a Backpressure in an Axisymmetric
Scramjet Model</h1>.
Degree: Aerospace and Mechanical Engineering, 2019, University of Notre Dame
URL: https://curate.nd.edu/show/02870v8624p
► The supersonic combustion ramjet (scramjet) engine has been developed as a promising propulsion system for hypersonic vehicles. The primary objective of the current study…
(more)
▼ The supersonic combustion ramjet (
scramjet)
engine has been developed as a promising propulsion system for
hypersonic vehicles. The primary objective of the current study is
to investigate the flow features under various backpressure
conditions in axisymmetric scramjets. Particularly, the current
study focuses on the flow behaviors under inlet unstart, a critical
phenomenon resulting in a loss of thrust. Three angled-wedge inlets
(
α = 0-, 6-, and 9-degree) were used to
investigate the effect of the incident shock wave on unstart. Flow
visualizations and wall pressure measurements provided qualitative
and quantitative results. The fuel injection was
used to increase the backpressure in low-enthalpy conditions to
trigger unstart. The flow features were investigated by using a
flow visualization technique of the planar laser Rayleigh
scattering (PLRS). For the flow without jet injection, the PLRS
visualized flow features such as the axisymmetric flow structure,
transition and the growth of the boundary layer, the conical
incident shock wave, the shock wave-boundary layer interaction
(SBLI), and the shock-shock interaction (SSI). The wall pressure
measurement showed that there were pressure gradients in the
isolator imposed by the incident shock wave for
α = 6 and 9, whereas minor pressure gradients
were present for
α = 0. For the choked flow,
the PLRS images depicted the pseudoshock propagating upstream
within the isolator. The pseudoshock advanced upstream as the J
increased and anchored at a certain location. The pseudoshock
location fluctuated to a certain extent for all the inlets, but the
pseudoshock for
α = 0 fluctuated more than
α = 6 and 9. The wall pressure measurement
suggested the presence of the termination shock wave and the
subsonic acceleration in the combustor. The schlieren
imaging system was used to determine the unstart thresholds. The
unstart threshold in terms of the jet momentum ratio was highest
for
α = 6 and lowest for
α = 0. The lower unstart threshold of
α = 9 compared to
α = 6
was attributed to the higher downstream pressure of the baseflow
for
α = 9; less pressure rise was needed to
choke the flow for
α = 9. The smaller capture
area of
α = 0 may be the cause of the lowest
unstart threshold of
α = 0; the lower total
momentum of the captured flow for
α = 0 was
not able to hold the propagating unstart shock. The PLRS visualized
the unstart shock that propagates upstream. In addition, the
quasi-steady and transient motions of the unstart shock were
observed due to the adverse pressure gradient imposed by the
incident shock wave. The unstart shock velocity was estimated from
the wall pressure data, and the shock velocities for
α = 6 and 9 suggested the transient and the
quasi-steady motions of the unstart shock. The
effect of combustion heat release on unstart was investigated under
high-enthalpy conditions. Direct flame imaging and CH*…
Advisors/Committee Members: David Go, Committee Member, Aleksander Jemcov, Committee Member, Seong-kyun Im, Research Director.
Subjects/Keywords: scramjet; unstart
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APA (6th Edition):
Kato, N. (2019). Flow Phenomena under a Backpressure in an Axisymmetric
Scramjet Model</h1>. (Thesis). University of Notre Dame. Retrieved from https://curate.nd.edu/show/02870v8624p
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):
Kato, Nozomu. “Flow Phenomena under a Backpressure in an Axisymmetric
Scramjet Model</h1>.” 2019. Thesis, University of Notre Dame. Accessed January 24, 2021.
https://curate.nd.edu/show/02870v8624p.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Kato, Nozomu. “Flow Phenomena under a Backpressure in an Axisymmetric
Scramjet Model</h1>.” 2019. Web. 24 Jan 2021.
Vancouver:
Kato N. Flow Phenomena under a Backpressure in an Axisymmetric
Scramjet Model</h1>. [Internet] [Thesis]. University of Notre Dame; 2019. [cited 2021 Jan 24].
Available from: https://curate.nd.edu/show/02870v8624p.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Kato N. Flow Phenomena under a Backpressure in an Axisymmetric
Scramjet Model</h1>. [Thesis]. University of Notre Dame; 2019. Available from: https://curate.nd.edu/show/02870v8624p
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of Notre Dame
3.
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 (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 January 24, 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. 24 Jan 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 Jan 24].
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 Minnesota
4.
Tylczak, Erik.
Computational Study of Shock/Plume Interactions Between Multiple Jets in Supersonic Crossflow.
Degree: PhD, Aerospace Engineering and Mechanics, 2016, University of Minnesota
URL: http://hdl.handle.net/11299/182745
► The interaction of multiple jets in supersonic crossflow is simulated using hybrid Reynolds- Averaged Navier Stokes and Large Eddy Simulation turbulence models. The blockage of…
(more)
▼ The interaction of multiple jets in supersonic crossflow is simulated using hybrid Reynolds- Averaged Navier Stokes and Large Eddy Simulation turbulence models. The blockage of a jet generates a curved bow shock, and in multi-jet flows, each shock impinges on the other fuel plumes. The curved nature of each shock generates vorticity directly, and the impingement of each shock on the vortical structures within the adjacent fuel plumes strengthens vortical structures already present. These stirring motions are the major driver of fuel-air mixing, and so mixing enhancement is predicted to occur in multi-port configurations. The primary geometry considered is that of the combustion duct at the Calspan- University of Buffalo Research Center 48” Large Energy National Shock (LENS) tunnel. This geometry was developed to be representative of the geometry and flow physics of the Flight 2 test vehicle of the Hypersonic International Flight Research Experimenta- tion Program (HiFIRE-2). This geometry takes the form of a symmetric pair of external compression ramps that feed an isolator of approximately 4” × 1” cross-section. Nine interdigitated flush-wall injectors, four on one wall and five on the other, inject hydrogen at an angle of 30 degrees to the freestream. Two freestream flow conditions are consid- ered: approximately Mach 7.2 at a static temperature of 214K and a density of 0.039 kg/m3 for the five-injector case, and approximately Mach 8.9 at a static temperature of 167K and density of 0.014 kg/m3 for the nine-injector case. Validation computations are performed on a single-port experiment with an imposed shock wave. Unsteady calculations are performed on five-port and nine-port configura- tions, and the five-port configuration is compared to calculations performed with only a single active port on the same geometry. Analysis of statistical data demonstrates enhanced mixing in the multi-port configurations in regions where shock impingement occurs.
Subjects/Keywords: CFD; Crossflow; Injection; Mixing; Scramjet
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APA (6th Edition):
Tylczak, E. (2016). Computational Study of Shock/Plume Interactions Between Multiple Jets in Supersonic Crossflow. (Doctoral Dissertation). University of Minnesota. Retrieved from http://hdl.handle.net/11299/182745
Chicago Manual of Style (16th Edition):
Tylczak, Erik. “Computational Study of Shock/Plume Interactions Between Multiple Jets in Supersonic Crossflow.” 2016. Doctoral Dissertation, University of Minnesota. Accessed January 24, 2021.
http://hdl.handle.net/11299/182745.
MLA Handbook (7th Edition):
Tylczak, Erik. “Computational Study of Shock/Plume Interactions Between Multiple Jets in Supersonic Crossflow.” 2016. Web. 24 Jan 2021.
Vancouver:
Tylczak E. Computational Study of Shock/Plume Interactions Between Multiple Jets in Supersonic Crossflow. [Internet] [Doctoral dissertation]. University of Minnesota; 2016. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/11299/182745.
Council of Science Editors:
Tylczak E. Computational Study of Shock/Plume Interactions Between Multiple Jets in Supersonic Crossflow. [Doctoral Dissertation]. University of Minnesota; 2016. Available from: http://hdl.handle.net/11299/182745
North Carolina State University
5.
Keistler, Patrick G.
A Variable Turbulent Prandtl and Schmidt Number Model Study for Scramjet Applications.
Degree: PhD, Aerospace Engineering, 2009, North Carolina State University
URL: http://www.lib.ncsu.edu/resolver/1840.16/3373
► A turbulence model that allows for the calculation of the variable turbulent Prandtl (Prt) and Schmidt (Sct) numbers as part of the solution is presented.…
(more)
▼ A turbulence model that allows for the calculation of the variable turbulent Prandtl (Prt) and Schmidt (Sct) numbers as part of the solution is presented. The model also accounts for the interactions between turbulence and chemistry by modeling the corresponding terms. Four equations are added to the baseline k-ζ turbulence model: two equations for enthalpy variance and its dissipation rate to calculate the turbulent diffusivity, and two equations for the concentrations variance and its dissipation rate to calculate the turbulent diffusion coefficient. The underlying turbulence model already accounts for compressibility effects. The variable Prt/Sct turbulence model is validated and tuned by simulating a wide variety of experiments. Included in the experiments are two-dimensional, axisymmetric, and three-dimensional mixing and combustion cases. The combustion cases involved either hydrogen and air, or hydrogen, ethylene, and air. Two chemical kinetic models are employed for each of these situations. For the hydrogen and air cases, a seven species/seven reaction model where the reaction rates are temperature dependent and a nine species/nineteen reaction model where the reaction rates are dependent on both pressure and temperature are used. For the cases involving ethylene, a 15 species/44 reaction reduced model that is both pressure and temperature dependent is used, along with a 22 species/18 global reaction reduced model that makes use of the quasi-steady-state approximation.
In general, fair to good agreement is indicated for all simulated experiments. The turbulence/chemistry interaction terms are found to have a significant impact on flame location for the two-dimensional combustion case, with excellent experimental agreement when the terms are included. In most cases, the hydrogen chemical mechanisms behave nearly identically, but for one case, the pressure dependent model would not auto-ignite at the same conditions as the experiment and the other chemical model. The model was artificially ignited in that case. For the cases involving ethylene combustion, the chemical model has a profound impact on the flame size, shape, and ignition location. However, without quantitative experimental data, it is difficult to determine which one is more suitable for this particular application.
Advisors/Committee Members: Fred R. DeJarnette, Committee Member (advisor), Jack R. Edwards, Committee Member (advisor), Hassan A. Hassan, Committee Chair (advisor), Pierre A. Gremaud, Committee Member (advisor).
Subjects/Keywords: scramjet; schmidt; prandtl; turbulence modeling; cfd
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APA (6th Edition):
Keistler, P. G. (2009). A Variable Turbulent Prandtl and Schmidt Number Model Study for Scramjet Applications. (Doctoral Dissertation). North Carolina State University. Retrieved from http://www.lib.ncsu.edu/resolver/1840.16/3373
Chicago Manual of Style (16th Edition):
Keistler, Patrick G. “A Variable Turbulent Prandtl and Schmidt Number Model Study for Scramjet Applications.” 2009. Doctoral Dissertation, North Carolina State University. Accessed January 24, 2021.
http://www.lib.ncsu.edu/resolver/1840.16/3373.
MLA Handbook (7th Edition):
Keistler, Patrick G. “A Variable Turbulent Prandtl and Schmidt Number Model Study for Scramjet Applications.” 2009. Web. 24 Jan 2021.
Vancouver:
Keistler PG. A Variable Turbulent Prandtl and Schmidt Number Model Study for Scramjet Applications. [Internet] [Doctoral dissertation]. North Carolina State University; 2009. [cited 2021 Jan 24].
Available from: http://www.lib.ncsu.edu/resolver/1840.16/3373.
Council of Science Editors:
Keistler PG. A Variable Turbulent Prandtl and Schmidt Number Model Study for Scramjet Applications. [Doctoral Dissertation]. North Carolina State University; 2009. Available from: http://www.lib.ncsu.edu/resolver/1840.16/3373
University of Toronto
6.
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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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 January 24, 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. 24 Jan 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 Jan 24].
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
Clemson University
7.
Quinones, Matthew.
Numerical Analysis of Scramjet Cavity Flameholders at Varying Flight Mach Numbers.
Degree: MS, Mechanical Engineering, 2018, Clemson University
URL: https://tigerprints.clemson.edu/all_theses/3010
► Various types of air breathing engines are used as propulsion devices in aviation. At high flight velocities, the use of a ramjet or supersonic combustion…
(more)
▼ Various types of air breathing engines are used as propulsion devices in aviation. At high
flight velocities, the use of a ramjet or supersonic combustion ramjet (
scramjet) may be
preferred due to the natural compressibility of air at high speed. A
scramjet, while
similar to the ramjet, does not slow air to subsonic speeds prior to combustion, allowing
it to operate at much higher flight velocities at very high altitude. For this reason,
however, the residence time of air inside of the combustor is on the order of milliseconds,
requiring rapid mixing and ignition of the fuel to generate adequate thrust. To do this, a
flameholder is often used, which generates turbulence, shock waves, and maintains a
recirculation region through geometric effects. In this study, four geometry types
involving eighteen separate designs were chosen and analyzed using CFD software.
Isolator inlet Mach numbers of 2.2, 4, 6, 8, and 10 were selected to model varying flight
velocity, and hydrogen fuel was injected sonically at all injector locations with a single
step reaction mechanism applied for combustion. An existing square cavity model was
chosen and modified to produce slanted cavity, double cavity, and combined strut-cavity
designs. The flameholders were analyzed in a non-reacting simulation to observe their
effects on the flow field and fuel mixing efficiency. Reacting simulations were
performed for each flameholder to investigate flame stabilization capabilities, thermal
choking, stagnation pressure losses and drag generated inside of the combustor. Results
show that all designs sustain a flame during combustion at all flight Mach numbers.
However, the square cavity with a back cavity injector does this while limiting losses and
drag due to shock wave formation, thermal choking, and geometric effects in the flow.
Advisors/Committee Members: Dr. Richard Miller, Committee Chair, Dr. John Saylor, Dr. Xiangchun Xuan.
Subjects/Keywords: Flame holder; Flame stability; Ramjet; Scramjet
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APA (6th Edition):
Quinones, M. (2018). Numerical Analysis of Scramjet Cavity Flameholders at Varying Flight Mach Numbers. (Masters Thesis). Clemson University. Retrieved from https://tigerprints.clemson.edu/all_theses/3010
Chicago Manual of Style (16th Edition):
Quinones, Matthew. “Numerical Analysis of Scramjet Cavity Flameholders at Varying Flight Mach Numbers.” 2018. Masters Thesis, Clemson University. Accessed January 24, 2021.
https://tigerprints.clemson.edu/all_theses/3010.
MLA Handbook (7th Edition):
Quinones, Matthew. “Numerical Analysis of Scramjet Cavity Flameholders at Varying Flight Mach Numbers.” 2018. Web. 24 Jan 2021.
Vancouver:
Quinones M. Numerical Analysis of Scramjet Cavity Flameholders at Varying Flight Mach Numbers. [Internet] [Masters thesis]. Clemson University; 2018. [cited 2021 Jan 24].
Available from: https://tigerprints.clemson.edu/all_theses/3010.
Council of Science Editors:
Quinones M. Numerical Analysis of Scramjet Cavity Flameholders at Varying Flight Mach Numbers. [Masters Thesis]. Clemson University; 2018. Available from: https://tigerprints.clemson.edu/all_theses/3010
University of New South Wales
8.
Van Pelt, Hilbert.
Shock Vector Control applied to a Converging-Diverging Nozzle and a Hypersonic Vehicle.
Degree: Aerospace & Mechanical Engineering, 2018, University of New South Wales
URL: http://handle.unsw.edu.au/1959.4/60744
;
https://unsworks.unsw.edu.au/fapi/datastream/unsworks:53083/SOURCE02?view=true
► Shock vector control (SVC) is a technique to control airplanes and spaceplanes by means of fluidinjection into the nozzle flow. SVC is part of fluidic…
(more)
▼ Shock vector control (SVC) is a technique to control airplanes and spaceplanes by means of fluidinjection into the nozzle flow. SVC is part of fluidic thrust vectoring, a technique whereby acontrol force can be generated by means of fluid injection in the engine exhaust.For this thesis the application of shock vector control to a converging-diverging nozzle and ahypersonic vehicle was investigated. The experiments with a converging-diverging nozzle withan injector took place in an experimental facility which houses extruded two-dimensional nozzleprofiles of 8 mm thick. The profiles are sealed by two windows on either side of the nozzle foroptical access. The nozzle is a linear converging-diverging nozzle with a sonic injector placed at71% length of the diverging section. On this geometry two-dimensional steady state simulationshave been performed, as well as time accurate simulations and a large eddy simulation. It isshown that the injected mass creates a blockage for the incoming flow generating a shock wave.This blockage causes the core flow to separate before the injector, meaning that there is a highpressure region before the injector creating the sideforce. It was demonstrated numerically andexperimentally that the sideforce, relative to the axial force, has a linear relationship with theinjected mass until the shock wave hits the opposite nozzle wall causing a shockwave boundarylayer interaction that is detrimental to the net sideforce.It is shown that this type of nozzle flow can be modelled with analytical relations based onblast wave theory. The established analytical methods were calibrated based on an orifice injectorand therefore the calibration constant needed to be changed from 1.2 to 2.36 because of the fullwidth injector. In an orifice injector the cross flow can go around the injected fluid, while in atwo-dimensional flow case all the nozzle flow is being turned by the injected fluid, and thereforethe calibration constant needs to be increased.It was demonstrated that for the small scale experiment the force establishment time is 0.2msec, which is faster than most solenoid valves available on the market. Analytically it is shownthat the establishment time will increase with the size of the nozzle, but it will remain fast enoughto be used as a control system for a full scale rocket or
scramjet.The experimental supersonic geometry was modelled with Large eddy simulation (LES) andcompared to Reynolds averaged Navier-Stokes (RANS) simulations. It was shown that the flowshows three-dimensional effects at the point where the sidewall meets the injector, but is moreiiitwo-dimensional towards the middle of the nozzle.To test the hypersonic implementation of SVC a
scramjet of similar shape to the X-43 wasdesigned with a single slot injector close to the trailing edge of the nozzle. The pressure in thehypersonic environment is orders of magnitude lower than in the supersonic experiment and thereforea large expansion of the Mach disk at the injection point was seen. An increase of sideforceand moment is present…
Advisors/Committee Members: Neely, Andrew James, Aerospace & Mechanical Engineering, UNSW Canberra, UNSW, Young, John, Aerospace & Mechanical Engineering, UNSW Canberra, UNSW.
Subjects/Keywords: Scramjet; Thrust vectoring; Shock vector control
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APA (6th Edition):
Van Pelt, H. (2018). Shock Vector Control applied to a Converging-Diverging Nozzle and a Hypersonic Vehicle. (Doctoral Dissertation). University of New South Wales. Retrieved from http://handle.unsw.edu.au/1959.4/60744 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:53083/SOURCE02?view=true
Chicago Manual of Style (16th Edition):
Van Pelt, Hilbert. “Shock Vector Control applied to a Converging-Diverging Nozzle and a Hypersonic Vehicle.” 2018. Doctoral Dissertation, University of New South Wales. Accessed January 24, 2021.
http://handle.unsw.edu.au/1959.4/60744 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:53083/SOURCE02?view=true.
MLA Handbook (7th Edition):
Van Pelt, Hilbert. “Shock Vector Control applied to a Converging-Diverging Nozzle and a Hypersonic Vehicle.” 2018. Web. 24 Jan 2021.
Vancouver:
Van Pelt H. Shock Vector Control applied to a Converging-Diverging Nozzle and a Hypersonic Vehicle. [Internet] [Doctoral dissertation]. University of New South Wales; 2018. [cited 2021 Jan 24].
Available from: http://handle.unsw.edu.au/1959.4/60744 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:53083/SOURCE02?view=true.
Council of Science Editors:
Van Pelt H. Shock Vector Control applied to a Converging-Diverging Nozzle and a Hypersonic Vehicle. [Doctoral Dissertation]. University of New South Wales; 2018. Available from: http://handle.unsw.edu.au/1959.4/60744 ; https://unsworks.unsw.edu.au/fapi/datastream/unsworks:53083/SOURCE02?view=true
Rutgers University
9.
Pizzaia, Adam R., 1989-.
Side view and end view mixing measurements of a sonic transverse jet in a supersonic crossflow.
Degree: MS, Mechanical and Aerospace Engineering, 2014, Rutgers University
URL: https://rucore.libraries.rutgers.edu/rutgers-lib/44199/
► The primary investigation was to determine if mixing of a sonic transverse jet in the side-view plane was correlated to or equal to the mixing…
(more)
▼ The primary investigation was to determine if mixing of a sonic transverse jet in the side-view plane was correlated to or equal to the mixing in the end-view plane. The analysis was conducted by imaging a sonic jet in a supersonic crossflow at jet momentum ratios of 1.2, 2.7, and 5.2. Because of the difficulty of placing a camera streamwise to a supersonic crossflow to obtain end view images, off axis imaging was carried out. Overall, the jets appear to be more symmetric when injected into the thin boundary layer than in the thick boundary layer. Centerline probability density functions in the end view plane were compared with probability density functions in the side view plane. The probability of finding unmixed fluid at the windward side of the jet was higher than finding unmixed fluid toward the center or near the orifice plane. Different paths other than those that fall on the mid planes were analyzed and compared in the end-view plane. The effect of the boundary layer thickness on the penetration and mixing of the sonic jet in supersonic flow was investigated.
Advisors/Committee Members: Knight, Doyle D (chair), Diez, Francisco J (internal member), Rossmann, Tobias (outside member).
Subjects/Keywords: Airplanes – Scramjet engines; Aerodynamics, Supersonic; Supersonic planes
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APA (6th Edition):
Pizzaia, Adam R., 1. (2014). Side view and end view mixing measurements of a sonic transverse jet in a supersonic crossflow. (Masters Thesis). Rutgers University. Retrieved from https://rucore.libraries.rutgers.edu/rutgers-lib/44199/
Chicago Manual of Style (16th Edition):
Pizzaia, Adam R., 1989-. “Side view and end view mixing measurements of a sonic transverse jet in a supersonic crossflow.” 2014. Masters Thesis, Rutgers University. Accessed January 24, 2021.
https://rucore.libraries.rutgers.edu/rutgers-lib/44199/.
MLA Handbook (7th Edition):
Pizzaia, Adam R., 1989-. “Side view and end view mixing measurements of a sonic transverse jet in a supersonic crossflow.” 2014. Web. 24 Jan 2021.
Vancouver:
Pizzaia, Adam R. 1. Side view and end view mixing measurements of a sonic transverse jet in a supersonic crossflow. [Internet] [Masters thesis]. Rutgers University; 2014. [cited 2021 Jan 24].
Available from: https://rucore.libraries.rutgers.edu/rutgers-lib/44199/.
Council of Science Editors:
Pizzaia, Adam R. 1. Side view and end view mixing measurements of a sonic transverse jet in a supersonic crossflow. [Masters Thesis]. Rutgers University; 2014. Available from: https://rucore.libraries.rutgers.edu/rutgers-lib/44199/
10.
Rodrigues, Luís.
Thermodynamic analysis and optimization of a scramjet engine with thermal management system.
Degree: 2010, RCAAP
URL: https://www.rcaap.pt/detail.jsp?id=oai:ubibliorum.ubi.pt:10400.6/3686
► Thermal management of the scramjet engine is one of the key issues of the challenges brought by the development of hypersonic airbreathing vehicles. A Closed…
(more)
▼ Thermal management of the
scramjet engine is one of the key issues of the challenges brought by the development of hypersonic airbreathing vehicles. A Closed Brayton Cycle thermal management system for a regenerative cooled
scramjet is introduced with the goal of reducing the hydrogen fuel flow for cooling.
Part of the heat absorbed from fuel is converted into other forms of energy to decrease the heat that must be taken away by hydrogen fuel. Reducing this heat increases the fuel heat sink (cooling capacity)
without requiring excess fuel for cooling and eliminating the need to search for a new coolant. The proposed thermal cycle reduces the fuel flow for cooling, and this way, the fuel on board assures the cooling requirements for the whole hypersonic vehicle. The basic concept and working principle are introduced:
a thermodynamic cycle analysis is performed to demonstrate the system performance gains of Closed Brayton Cycle (CBC) Thermal Management System (TMS) over the conventional system with regenerative
cooling. It was shown that the Closed Brayton Cycle Thermal Management Systems presents a high performance gain when compared to conventional regenerative cooling due to the reduction of fuel
flow for cooling and additional power output.
Advisors/Committee Members: Brójo, Francisco Miguel Ribeiro Proença.
Subjects/Keywords: Scramjet (motor) - Veículos hipersónicos; Arrefecimento regenerativo - Motor - Veículo hipersónico; Arrefecimento regenerativo - Veículo hipersónico; Gestão térmica - Scramjet (motor) - Veículos hipersónicos
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APA (6th Edition):
Rodrigues, L. (2010). Thermodynamic analysis and optimization of a scramjet engine with thermal management system. (Thesis). RCAAP. Retrieved from https://www.rcaap.pt/detail.jsp?id=oai:ubibliorum.ubi.pt:10400.6/3686
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):
Rodrigues, Luís. “Thermodynamic analysis and optimization of a scramjet engine with thermal management system.” 2010. Thesis, RCAAP. Accessed January 24, 2021.
https://www.rcaap.pt/detail.jsp?id=oai:ubibliorum.ubi.pt:10400.6/3686.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Rodrigues, Luís. “Thermodynamic analysis and optimization of a scramjet engine with thermal management system.” 2010. Web. 24 Jan 2021.
Vancouver:
Rodrigues L. Thermodynamic analysis and optimization of a scramjet engine with thermal management system. [Internet] [Thesis]. RCAAP; 2010. [cited 2021 Jan 24].
Available from: https://www.rcaap.pt/detail.jsp?id=oai:ubibliorum.ubi.pt:10400.6/3686.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Rodrigues L. Thermodynamic analysis and optimization of a scramjet engine with thermal management system. [Thesis]. RCAAP; 2010. Available from: https://www.rcaap.pt/detail.jsp?id=oai:ubibliorum.ubi.pt:10400.6/3686
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Virginia Tech
11.
Prebola, John L. Jr.
Performance of a Plasma Torch with Hydrocarbon Feedstocks for Use in Scramjet Combustion.
Degree: MS, Aerospace and Ocean Engineering, 1998, Virginia Tech
URL: http://hdl.handle.net/10919/36941
► Research was conducted at Virginia Tech on a high-pressure uncooled plasma torch to study torch operational characteristics with hydrocarbon feedstocks and to determine the feasibility…
(more)
▼ Research was conducted at Virginia Tech on a high-pressure uncooled plasma torch to study torch operational characteristics with hydrocarbon feedstocks and to determine the feasibility of using the torch as an igniter in
scramjet applications. Operational characteristics studied included electrical properties, such as arc stability, voltage-current characteristics and start/re-start capabilities, and mechanical properties, such as coking, electrode erosion and transient to steady-state torch body temperature trends. Possible use of the plasma torch as an igniter in high-speed combustion environments was investigated through the use of emission spectroscopy and a NASA chemical kinetics code.
All feedstocks tested; argon, methane, ethylene and propylene, were able to start. The voltage data indicated that there were two preferred operating modes, which were well defined for methane. For all gases, a higher current setting, on the order of 40 A, led to more stable torch operation. A low intensity, high frequency current applied to the torch, along with the primary DC current, resulted in virtual elimination of soot deposits on the anodes. Electrode erosion was found to multiply each time the complexity of the hydrocarbon was increased. Audio and high-speed visual analysis led to identification of 180 Hz plasma formation cycle, related to the three-phase power supply. The spectroscopic analysis aided in the identification of combustion enhancing radicals being produced by the torch, and results of the chemical kinetics analysis verified combustion enhancement and radical production through the use of a basic plasma model. Overall, the results of this study indicate that the plasma torch is a promising source for
scramjet ignition, and further study is warranted.
Advisors/Committee Members: Schetz, Joseph A. (committeechair), O'Brien, Walter F. Jr. (committee member), Yates, Charlie L. (committee member).
Subjects/Keywords: plasma torch; hydrocarbon feedstocks; scramjet; scramjet combustion
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APA (6th Edition):
Prebola, J. L. J. (1998). Performance of a Plasma Torch with Hydrocarbon Feedstocks for Use in Scramjet Combustion. (Masters Thesis). Virginia Tech. Retrieved from http://hdl.handle.net/10919/36941
Chicago Manual of Style (16th Edition):
Prebola, John L Jr. “Performance of a Plasma Torch with Hydrocarbon Feedstocks for Use in Scramjet Combustion.” 1998. Masters Thesis, Virginia Tech. Accessed January 24, 2021.
http://hdl.handle.net/10919/36941.
MLA Handbook (7th Edition):
Prebola, John L Jr. “Performance of a Plasma Torch with Hydrocarbon Feedstocks for Use in Scramjet Combustion.” 1998. Web. 24 Jan 2021.
Vancouver:
Prebola JLJ. Performance of a Plasma Torch with Hydrocarbon Feedstocks for Use in Scramjet Combustion. [Internet] [Masters thesis]. Virginia Tech; 1998. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/10919/36941.
Council of Science Editors:
Prebola JLJ. Performance of a Plasma Torch with Hydrocarbon Feedstocks for Use in Scramjet Combustion. [Masters Thesis]. Virginia Tech; 1998. Available from: http://hdl.handle.net/10919/36941
Purdue University
12.
Williams, Joseph.
Trajectory optimization using indirect methods and parametric scramjet cycle analysis.
Degree: MSin Aeronautics and Astronautics, Aeronautics and Astronautics, 2016, Purdue University
URL: https://docs.lib.purdue.edu/open_access_theses/908
► This study investigates the solution of time sensitive regional strike trajectories for hypersonic missiles. This minimum time trajectory is suspected to be best performed…
(more)
▼ This study investigates the solution of time sensitive regional strike trajectories for hypersonic missiles. This minimum time trajectory is suspected to be best performed by
scramjet powered hypersonic missiles which creates strong coupled interaction between the flight dynamics and the performance of the engine. Comprehensive engine models are necessary to gain better insight into
scramjet propulsion. Separately, robust and comprehensive trajectory analysis provides references for vehicles to fly along. However, additional observation and understanding is obtained by integrating the propulsion model inside the trajectory framework. Going beyond curve fitted thrusting models, an integrated
scramjet cycle analysis offers rapid trade studies on engine parameters and enables the identification of the most significant and optimal engine parameters for the mission as a whole. Regularization of bang-bang control problems by use of the Epsilon-Trig regularization method has created the possibility to preserve the original equations of motion while still solving these problems through indirect methods. Indirect methods incorporate mathematical information from the optimal control problem to provide high quality, integrated solutions. The minimum time optimal trajectory of a rocket propelled missile is compared to that of a
scramjet powered missile to determine the advantages of
scramjet technology in this application.
Advisors/Committee Members: Michael J. Grant, Michael J. Grant, William Crossley, Stephen Heister.
Subjects/Keywords: Applied sciences; Indirect; Optimization; Powered; Regularization; Scramjet; Trajectory; Aerospace Engineering
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APA (6th Edition):
Williams, J. (2016). Trajectory optimization using indirect methods and parametric scramjet cycle analysis. (Thesis). Purdue University. Retrieved from https://docs.lib.purdue.edu/open_access_theses/908
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):
Williams, Joseph. “Trajectory optimization using indirect methods and parametric scramjet cycle analysis.” 2016. Thesis, Purdue University. Accessed January 24, 2021.
https://docs.lib.purdue.edu/open_access_theses/908.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Williams, Joseph. “Trajectory optimization using indirect methods and parametric scramjet cycle analysis.” 2016. Web. 24 Jan 2021.
Vancouver:
Williams J. Trajectory optimization using indirect methods and parametric scramjet cycle analysis. [Internet] [Thesis]. Purdue University; 2016. [cited 2021 Jan 24].
Available from: https://docs.lib.purdue.edu/open_access_theses/908.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Williams J. Trajectory optimization using indirect methods and parametric scramjet cycle analysis. [Thesis]. Purdue University; 2016. Available from: https://docs.lib.purdue.edu/open_access_theses/908
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of Michigan
13.
Mbagwu, Chukwuka.
Combustion Efficiency, Flameout Operability Limits and General Design Optimization for Integrated Ramjet-Scramjet Hypersonic Vehicles.
Degree: PhD, Aerospace Engineering, 2017, University of Michigan
URL: http://hdl.handle.net/2027.42/140954
► High speed, air-breathing hypersonic vehicles encounter a varied range of engine and operating conditions traveling along cruise/ascent missions at high altitudes and dynamic pressures. Variations…
(more)
▼ High speed, air-breathing hypersonic vehicles encounter a varied range of engine and operating conditions traveling along cruise/ascent missions at high altitudes and dynamic pressures. Variations of ambient pressure, temperature, Mach number, and dynamic pressure can affect the combustion conditions in conflicting ways. Computations were performed to understand propulsion tradeoffs that occur when a hypersonic vehicle travels along an ascent trajectory. Proper Orthogonal Decomposition methods were applied for the reduction of flamelet chemistry data in an improved combustor model. Two operability limits are set by requirements that combustion efficiency exceed selected minima and flameout be avoided. A method for flameout prediction based on empirical Damkohler number measurements is presented. Operability limits are plotted that define allowable flight corridors on an altitude versus flight Mach number performance map; fixed-acceleration ascent trajectories were considered for this study.
Several design rules are also presented for a hypersonic waverider with a dual-mode
scramjet engine. Focus is placed on ``vehicle integration" design, differing from previous ``propulsion-oriented" design optimization. The well-designed waverider falls between that of an aircraft (high lift-to-drag ratio) and a rocket (high thrust-to-drag ratio). 84 variations of an X-43-like vehicle were run using the MASIV
scramjet reduced order model to examine performance tradeoffs. Informed by the vehicle design study, variable-acceleration trajectory optimization was performed for three constant dynamic pressures ascents. Computed flameout operability limits were implemented as additional constraints to the optimization problem.
The Michigan-AFRL
Scramjet In-Vehicle (MASIV) waverider model includes finite-rate chemistry, applied scaling laws for 3-D turbulent mixing, ram-scram transition and an empirical value of the flameout Damkohler number. A reduced-order modeling approach is justified (in lieu of higher-fidelity computational simulations) because all vehicle forces are computed multiple thousands of times to generate multi-dimensional performance maps.
The findings of this thesis work present a number of compelling conclusions. It is found that the ideal operating conditions of a
scramjet engine are heavily dependent on the ambient and combustor pressure (and less strongly on temperature). Combustor pressures of approximately 1.0 bar or greater achieve the highest combustion efficiency, in line with industry standards of more than 0.5 bar. Ascent trajectory analysis of combustion efficiency and lean-limit flameout dictate best operation at higher dynamic pressures and lower altitudes, but these goals are traded off by current structural limitations whereby dynamic pressures must remain below 100 kPa. Hypersonic waverider designs varied between an "airplane" and a "rocket" are found to have better performance with the latter design, with controllability and minimum elevon/rudder surface area as a stability constraint for the vehicle…
Advisors/Committee Members: Driscoll, James F (committee member), Johnsen, Eric (committee member), Martins, Joaquim R R A (committee member), Raman, Venkatramanan (committee member).
Subjects/Keywords: hypersonic vehicles; scramjet; optimization; combustion; Aerospace Engineering; Engineering
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APA (6th Edition):
Mbagwu, C. (2017). Combustion Efficiency, Flameout Operability Limits and General Design Optimization for Integrated Ramjet-Scramjet Hypersonic Vehicles. (Doctoral Dissertation). University of Michigan. Retrieved from http://hdl.handle.net/2027.42/140954
Chicago Manual of Style (16th Edition):
Mbagwu, Chukwuka. “Combustion Efficiency, Flameout Operability Limits and General Design Optimization for Integrated Ramjet-Scramjet Hypersonic Vehicles.” 2017. Doctoral Dissertation, University of Michigan. Accessed January 24, 2021.
http://hdl.handle.net/2027.42/140954.
MLA Handbook (7th Edition):
Mbagwu, Chukwuka. “Combustion Efficiency, Flameout Operability Limits and General Design Optimization for Integrated Ramjet-Scramjet Hypersonic Vehicles.” 2017. Web. 24 Jan 2021.
Vancouver:
Mbagwu C. Combustion Efficiency, Flameout Operability Limits and General Design Optimization for Integrated Ramjet-Scramjet Hypersonic Vehicles. [Internet] [Doctoral dissertation]. University of Michigan; 2017. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/2027.42/140954.
Council of Science Editors:
Mbagwu C. Combustion Efficiency, Flameout Operability Limits and General Design Optimization for Integrated Ramjet-Scramjet Hypersonic Vehicles. [Doctoral Dissertation]. University of Michigan; 2017. Available from: http://hdl.handle.net/2027.42/140954
Vanderbilt University
14.
Grady, Nathan Ryan.
Laser Diagnostics of Turbulent Flames in High Speed Flows.
Degree: PhD, Mechanical Engineering, 2015, Vanderbilt University
URL: http://hdl.handle.net/1803/13454
► High speed turbulent flames can be found in most aerospace propulsion devices. In order to understand the behavior and phenomenological nature of combustion in these…
(more)
▼ High speed turbulent flames can be found in most aerospace propulsion devices. In order to understand the behavior and phenomenological nature of combustion in these devices time resolved, non-invasive, in situ laser diagnostic measurements are needed. In this dissertation, laser diagnostics are used to study freely propagating turbulent premixed flames, and non premixed flames in a model
scramjet combustor. In addition, a new laser method is developed to measure velocity profiles in reacting regions without seed particles.
Premixed turbulent flames have been studied extensively over burner stabilized flames. However, the flame propagation in these burners is usually affected by the burner itself resulting in a geometric dependent propagation. Therefore, a fundamental understanding of the interaction between turbulence and premixed flames is obscured. Alternatively, spherically propagating flames (also known as flame kernels) that do not have these geometric dependencies can be studied using flame bombs. While flame bomb studies have been very successful in elucidating the interaction between turbulence and premixed flames, they typically have limited optical access and a mean radial inflow which can inhibit flame propagation. Therefore, a new means of studying flame kernels is described using a turbulent wind tunnel where flame kernels are allowed to freely propagate downstream. The results obtained in this new device are compared with traditional flame bomb measurements.
The reaction progress of a non-premixed combustion inside a cavity-piloted
scramjet combustor was determined by measuring all major species and temperature using spontaneous UV Raman scattering. A 70% CH4/30% H2 fuel blend was used to approximate the reactivity of liquid jet fuels, and minimize the number of Raman spectral lines to ensure tractable data sets. Inside the cavity, H2 fuel quickly burnt off while the CH4 and CO persisted until the fuel path reached the cavity shear layer.
Finally, a new molecular tagging velocimetry method of obtaining non-seeded velocity measurements inside a reaction zone is described. This method uses a two-photon process to dissociate H2O and create vibrationally excited OH photofragments which can then be differentiated from flame generated OH radicals that are predominately in the vibrational ground state to make single-shot velocity measurements.
Advisors/Committee Members: A.V. Anilkumar (committee member), Suresh Menon (committee member), Joseph Wehrmeyer (committee member), Greg Walker (committee member), Robert W. Pitz (Committee Chair).
Subjects/Keywords: flame kernel; molecular tagging velocimetry; scramjet; turbulent combustion; laser diagnostics
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APA (6th Edition):
Grady, N. R. (2015). Laser Diagnostics of Turbulent Flames in High Speed Flows. (Doctoral Dissertation). Vanderbilt University. Retrieved from http://hdl.handle.net/1803/13454
Chicago Manual of Style (16th Edition):
Grady, Nathan Ryan. “Laser Diagnostics of Turbulent Flames in High Speed Flows.” 2015. Doctoral Dissertation, Vanderbilt University. Accessed January 24, 2021.
http://hdl.handle.net/1803/13454.
MLA Handbook (7th Edition):
Grady, Nathan Ryan. “Laser Diagnostics of Turbulent Flames in High Speed Flows.” 2015. Web. 24 Jan 2021.
Vancouver:
Grady NR. Laser Diagnostics of Turbulent Flames in High Speed Flows. [Internet] [Doctoral dissertation]. Vanderbilt University; 2015. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/1803/13454.
Council of Science Editors:
Grady NR. Laser Diagnostics of Turbulent Flames in High Speed Flows. [Doctoral Dissertation]. Vanderbilt University; 2015. Available from: http://hdl.handle.net/1803/13454
North Carolina State University
15.
Star, Jason Blue.
Numerical Simulation of Scramjet Combustion in a Shock Tunnel.
Degree: PhD, Aerospace Engineering, 2005, North Carolina State University
URL: http://www.lib.ncsu.edu/resolver/1840.16/5028
► Three-dimensional computational simulations of reactive flowfields within a hydrogen-fueled scramjet-like geometry experimentally tested in a free piston shock tunnel are presented. The experimental configuration (Odam…
(more)
▼ Three-dimensional computational simulations of reactive flowfields within a hydrogen-fueled
scramjet-like geometry experimentally tested in a free piston shock tunnel are presented. The experimental configuration (Odam and Paull, AIAA Paper 2003-5244) involves injection of hydrogen fuel into the
scramjet inlet, followed by mixing, shock-induced ignition, and combustion. The predictions for both fuel-off and fuel-on conditions were observed to be sensitive to the choice of the wall temperature boundary conditions. The best comparison with experimental data were achieved through the implementation of an approach that involves a simplified conjugate heat transfer model that couples the heat conduction through the wall with the heat conduction of the fluid within the boundary layer. This approach is able to predict thermal loads on the walls of the
scramjet model due to shock wave interactions and due to heat release. As such, it is able to more accurately represent the physical temperature response of the engine model. Also shown to produce very good agreement with the statistically-steady experimental data was the isothermal ghost-cell boundary condition, which is based on a simplification of the time-dependent conjugate heat transfer boundary condition. This simplified boundary condition assumes a linear temperature distribution within the wall based on the effective depth that an applied heat load would penetrate, thus, it also allows the actual wall temperature to vary in response to the applied heat load.
Results for fuel-off simulations showed that the solution generated by a steady-state simulation implementing the isothermal ghost-cell wall boundary condition was very comparable with the statistically-steady solution obtained from a fully transient simulation with coupled heat conduction within the walls.
When integrated in a fully time-accurate manner, the fuel-on simulations showed a striking sensitivity to the modeled rate of air ingestion into the engine. For experimental data that showed steady combustion, the transient simulations resulted in either a steady combusting solution or a progression toward engine unstart, depending on the modeled rate of air ingestion. Also, for experimental data that showed an unsteady thermal choking event leading to eventual unstart, the transient simulations were able to predict both unstart and steady combustion, once again depending on the air ingestion rate. In all cases, the modeled air ingestion process is an approximation of the actual experimental process, in that uniform conditions are imposed as linear functions of time over the inlet plane.
The computational results also provide some support for a 'radical-farming' hypothesis, proposed to explain the ability of the hydrogen-air mixture to auto-ignite at relatively low inlet contraction ratios.
Advisors/Committee Members: Jack R. Edwards Jr., Committee Chair (advisor), Hassan A. Hassan, Committee Member (advisor), D. S. McRae, Committee Member (advisor), William L. Roberts, Committee Member (advisor).
Subjects/Keywords: propulsion; computational; CFD; scramjet
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APA (6th Edition):
Star, J. B. (2005). Numerical Simulation of Scramjet Combustion in a Shock Tunnel. (Doctoral Dissertation). North Carolina State University. Retrieved from http://www.lib.ncsu.edu/resolver/1840.16/5028
Chicago Manual of Style (16th Edition):
Star, Jason Blue. “Numerical Simulation of Scramjet Combustion in a Shock Tunnel.” 2005. Doctoral Dissertation, North Carolina State University. Accessed January 24, 2021.
http://www.lib.ncsu.edu/resolver/1840.16/5028.
MLA Handbook (7th Edition):
Star, Jason Blue. “Numerical Simulation of Scramjet Combustion in a Shock Tunnel.” 2005. Web. 24 Jan 2021.
Vancouver:
Star JB. Numerical Simulation of Scramjet Combustion in a Shock Tunnel. [Internet] [Doctoral dissertation]. North Carolina State University; 2005. [cited 2021 Jan 24].
Available from: http://www.lib.ncsu.edu/resolver/1840.16/5028.
Council of Science Editors:
Star JB. Numerical Simulation of Scramjet Combustion in a Shock Tunnel. [Doctoral Dissertation]. North Carolina State University; 2005. Available from: http://www.lib.ncsu.edu/resolver/1840.16/5028
University of Toronto
16.
West, Jonathan Philip.
Numerical Analysis of Ethylene Injection in the Inlet of a Mach Six Scramjet.
Degree: 2011, University of Toronto
URL: http://hdl.handle.net/1807/29643
► A scramjet inlet was designed for use on a small scale, Mach six, ethylene-fuelled vehicle. The inlet used strut-based cantilevered fuel injectors and a well-defined…
(more)
▼ A scramjet inlet was designed for use on a small scale, Mach six, ethylene-fuelled vehicle. The inlet used strut-based cantilevered fuel injectors and a well-defined mixing duct to mix fuel prior to the combustor. Designed using theoretical and numerical analyses, the resulting inlet configuration consisted of a single body shock inlet with vertical fuel injector struts and four cantilevered injectors per strut side. This inlet was 80 cm long and 42 cm high. Numerical analysis of the vehicle was conducted with computational fluid dynamics by solving the Favre-Averaged Navier-Stokes equations; turbulence was simulated using the Wilcox k-ω model. Multispecies simulations in two and three dimensions were used to evaluate the design. Analysis of the simulated flow features, thrust potential and mixing efficiency demonstrated favourable vehicle performance. In particular, the inlet allowed for complete combustion when lean equivalence ratios of less than 0.7 were used.
MAST
Advisors/Committee Members: Sislian, Jean Pascal, Aerospace Science and Engineering.
Subjects/Keywords: scramjet; ethylene; hypersonic; fuel injection; fuel mixing; computational fluid dynamics; 0538
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APA (6th Edition):
West, J. P. (2011). Numerical Analysis of Ethylene Injection in the Inlet of a Mach Six Scramjet. (Masters Thesis). University of Toronto. Retrieved from http://hdl.handle.net/1807/29643
Chicago Manual of Style (16th Edition):
West, Jonathan Philip. “Numerical Analysis of Ethylene Injection in the Inlet of a Mach Six Scramjet.” 2011. Masters Thesis, University of Toronto. Accessed January 24, 2021.
http://hdl.handle.net/1807/29643.
MLA Handbook (7th Edition):
West, Jonathan Philip. “Numerical Analysis of Ethylene Injection in the Inlet of a Mach Six Scramjet.” 2011. Web. 24 Jan 2021.
Vancouver:
West JP. Numerical Analysis of Ethylene Injection in the Inlet of a Mach Six Scramjet. [Internet] [Masters thesis]. University of Toronto; 2011. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/1807/29643.
Council of Science Editors:
West JP. Numerical Analysis of Ethylene Injection in the Inlet of a Mach Six Scramjet. [Masters Thesis]. University of Toronto; 2011. Available from: http://hdl.handle.net/1807/29643
Oregon State University
17.
Nahorniak, Matthew T.
Feasibility of Lorentz mixing to enhance combustion in supersonic diffusion flames.
Degree: MS, Mechanical Engineering, 1996, Oregon State University
URL: http://hdl.handle.net/1957/34208
► The purpose of this research was to determine if it is feasible to apply Lorentz mixing to supersonic diffusion flames, such as those found in…
(more)
▼ The purpose of this research was to determine if it is feasible to apply Lorentz
mixing to supersonic diffusion flames, such as those found in
SCRAMjet engines. The
combustion rate in supersonic diffusion flames is limited by the rate at which air and fuel
mix. Lorentz mixing increases turbulence within a flow, which increases the rate at which
species mix and thus increases the rate of combustion.
In order to determine the feasibility of Lorentz mixing for this application, a two-dimensional model of supersonic reacting flow with the application of a Lorentz force has been examined numerically. The flow model includes the complete Navier-Stokes equations, the ideal gas law, and terms to account for diffusion of chemical species, heat release due to chemical reaction, change in species density due to chemical reaction, and the Lorentz forces applied during Lorentz mixing. In addition, the Baldwin-Lomax turbulence model is used to approximate turbulent transport properties.
A FORTRAN program using the MacCormack method, a commonly used computational fluid dynamics algorithm, was used to solve the governing equations. The accuracy of the program was verified by using the program to model flows with known solutions.
Results were obtained for flows with Lorentz forces applied over a series of power levels and frequencies. The results show significant increases in the rate of combustion
when Lorentz mixing is applied. The amount of power required to drive Lorentz mixing is small relative to the rate at which energy is released in the chemical reaction. An optimum frequency at which to apply Lorentz mixing was also found for the flow being considered.
The results of the current study show that Lorentz mixing looks promising for increasing combustion rates in supersonic reacting flows, and that future study is warranted. In particular, researchers attempting to improve combustion in
SCRAMjet engines may want to consider Lorentz mixing as a way to improve combustion.
Advisors/Committee Members: Pattee, Heidi A. (advisor).
Subjects/Keywords: Airplanes – Scramjet engines – Mathematical models
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APA (6th Edition):
Nahorniak, M. T. (1996). Feasibility of Lorentz mixing to enhance combustion in supersonic diffusion flames. (Masters Thesis). Oregon State University. Retrieved from http://hdl.handle.net/1957/34208
Chicago Manual of Style (16th Edition):
Nahorniak, Matthew T. “Feasibility of Lorentz mixing to enhance combustion in supersonic diffusion flames.” 1996. Masters Thesis, Oregon State University. Accessed January 24, 2021.
http://hdl.handle.net/1957/34208.
MLA Handbook (7th Edition):
Nahorniak, Matthew T. “Feasibility of Lorentz mixing to enhance combustion in supersonic diffusion flames.” 1996. Web. 24 Jan 2021.
Vancouver:
Nahorniak MT. Feasibility of Lorentz mixing to enhance combustion in supersonic diffusion flames. [Internet] [Masters thesis]. Oregon State University; 1996. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/1957/34208.
Council of Science Editors:
Nahorniak MT. Feasibility of Lorentz mixing to enhance combustion in supersonic diffusion flames. [Masters Thesis]. Oregon State University; 1996. Available from: http://hdl.handle.net/1957/34208
University of Kansas
18.
Tourani, Chandraprakash Chandra.
COMPUTATIONAL SIMULATION OF SCRAMJET COMBUSTORS - A COMPARISON BETWEEN QUASI-ONE DIMENSIONAL AND 2-D NUMERICAL SIMULATIONS.
Degree: MS, Aerospace Engineering, 2011, University of Kansas
URL: http://hdl.handle.net/1808/7635
► 1-D simulations based on the quasi-one-dimensional equations of fluid motion plus an ignition delay model and 2-D numerical simulations based on Reynolds-Averaged Navier-Stokes (RANS) equations…
(more)
▼ 1-D simulations based on the quasi-one-dimensional equations of fluid motion plus an ignition delay model and 2-D numerical simulations based on Reynolds-Averaged Navier-Stokes (RANS) equations have been performed for two different
scramjet combustors. The combustor configurations at DLR and NASA's SCHOLAR Supersonic Combustor have been used as test cases for the 1-D and 2-D simulations. Comparisons between the published 3-D computational and experimental results and quasi-one-dimensional and 2-D simulations have been performed. The quasi-one dimensional modeling of NASA's SCHOLAR supersonic combustor captures the trends in Mach number, static pressure and static temperature for both cold flow and combustion case. The comparison with experimental result for combustion case reveals a close agreement with the pressure peak and the presence of an ignition delay. Thus, 1-D simulation very closely predicts the flow evolution within the combustor. On the other hand, for DLR supersonic combustor, due to the lack of oblique wave (i.e. shock waves and expansion waves) and shear dominated viscous flow simulation, 1-D model severely fails to predict the trend followed by the experimental result along the centerline of the combustor. However, the 1-D model is able to match the overall flow velocity achieved within the combustor downstream of the wedge at approximately six wedge chord lengths.
Advisors/Committee Members: Farokhi, Saeed (advisor), Taghavi, Ray (cmtemember), Keshmiri, Shahriar (cmtemember).
Subjects/Keywords: Aerospace engineering; Air-breathing propulsion; Combustor; Hypersonic; Quasi one-dimensional; Scramjet
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to Zotero / EndNote / Reference
Manager
APA (6th Edition):
Tourani, C. C. (2011). COMPUTATIONAL SIMULATION OF SCRAMJET COMBUSTORS - A COMPARISON BETWEEN QUASI-ONE DIMENSIONAL AND 2-D NUMERICAL SIMULATIONS. (Masters Thesis). University of Kansas. Retrieved from http://hdl.handle.net/1808/7635
Chicago Manual of Style (16th Edition):
Tourani, Chandraprakash Chandra. “COMPUTATIONAL SIMULATION OF SCRAMJET COMBUSTORS - A COMPARISON BETWEEN QUASI-ONE DIMENSIONAL AND 2-D NUMERICAL SIMULATIONS.” 2011. Masters Thesis, University of Kansas. Accessed January 24, 2021.
http://hdl.handle.net/1808/7635.
MLA Handbook (7th Edition):
Tourani, Chandraprakash Chandra. “COMPUTATIONAL SIMULATION OF SCRAMJET COMBUSTORS - A COMPARISON BETWEEN QUASI-ONE DIMENSIONAL AND 2-D NUMERICAL SIMULATIONS.” 2011. Web. 24 Jan 2021.
Vancouver:
Tourani CC. COMPUTATIONAL SIMULATION OF SCRAMJET COMBUSTORS - A COMPARISON BETWEEN QUASI-ONE DIMENSIONAL AND 2-D NUMERICAL SIMULATIONS. [Internet] [Masters thesis]. University of Kansas; 2011. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/1808/7635.
Council of Science Editors:
Tourani CC. COMPUTATIONAL SIMULATION OF SCRAMJET COMBUSTORS - A COMPARISON BETWEEN QUASI-ONE DIMENSIONAL AND 2-D NUMERICAL SIMULATIONS. [Masters Thesis]. University of Kansas; 2011. Available from: http://hdl.handle.net/1808/7635
University of Kansas
19.
Hwang, Seung-Jae.
Numerical Simulation of Enhanced Mixing in Scramjet Combustor Using Ramp, Tabs and Suction Collar.
Degree: PhD, Aerospace Engineering, 2011, University of Kansas
URL: http://hdl.handle.net/1808/8080
► Numerical simulations of the scramjet combustor by using the commercial CFD code Fluent with the coupled implicit method with second-order accurate discretization have been obtained…
(more)
▼ Numerical simulations of the
scramjet combustor by using the commercial CFD code Fluent with the coupled implicit method with second-order accurate discretization have been obtained for the reacting flows with the parallel fuel injection (ramp injection) and normal fuel injection (wall injection) schemes. Incorporated in the
scramjet combustors are delta tabs and suction collars of two types as means of mixing enhancement. The main mechanism of the tabs and suction collars for mixing enhancement is the generation of streamwise vorticity and providing outstanding flameholding capability along with the induced global instability of the shear layer. The idea has been previously recommended for mixing enhancement of the
scramjet combustor, but no experimental or computational data on the combustor performance has been reported, yet. The finite rate reaction model is used for the species transport model that only considers four species, H2, O2, H2O and N2. Vitiated air (mass fraction of O2, H2O, and N2 being 0.198, 0.139, and 0.663, respectively) enters the combustor at Mach number of 2.5 at a stagnation temperature and pressure of 1500 K and 101,325 Pa, respectively. The equivalence ratio is fixed at 0.45 in the present study. An optimization study of the combinations of the tabs and suction collars has been performed. Uninstalled thrust force for the optimal combination which was composed of the relieved ramp, 4 delta tabs, suction collar type I and 4 delta tabs in the fuel inlet scheme produced an additional 73% increase in thrust with only an additional 3.37% loss of the total pressure compared to the ramp injection alone, i.e., the baseline case. The numerical results clearly indicate that the fuel injection schemes investigated in the present study are more efficient than a strut or multi-staged strut and wall injection scheme.
Advisors/Committee Members: Farokhi, Saeed (advisor), Taghavi, Ray (cmtemember), Barrett-Gonzalez, Ron (cmtemember), Keshmiri, Shahriar (cmtemember), Huang, Weizhang (cmtemember).
Subjects/Keywords: Aerospace engineering; Combustor; Enhanced; Mixing; Scramjet; Suction collar; Tabs
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APA ·
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Export
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APA (6th Edition):
Hwang, S. (2011). Numerical Simulation of Enhanced Mixing in Scramjet Combustor Using Ramp, Tabs and Suction Collar. (Doctoral Dissertation). University of Kansas. Retrieved from http://hdl.handle.net/1808/8080
Chicago Manual of Style (16th Edition):
Hwang, Seung-Jae. “Numerical Simulation of Enhanced Mixing in Scramjet Combustor Using Ramp, Tabs and Suction Collar.” 2011. Doctoral Dissertation, University of Kansas. Accessed January 24, 2021.
http://hdl.handle.net/1808/8080.
MLA Handbook (7th Edition):
Hwang, Seung-Jae. “Numerical Simulation of Enhanced Mixing in Scramjet Combustor Using Ramp, Tabs and Suction Collar.” 2011. Web. 24 Jan 2021.
Vancouver:
Hwang S. Numerical Simulation of Enhanced Mixing in Scramjet Combustor Using Ramp, Tabs and Suction Collar. [Internet] [Doctoral dissertation]. University of Kansas; 2011. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/1808/8080.
Council of Science Editors:
Hwang S. Numerical Simulation of Enhanced Mixing in Scramjet Combustor Using Ramp, Tabs and Suction Collar. [Doctoral Dissertation]. University of Kansas; 2011. Available from: http://hdl.handle.net/1808/8080
University of Cambridge
20.
Cocks, Peter.
Large eddy simulation of supersonic combustion with application to scramjet engines.
Degree: PhD, 2011, University of Cambridge
URL: https://doi.org/10.17863/CAM.14003
;
https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.541905
► This work evaluates the capabilities of the RANS and LES techniques for the simulation of high speed reacting flows. These methods are used to gain…
(more)
▼ This work evaluates the capabilities of the RANS and LES techniques for the simulation of high speed reacting flows. These methods are used to gain further insight into the physics encountered and regimes present in supersonic combustion. The target application of this research is the scramjet engine, a propulsion system of great promise for efficient hypersonic flight. In order to conduct this work a new highly parallelised code, PULSAR, is developed. PULSAR is capable of simulating complex chemistry combustion in highly compressible flows, based on a second order upwind method to provide a monotonic solution in the presence of high gradient physics. Through the simulation of a non-reacting supersonic coaxial helium jet the RANS method is shown to be sensitive to constants involved in the modelling process. The LES technique is more computationally demanding but is shown to be much less sensitive to these model parameters. Nevertheless, LES results are shown to be sensitive to the nature of turbulence at the inflow; however this information can be experimentally obtained. The SCHOLAR test case is used to validate the reacting aspects of PULSAR. Comparing RANS results from laminar chemistry and assumed PDF combustion model simulations, the influence of turbulence-chemistry interactions in supersonic combustion is shown to be small. In the presence of reactions, the RANS results are sensitive to inflow turbulence, due to its influence on mixing. From complex chemistry simulations the combustion behaviour is evaluated to sit between the flamelet and distributed reaction regimes. LES results allow an evaluation of the physics involved, with a pair of coherent vortices identified as the dominant influence on mixing for the oblique wall fuel injection method. It is shown that inflow turbulence has a significant impact on the behaviour of these vortices and hence it is vital for turbulence intensities and length scales to be measured by experimentalists, in order for accurate simulations to be possible.
Subjects/Keywords: 620; Scramjet; Supersonic combustion; Assumed pdf combustion model; Large eddy simulation
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APA ·
Chicago ·
MLA ·
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CSE |
Export
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Manager
APA (6th Edition):
Cocks, P. (2011). Large eddy simulation of supersonic combustion with application to scramjet engines. (Doctoral Dissertation). University of Cambridge. Retrieved from https://doi.org/10.17863/CAM.14003 ; https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.541905
Chicago Manual of Style (16th Edition):
Cocks, Peter. “Large eddy simulation of supersonic combustion with application to scramjet engines.” 2011. Doctoral Dissertation, University of Cambridge. Accessed January 24, 2021.
https://doi.org/10.17863/CAM.14003 ; https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.541905.
MLA Handbook (7th Edition):
Cocks, Peter. “Large eddy simulation of supersonic combustion with application to scramjet engines.” 2011. Web. 24 Jan 2021.
Vancouver:
Cocks P. Large eddy simulation of supersonic combustion with application to scramjet engines. [Internet] [Doctoral dissertation]. University of Cambridge; 2011. [cited 2021 Jan 24].
Available from: https://doi.org/10.17863/CAM.14003 ; https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.541905.
Council of Science Editors:
Cocks P. Large eddy simulation of supersonic combustion with application to scramjet engines. [Doctoral Dissertation]. University of Cambridge; 2011. Available from: https://doi.org/10.17863/CAM.14003 ; https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.541905
University of Cambridge
21.
Cocks, Peter.
Large eddy simulation of supersonic combustion with application to scramjet engines.
Degree: PhD, 2011, University of Cambridge
URL: http://www.dspace.cam.ac.uk/handle/1810/239344https://www.repository.cam.ac.uk/bitstream/1810/239344/2/license.txt
;
https://www.repository.cam.ac.uk/bitstream/1810/239344/5/thesis.pdf.txt
;
https://www.repository.cam.ac.uk/bitstream/1810/239344/6/thesis.pdf.jpg
► This work evaluates the capabilities of the RANS and LES techniques for the simulation of high speed reacting flows. These methods are used to gain…
(more)
▼ This work evaluates the capabilities of the RANS and LES techniques for the simulation of high speed reacting flows. These methods are used to gain further insight into the physics encountered and regimes present in supersonic combustion. The target application of this research is the scramjet engine, a propulsion system of great promise for efficient hypersonic flight. In order to conduct this work a new highly parallelised code, PULSAR, is developed. PULSAR is capable of simulating complex chemistry combustion in highly compressible flows, based on a second order upwind method to provide a monotonic solution in the presence of high gradient physics.
Through the simulation of a non-reacting supersonic coaxial helium jet the RANS method is shown to be sensitive to constants involved in the modelling process. The LES technique is more computationally demanding but is shown to be much less sensitive to these model parameters. Nevertheless, LES results are shown to be sensitive to the nature of turbulence at the inflow; however this information can be experimentally obtained.
The SCHOLAR test case is used to validate the reacting aspects of PULSAR. Comparing RANS results from laminar chemistry and assumed PDF combustion model simulations, the influence of turbulence-chemistry interactions in supersonic combustion is shown to be small. In the presence of reactions, the RANS results are sensitive to inflow turbulence, due to its influence on mixing. From complex chemistry simulations the combustion behaviour is evaluated to sit between the flamelet and distributed reaction regimes. LES results allow an evaluation of the physics involved, with a pair of coherent vortices identified as the dominant influence on mixing for the oblique wall fuel injection method. It is shown that inflow turbulence has a significant impact on the behaviour of these vortices and hence it is vital for turbulence intensities and length scales to be measured by experimentalists, in order for accurate simulations to be possible.
Subjects/Keywords: Scramjet; Supersonic combustion; Assumed pdf combustion model; Large eddy simulation
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❌
APA ·
Chicago ·
MLA ·
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CSE |
Export
to Zotero / EndNote / Reference
Manager
APA (6th Edition):
Cocks, P. (2011). Large eddy simulation of supersonic combustion with application to scramjet engines. (Doctoral Dissertation). University of Cambridge. Retrieved from http://www.dspace.cam.ac.uk/handle/1810/239344https://www.repository.cam.ac.uk/bitstream/1810/239344/2/license.txt ; https://www.repository.cam.ac.uk/bitstream/1810/239344/5/thesis.pdf.txt ; https://www.repository.cam.ac.uk/bitstream/1810/239344/6/thesis.pdf.jpg
Chicago Manual of Style (16th Edition):
Cocks, Peter. “Large eddy simulation of supersonic combustion with application to scramjet engines.” 2011. Doctoral Dissertation, University of Cambridge. Accessed January 24, 2021.
http://www.dspace.cam.ac.uk/handle/1810/239344https://www.repository.cam.ac.uk/bitstream/1810/239344/2/license.txt ; https://www.repository.cam.ac.uk/bitstream/1810/239344/5/thesis.pdf.txt ; https://www.repository.cam.ac.uk/bitstream/1810/239344/6/thesis.pdf.jpg.
MLA Handbook (7th Edition):
Cocks, Peter. “Large eddy simulation of supersonic combustion with application to scramjet engines.” 2011. Web. 24 Jan 2021.
Vancouver:
Cocks P. Large eddy simulation of supersonic combustion with application to scramjet engines. [Internet] [Doctoral dissertation]. University of Cambridge; 2011. [cited 2021 Jan 24].
Available from: http://www.dspace.cam.ac.uk/handle/1810/239344https://www.repository.cam.ac.uk/bitstream/1810/239344/2/license.txt ; https://www.repository.cam.ac.uk/bitstream/1810/239344/5/thesis.pdf.txt ; https://www.repository.cam.ac.uk/bitstream/1810/239344/6/thesis.pdf.jpg.
Council of Science Editors:
Cocks P. Large eddy simulation of supersonic combustion with application to scramjet engines. [Doctoral Dissertation]. University of Cambridge; 2011. Available from: http://www.dspace.cam.ac.uk/handle/1810/239344https://www.repository.cam.ac.uk/bitstream/1810/239344/2/license.txt ; https://www.repository.cam.ac.uk/bitstream/1810/239344/5/thesis.pdf.txt ; https://www.repository.cam.ac.uk/bitstream/1810/239344/6/thesis.pdf.jpg
22.
Choi, Jongseong.
Parametric Scramjet Analysis.
Degree: M.S. in Engineering Science, Mechanical Engineering, 2014, University of Mississippi
URL: https://egrove.olemiss.edu/etd/535
► The performance of a hypersonic flight vehicle will depend on existing materials and fuels; this work presents the performance of the ideal scramjet engine for…
(more)
▼ The performance of a hypersonic flight vehicle will depend on existing materials and fuels; this work presents the performance of the ideal
scramjet engine for three different combustion chamber materials and three different candidate fuels. Engine performance is explored by parametric cycle analysis for the ideal
scramjet as a function of material maximum service temperature and the lower heating value of jet engine fuels. The thermodynamic analysis is based on the Brayton cycle as similarly employed in describing the performance of the ramjet, turbojet, and fanjet ideal engines. The objective of this work is to explore material operating temperatures and fuel possibilities for the combustion chamber of a
scramjet propulsion system to show how they relate to
scramjet performance and the seven
scramjet engine parameters: specific thrust, fuel-to-air ratio, thrust-specific fuel consumption, thermal efficiency, propulsive efficiency, overall efficiency, and thrust flux. The information presented in this work has not been done by others in the scientific literature. This work yields simple algebraic equations for
scramjet performance which are similar to that of the ideal ramjet, ideal turbojet and ideal turbofan engines.
Advisors/Committee Members: Jeffrey A. Roux.
Subjects/Keywords: Efficiency; Heat Transfer; Mass Flow Rate; Parametric; Ramjet; Scramjet; Aerospace Engineering
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APA ·
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Export
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APA (6th Edition):
Choi, J. (2014). Parametric Scramjet Analysis. (Thesis). University of Mississippi. Retrieved from https://egrove.olemiss.edu/etd/535
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):
Choi, Jongseong. “Parametric Scramjet Analysis.” 2014. Thesis, University of Mississippi. Accessed January 24, 2021.
https://egrove.olemiss.edu/etd/535.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Choi, Jongseong. “Parametric Scramjet Analysis.” 2014. Web. 24 Jan 2021.
Vancouver:
Choi J. Parametric Scramjet Analysis. [Internet] [Thesis]. University of Mississippi; 2014. [cited 2021 Jan 24].
Available from: https://egrove.olemiss.edu/etd/535.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Choi J. Parametric Scramjet Analysis. [Thesis]. University of Mississippi; 2014. Available from: https://egrove.olemiss.edu/etd/535
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of Notre Dame
23.
Qili Liu.
Turbulent Combustion in High Mach Number
Flows</h1>.
Degree: Aerospace and Mechanical Engineering, 2015, University of Notre Dame
URL: https://curate.nd.edu/show/zc77sn02w60
► The primary goals of this study are to investigate turbulent combustion dynamics in high Mach number flows and to understand the influence of turbulence…
(more)
▼ The primary goals of this study are to
investigate turbulent combustion dynamics in high Mach number flows
and to understand the influence of turbulence on the high-speed
combustion phenomena. To conduct this unprecedented experimental
investigation, an arc-heated hypersonic wind tunnel facility with
1-second test time is used for generating high Mach number (Mach
4.5, 6 and 9) and high-enthalpy flows of total temperature up to
3,500 K. An integrated model
scramjet is installed in the test
section of the hypersonic tunnel for experimental investigations in
well-defined flows. The freestream turbulence is
manipulated using various combinations of meshes and a turbulence
damper (a component enforcing slow flow expansion and rapid
contraction). It is well known that fine meshes can break
large-scale flow structures into small-scale structures to
facilitate turbulence dissipation. In addition, the contraction
component (turbulence damper) is designed to accelerate gas
mixtures toward sonic point at the C/D nozzle throat while
dampening the streamwise velocity fluctuations. The turbulence
levels in flows are measured qualitatively and quantitatively using
Rayleigh scattering and Pitot probe, respectively.
Flow field visualizations in a model
scramjet are
conducted in room temperature flows at different turbulence levels
using CO2-seeded Rayleigh scattering method. The three-dimensional
flow structures are acquired by scanning the two-dimensional laser
sheet for Rayleigh scattering over the entire
scramjet flow volume.
Flow structures visualized using this novel technique depict that
the jet-induced secondary flows significantly affect the subsequent
flow structures. In addition, it is obvious that the increased
turbulence level in freestream triggers earlier boundary layer
transition and separation. Although these investigations on the 3D
flow structures in the model
scramjet are conducted in low-enthalpy
flows, numerous important insights on the rather generic flow
behaviors that will appear also in high-enthalpy flows are provided
with various turbulent levels and inlet geometries.
Planar laser induced fluorescence (PLIF) method is used as
the most important tool for investigating the turbulent flame in
supersonic flows. Ethylene (C2H4) is directly injected into the
model
scramjet under various freestream turbulence conditions. The
effects of the freestream turbulence, inlet geometries, and other
flow properties including total pressure/temperature are
studied.
Advisors/Committee Members: Flint O. Thomas, Committee Member, Hyungrok Do, Committee Chair, Scott C. Morris, Committee Member, Joseph M. Powers, Committee Member.
Subjects/Keywords: scramjet; hypersonic flows; flow visualization; ethylene flame; turbulence control; combustion diagnostics
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APA ·
Chicago ·
MLA ·
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CSE |
Export
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Manager
APA (6th Edition):
Liu, Q. (2015). Turbulent Combustion in High Mach Number
Flows</h1>. (Thesis). University of Notre Dame. Retrieved from https://curate.nd.edu/show/zc77sn02w60
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):
Liu, Qili. “Turbulent Combustion in High Mach Number
Flows</h1>.” 2015. Thesis, University of Notre Dame. Accessed January 24, 2021.
https://curate.nd.edu/show/zc77sn02w60.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Liu, Qili. “Turbulent Combustion in High Mach Number
Flows</h1>.” 2015. Web. 24 Jan 2021.
Vancouver:
Liu Q. Turbulent Combustion in High Mach Number
Flows</h1>. [Internet] [Thesis]. University of Notre Dame; 2015. [cited 2021 Jan 24].
Available from: https://curate.nd.edu/show/zc77sn02w60.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Liu Q. Turbulent Combustion in High Mach Number
Flows</h1>. [Thesis]. University of Notre Dame; 2015. Available from: https://curate.nd.edu/show/zc77sn02w60
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
University of Minnesota
24.
Peterson, David Michael.
Simulations of injection, mixing, and combustion in supersonic flow using a hybrid RANS/LES approach.
Degree: PhD, Aerospace Engineering and Mechanics, 2011, University of Minnesota
URL: http://purl.umn.edu/117465
► There is a great need for accurate and reliable numerical simulation of injection, mixing, and combustion in supersonic combustion ramjet engines. This study seeks to…
(more)
▼ There is a great need for accurate and reliable numerical simulation of injection, mixing, and combustion in supersonic combustion ramjet engines. This study seeks to improve the accuracy and reliability which these flow can be simulated with by investigating the use of recent improvements in turbulence modeling and numerical methods. The present numerical simulations use implicit time integration and low-dissipation flux evaluation schemes in an unstructured grid framework. A hybrid Reynolds-Averaged Navier-Stokes and large-eddy simulation approach is used to model turbulence. The large-scale turbulent structure of the flow is resolved, while the near-wall structure is fully modeled. The effects of numerics, grid resolution, and boundary conditions are investigated. The simulation approach is thoroughly validated against available experimental data at a variety of flow conditions. The simulations focus on the injection of fuel through circular injector ports that are oriented either normal to the supersonic crossflow, or at a low angle with respect to the crossflow. The instantaneous flow structure resolved by the simulations is qualitatively compared to experimental flowfield visualization. Quantitative comparisons are made to mean wall pressure, mean velocity, turbulence quantities, and mean mixing data. The simulations are found to do very well at predicting the mean flowfield as well as fluctuations in velocity and injectant concentration. The simulation approach is then used to simulate the flow within a model supersonic combustor. The focus is on the non-reacting case. The simulation results are found to agree well with experimental measurements of temperature and species concentrations. The flow is examined to improve understanding of the mixing within the model combustor. Preliminary results for a simulation including hydrogen combustion are also presented.
Subjects/Keywords: Combustion; Hypersonic; LES; Mixing; Scramjet; Supersonic; Aerospace Engineering and Mechanics
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APA (6th Edition):
Peterson, D. M. (2011). Simulations of injection, mixing, and combustion in supersonic flow using a hybrid RANS/LES approach. (Doctoral Dissertation). University of Minnesota. Retrieved from http://purl.umn.edu/117465
Chicago Manual of Style (16th Edition):
Peterson, David Michael. “Simulations of injection, mixing, and combustion in supersonic flow using a hybrid RANS/LES approach.” 2011. Doctoral Dissertation, University of Minnesota. Accessed January 24, 2021.
http://purl.umn.edu/117465.
MLA Handbook (7th Edition):
Peterson, David Michael. “Simulations of injection, mixing, and combustion in supersonic flow using a hybrid RANS/LES approach.” 2011. Web. 24 Jan 2021.
Vancouver:
Peterson DM. Simulations of injection, mixing, and combustion in supersonic flow using a hybrid RANS/LES approach. [Internet] [Doctoral dissertation]. University of Minnesota; 2011. [cited 2021 Jan 24].
Available from: http://purl.umn.edu/117465.
Council of Science Editors:
Peterson DM. Simulations of injection, mixing, and combustion in supersonic flow using a hybrid RANS/LES approach. [Doctoral Dissertation]. University of Minnesota; 2011. Available from: http://purl.umn.edu/117465
University of Kansas
25.
Smith, Leslie Ann.
The Effect of Pulsed Injection on Shear Layer Dynamics in a Scramjet Combustion Chamber.
Degree: PhD, Aerospace Engineering, 2015, University of Kansas
URL: http://hdl.handle.net/1808/19415
► One of the greatest problems that scramjet research faces is fuel air mixing. The residence time for a scramjet engine, or the time it takes…
(more)
▼ One of the greatest problems that
scramjet research faces is fuel air mixing. The residence time for a
scramjet engine, or the time it takes for a volume of air to completely pass through the engine, is on the order of 0.1 ms. In that extremely short period of time fuel must be injected and fully mirco-mixed at stoichiometric ratios with the combustion chamber airflow. The fuel-air mixture must then be combusted and expanded through the nozzle to produce thrust. The goal of this research is to develop a new more efficient method of fuel air mixing within a
scramjet combustion chamber. A possible way to speed up the mixing process of parallel injection without incurring the total pressure losses that would occur in normal injection is to inject the fuel from the rear side of a backward facing step. Backward facing steps in supersonic flow produce a Prandtl-Meyer expansion fan followed by a shear layer. The instabilities in this shear layer have dominant resonant frequencies. It is believed that if fuel is injected in pulses that impinge on the shear layer at these dominant resonant frequencies that the shear layer will resonate. When the shear layer resonates the vortices that form in the shear layer will grow in magnitude, thus mixing the injected fuel with the air. To test this hypothesis a new test section was designed and built that features a one inch step under which an injector can be housed. This new test section was installed in the supersonic facility at the University of Kansas. Two injectors were also designed that each feature a face plate, one with eight injection ports arranged in a ring and one with 5 injection ports. Between the face plate and a back plate there is a cavity that houses a rotating valve that is powered by a pneumatic motor. Five valves were built: one with 8 teeth, one with 16 teeth, one with 5 teeth that are the same size as the gaps between the teeth, one with 5 teeth where the teeth are 50% larger than the gaps, and one with 5 teeth where the teeth are 50% smaller than the gaps. The 8 tooth valve and 16 tooth valve where used with the 8 port injector face plate. The 5 tooth valves were used with the 5 port injector face plate. As the valve rotates the teeth block and unblock the injection ports injecting carbon dioxide gas into the test section. The 8 port injector was tested over a range of frequencies from 1.6 kHz to 10.0 kHz. The 5 port injector was tested for each valve over a range of frequencies from 1.0 kHz to 4.0 kHz. Static pressure data was taken along the upper and lower walls of the test section by means of an array of pressure sensors. The pressure data from the test section was compared to results generated using a three dimensional CFD simulation of the test section. Overall the pressure data on the lower wall agreed reasonably well with the CFD simulation. The vorticity and turbulence contours generated by the STAR-CCM+ simulation suggest that as a pulse is injected into the test section from the step it causes the shear layer to curve outward near the point of…
Advisors/Committee Members: Farokhi, Saeed (advisor), Barrett, Ron (cmtemember), Taghavi, Ray (cmtemember), Yimer, Bedru (cmtemember), Keshmiri, Shawn (cmtemember).
Subjects/Keywords: Aerospace engineering; Injector; Micro-mixing; Pulsed Injection; Scramjet; Shear Layer
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APA ·
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APA (6th Edition):
Smith, L. A. (2015). The Effect of Pulsed Injection on Shear Layer Dynamics in a Scramjet Combustion Chamber. (Doctoral Dissertation). University of Kansas. Retrieved from http://hdl.handle.net/1808/19415
Chicago Manual of Style (16th Edition):
Smith, Leslie Ann. “The Effect of Pulsed Injection on Shear Layer Dynamics in a Scramjet Combustion Chamber.” 2015. Doctoral Dissertation, University of Kansas. Accessed January 24, 2021.
http://hdl.handle.net/1808/19415.
MLA Handbook (7th Edition):
Smith, Leslie Ann. “The Effect of Pulsed Injection on Shear Layer Dynamics in a Scramjet Combustion Chamber.” 2015. Web. 24 Jan 2021.
Vancouver:
Smith LA. The Effect of Pulsed Injection on Shear Layer Dynamics in a Scramjet Combustion Chamber. [Internet] [Doctoral dissertation]. University of Kansas; 2015. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/1808/19415.
Council of Science Editors:
Smith LA. The Effect of Pulsed Injection on Shear Layer Dynamics in a Scramjet Combustion Chamber. [Doctoral Dissertation]. University of Kansas; 2015. Available from: http://hdl.handle.net/1808/19415
University of Kansas
26.
Flesberg, Sonja.
Mixing Enhancement in a Scramjet Combustor Using Fuel Jet Injection Swirl.
Degree: PhD, Aerospace Engineering, 2015, University of Kansas
URL: http://hdl.handle.net/1808/19380
► The scramjet engine has proven to be a viable means of powering a hypersonic vehicle, especially after successful flights of the X-51 WaveRider and various…
(more)
▼ The
scramjet engine has proven to be a viable means of powering a hypersonic vehicle, especially after successful flights of the X-51 WaveRider and various Hy-SHOT test vehicles. The major challenge associated with operating a
scramjet engine is the short residence time of the fuel and oxidizer in the combustor. The fuel and oxidizer have only milliseconds to mix, ignite and combust in the combustion chamber. Combustion cannot occur until the fuel and oxidizer are mixed on a molecular level. Therefore the improvement of mixing is of utmost interest since this can increase combustion efficiency. This study investigated mixing enhancement of fuel and oxidizer within the combustion chamber of a
scramjet by introducing swirl to the fuel jet. The investigation was accomplished with numerical simulations using STAR-CCM+ computational fluid dynamic software. The geometry of the University of Virginia Supersonic Combustion Facility was used to model the isolator, combustor and nozzle of a
scramjet engine for simulation purposes. Experimental data from previous research at the facility was used to verify the simulation model before investigating the effect of fuel jet swirl on mixing. The model used coaxial fuel jet with a swirling annular jet. Single coaxial fuel jet and dual coaxial fuel jet configurations were simulated for the investigation. The coaxial fuel jets were modelled with a swirling annular jet and non-swirling core jet. Numerical analysis showed that fuel jet swirl not only increased mixing and entrainment of the fuel with the oxidizer but the mixing occurred further upstream than without fuel jet swirl. The burning efficiency was calculated for the all the configurations. An increase in burning efficiency indicated an increase in the mixing of H2 with O2. In the case of the single fuel jet models, the maximum burning efficiency increase due to fuel injection jet swirl was 23.3%. The research also investigated the possibility that interaction between two swirling jets would produce increased mixing and to study how the distance between the two fuel injector exits would affect mixing. Three swirl patterns were investigated: 1) the first swirl pattern as viewed by an observer looking downstream had the right fuel annular jet swirling counter clockwise and the left fuel annular jet swirling clockwise, 2) the second swirl pattern as viewed by an observer looking downstream had the right fuel jet swirling clockwise and the left fuel jet swirling counter clockwise, 3) the third swirl pattern as viewed by an observer looking downstream had both the right and left fuel jet swirling in the same clockwise direction. Each one of the swirl patterns were simulated with the distances between the center points of the fuel jets modelled 3, 4, and 5 times the fuel injector radius. The swirl pattern that produced the greatest increase in burning efficiency differed according to the fuel injector spacing. The maximum increase in burning efficiency compared to the corresponding non-swirling two jet baseline case was 24.6% and was…
Advisors/Committee Members: Taghavi, Ray (advisor), Farokhi, Saeed (advisor), Taghavi, Ray (cmtemember), Farokhi, Saeed (cmtemember), Hale, Richard (cmtemember), Barrett, Ronald (cmtemember), Yimer, Bedru (cmtemember).
Subjects/Keywords: Aerospace engineering; fuel jet injection swirl; mixing enhancement; scramjet combustion
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APA (6th Edition):
Flesberg, S. (2015). Mixing Enhancement in a Scramjet Combustor Using Fuel Jet Injection Swirl. (Doctoral Dissertation). University of Kansas. Retrieved from http://hdl.handle.net/1808/19380
Chicago Manual of Style (16th Edition):
Flesberg, Sonja. “Mixing Enhancement in a Scramjet Combustor Using Fuel Jet Injection Swirl.” 2015. Doctoral Dissertation, University of Kansas. Accessed January 24, 2021.
http://hdl.handle.net/1808/19380.
MLA Handbook (7th Edition):
Flesberg, Sonja. “Mixing Enhancement in a Scramjet Combustor Using Fuel Jet Injection Swirl.” 2015. Web. 24 Jan 2021.
Vancouver:
Flesberg S. Mixing Enhancement in a Scramjet Combustor Using Fuel Jet Injection Swirl. [Internet] [Doctoral dissertation]. University of Kansas; 2015. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/1808/19380.
Council of Science Editors:
Flesberg S. Mixing Enhancement in a Scramjet Combustor Using Fuel Jet Injection Swirl. [Doctoral Dissertation]. University of Kansas; 2015. Available from: http://hdl.handle.net/1808/19380
University of Maryland
27.
Bussey, Gillian Mary Harding.
ENTROPY CONSIDERATIONS APPLIED TO SHOCK UNSTEADINESS IN HYPERSONIC INLETS.
Degree: Aerospace Engineering, 2012, University of Maryland
URL: http://hdl.handle.net/1903/13503
► The stability of curved or rectangular shocks in hypersonic inlets in reponse to flow perturbations can be determined analytically from the principle of minimum entropy.…
(more)
▼ The stability of curved or rectangular shocks in hypersonic inlets in reponse to flow perturbations can be determined analytically from the principle of minimum entropy. Unsteady shock wave motion can have a significant effect on the flow in a hypersonic inlet or combustor. According to the principle of minimum entropy, a stable thermodynamic state is one with the lowest entropy gain. A model based on piston theory and its limits has been developed for applying the principle of minimum entropy to quasi-steady flow. Relations are derived for analyzing the time-averaged entropy gain flux across a shock for quasi-steady perturbations in atmospheric conditions and angle as a perturbation in entropy gain flux from the steady state. Initial results from sweeping a wedge at Mach 10 through several degrees in AEDC's Tunnel 9 indicates the bow shock becomes unsteady near the predicted normal Mach number. Several curved shocks of varying curvature are compared to a straight shock with the same mean normal Mach number, pressure ratio, or temperature ratio. The present work provides analysis and guidelines for designing an inlet robust to off- design flight or perturbations in flow conditions an inlet is likely to face. It also suggests that inlets with curved shocks are less robust to off-design flight than those with straight shocks such as rectangular inlets. Relations for evaluating entropy perturbations for highly unsteady flow across a shock and limits on their use were also developed. The normal Mach number at which a shock could be stable to high frequency upstream perturbations increases as the speed of the shock motion increases and slightly decreases as the perturbation size increases. The present work advances the principle of minimum entropy theory by providing additional validity for using the theory for time-varying flows and applying it to shocks, specifically those in inlets. While this analytic tool is applied in the present work for evaluating the stability of shocks in hypersonic inlets, it can be used for an arbitrary application with a shock.
Advisors/Committee Members: Lewis, Mark J (advisor).
Subjects/Keywords: Aerospace engineering; Physics; Entropy; Hypersonic inlets; Hypersonics; Scramjet; Shock waves; Thermodynamics
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APA (6th Edition):
Bussey, G. M. H. (2012). ENTROPY CONSIDERATIONS APPLIED TO SHOCK UNSTEADINESS IN HYPERSONIC INLETS. (Thesis). University of Maryland. Retrieved from http://hdl.handle.net/1903/13503
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):
Bussey, Gillian Mary Harding. “ENTROPY CONSIDERATIONS APPLIED TO SHOCK UNSTEADINESS IN HYPERSONIC INLETS.” 2012. Thesis, University of Maryland. Accessed January 24, 2021.
http://hdl.handle.net/1903/13503.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
MLA Handbook (7th Edition):
Bussey, Gillian Mary Harding. “ENTROPY CONSIDERATIONS APPLIED TO SHOCK UNSTEADINESS IN HYPERSONIC INLETS.” 2012. Web. 24 Jan 2021.
Vancouver:
Bussey GMH. ENTROPY CONSIDERATIONS APPLIED TO SHOCK UNSTEADINESS IN HYPERSONIC INLETS. [Internet] [Thesis]. University of Maryland; 2012. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/1903/13503.
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Council of Science Editors:
Bussey GMH. ENTROPY CONSIDERATIONS APPLIED TO SHOCK UNSTEADINESS IN HYPERSONIC INLETS. [Thesis]. University of Maryland; 2012. Available from: http://hdl.handle.net/1903/13503
Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation
Virginia Tech
28.
Cox-Stouffer, Susan K. Jr.
Numerical Simulation of Injection and Mixing in Supersonic Flow.
Degree: PhD, Aerospace and Ocean Engineering, 1997, Virginia Tech
URL: http://hdl.handle.net/10919/29628
► A numerical investigation of the performance of two candidate designs for injection into supersonic flow, including a comparison of two renormalized group theory (RNG) based…
(more)
▼ A numerical investigation of the performance of two candidate designs for injection into supersonic flow, including a comparison of two renormalized group theory (RNG) based k-epsilon turbulence models with a more conventional k-epsilon model. The chosen designs were an unswept ramp injector with four injection ports and a novel nine-hole injector array. The objectives of the investigation were to provide reliable computational solutions to the flowfields in question using both RNG and standard k-epsilon turbulence models and to compare the solutions to experiment, thereby to judge the relative performance of the turbulence models. A second objective of the investigation was to use the computed data to provide design insights for the nine-hole injector array.
This investigation made use of GASP(tm) version 2.2, a commercial computational fluid dynamics code that was augmented by the addition of one RNG-based k-epsilon turbulence model derived by Zhou, et. al. and one variant of Zhou's model, which was derived by the author. Mesh sequencing studies were performed to measure solution quality, with the fine mesh for the injector array containing roughly one million grid nodes and the fine mesh for the ramp injector containing more than six million grid nodes. Results of these studies indicated that the injector-array solution was significantly under-resolved in the farfield, though the quality was better in the vicinity of the injector itself. The ramp-injector solution, while not perfectly grid-resolved, showed much better grid convergence in both the nearfield and farfield. Accordingly, comparison with experiment was better for the ramp injector than for the injector array. For both injectors, the differences between solutions generated with RNG-based k-epsilon and standard k-epsilon turbulence models were negligibly small."
Despite inadequate grid resolution in the farfield, the computational investigation of the nine-hole injector array did yield several important design insights. Particularly, the significance to mixing and losses of the placement of the outer injectors of the second and third rows was determined.
Advisors/Committee Members: Schetz, Joseph A. (committeechair), Grossman, Bernard M. (committee member), Ng, Wing Fai (committee member), Walters, Robert W. (committee member), Yates, Charlie L. (committee member).
Subjects/Keywords: scramjet; RNG; CFD; ramp injector
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APA (6th Edition):
Cox-Stouffer, S. K. J. (1997). Numerical Simulation of Injection and Mixing in Supersonic Flow. (Doctoral Dissertation). Virginia Tech. Retrieved from http://hdl.handle.net/10919/29628
Chicago Manual of Style (16th Edition):
Cox-Stouffer, Susan K Jr. “Numerical Simulation of Injection and Mixing in Supersonic Flow.” 1997. Doctoral Dissertation, Virginia Tech. Accessed January 24, 2021.
http://hdl.handle.net/10919/29628.
MLA Handbook (7th Edition):
Cox-Stouffer, Susan K Jr. “Numerical Simulation of Injection and Mixing in Supersonic Flow.” 1997. Web. 24 Jan 2021.
Vancouver:
Cox-Stouffer SKJ. Numerical Simulation of Injection and Mixing in Supersonic Flow. [Internet] [Doctoral dissertation]. Virginia Tech; 1997. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/10919/29628.
Council of Science Editors:
Cox-Stouffer SKJ. Numerical Simulation of Injection and Mixing in Supersonic Flow. [Doctoral Dissertation]. Virginia Tech; 1997. Available from: http://hdl.handle.net/10919/29628
Virginia Tech
29.
Burger, Scott Kuhlman.
Investigation of Injectant Molecular Weight and Shock Impingement Effects on Transverse Injection Mixing in Supersonic Flow.
Degree: MS, Aerospace and Ocean Engineering, 2010, Virginia Tech
URL: http://hdl.handle.net/10919/31981
► This study examines the effect of varying injectant molecular weight on the penetration of transverse injection jets into a supersonic crossflow. The injectants considered here…
(more)
▼ This study examines the effect of varying injectant molecular weight on the penetration of transverse injection jets into a supersonic crossflow. The injectants considered here are methane (W=16.04), air (W=28.97) and carbon dioxide, (W=44.01). These results augment the previous results obtained at Virginia Tech for helium (W=4.00) injection under the same test conditions to provide a very wide range of molecular weights. Second, since shocks are ubiquitous in
scramjet combustors, their influence on penetration and mixing was also studied by arranging for an oblique shock to impinge near the injection station. The cases of a shock impinging upstream and downstream of the injector were both examined. One can anticipate an important influence of molecular weight here also because of the importance of density gradients on the generation of vorticity by baroclinic torque. Increasing molecular weight was found to increase penetration in general, as well as increase the lateral spreading of the plume. The majority of the data shows a weak dependency of the jet size on molecular weight, but there are indications that under certain circumstances large changes in the flow structure may occur due to molecular weight effects. The addition of an impinging shock is found to increase mixing and decrease penetration and plume size, especially with the shock impinging downstream of the injector.
Advisors/Committee Members: Schetz, Joseph A. (committeechair), Kochersberger, Kevin Bruce (committee member), Devenport, William J. (committee member).
Subjects/Keywords: molecular weight; mixing; injection; scramjet
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APA (6th Edition):
Burger, S. K. (2010). Investigation of Injectant Molecular Weight and Shock Impingement Effects on Transverse Injection Mixing in Supersonic Flow. (Masters Thesis). Virginia Tech. Retrieved from http://hdl.handle.net/10919/31981
Chicago Manual of Style (16th Edition):
Burger, Scott Kuhlman. “Investigation of Injectant Molecular Weight and Shock Impingement Effects on Transverse Injection Mixing in Supersonic Flow.” 2010. Masters Thesis, Virginia Tech. Accessed January 24, 2021.
http://hdl.handle.net/10919/31981.
MLA Handbook (7th Edition):
Burger, Scott Kuhlman. “Investigation of Injectant Molecular Weight and Shock Impingement Effects on Transverse Injection Mixing in Supersonic Flow.” 2010. Web. 24 Jan 2021.
Vancouver:
Burger SK. Investigation of Injectant Molecular Weight and Shock Impingement Effects on Transverse Injection Mixing in Supersonic Flow. [Internet] [Masters thesis]. Virginia Tech; 2010. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/10919/31981.
Council of Science Editors:
Burger SK. Investigation of Injectant Molecular Weight and Shock Impingement Effects on Transverse Injection Mixing in Supersonic Flow. [Masters Thesis]. Virginia Tech; 2010. Available from: http://hdl.handle.net/10919/31981
Virginia Tech
30.
Markell, Kyle Charles.
Exergy Methods for the Generic Analysis and Optimization of Hypersonic Vehicle Concepts.
Degree: MS, Mechanical Engineering, 2005, Virginia Tech
URL: http://hdl.handle.net/10919/31256
► This thesis work presents detailed results of the application of exergy-based methods to highly dynamic, integrated aerospace systems such as hypersonic vehicle concepts. In particular,…
(more)
▼ This thesis work presents detailed results of the application of exergy-based methods to highly dynamic, integrated aerospace systems such as hypersonic vehicle concepts. In particular, an exergy-based methodology is compared to a more traditional based measure by applying both to the synthesis/design and operational optimization of a hypersonic vehicle configuration comprised of an airframe sub-system and a propulsion sub-system consisting of inlet, combustor, and nozzle components. A number of key design and operational decision variables are identified as those which govern the hypersonic vehicle flow physics and thermodynamics and detailed one-dimensional models of each component and sub-system are developed. Rates of exergy loss as well as exergy destruction resulting from irreversible loss mechanisms are determined in each of the hypersonic vehicle sub-systems and their respective components.
Multiple optimizations are performed for both the propulsion sub-system only and for the entire hypersonic vehicle system for single mission segments and for a partial, three-segment mission. Three different objective functions are utilized in these optimizations with the specific goal of comparing exergy methods to a standard vehicle performance measure, namely, the vehicle overall efficiency. Results of these optimizations show that the exergy method presented here performs well when compared to the standard performance measure and, in a number of cases, leads to more optimal syntheses/designs in terms of the fuel mass flow rate required for a given task (e.g., for a fixed-thrust requirement or a given mission).
In addition to the various vehicle design optimizations, operational optimizations are conducted to examine the advantage if any of energy exchange to maintain shock-on-lip for both design and off-design conditions. Parametric studies of the hypersonic vehicle sub-systems and components are also conducted and provide further insights into the impacts that the design and operational decision variables and flow properties have on the rates of exergy destruction.
Advisors/Committee Members: von Spakovsky, Michael R. (committeechair), Moorhouse, David (committee member), Ellis, Michael W. (committee member), O'Brien, Walter F. Jr. (committee member).
Subjects/Keywords: scramjet; exergy; optimization; hypersonic's
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APA ·
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MLA ·
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Export
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APA (6th Edition):
Markell, K. C. (2005). Exergy Methods for the Generic Analysis and Optimization of Hypersonic Vehicle Concepts. (Masters Thesis). Virginia Tech. Retrieved from http://hdl.handle.net/10919/31256
Chicago Manual of Style (16th Edition):
Markell, Kyle Charles. “Exergy Methods for the Generic Analysis and Optimization of Hypersonic Vehicle Concepts.” 2005. Masters Thesis, Virginia Tech. Accessed January 24, 2021.
http://hdl.handle.net/10919/31256.
MLA Handbook (7th Edition):
Markell, Kyle Charles. “Exergy Methods for the Generic Analysis and Optimization of Hypersonic Vehicle Concepts.” 2005. Web. 24 Jan 2021.
Vancouver:
Markell KC. Exergy Methods for the Generic Analysis and Optimization of Hypersonic Vehicle Concepts. [Internet] [Masters thesis]. Virginia Tech; 2005. [cited 2021 Jan 24].
Available from: http://hdl.handle.net/10919/31256.
Council of Science Editors:
Markell KC. Exergy Methods for the Generic Analysis and Optimization of Hypersonic Vehicle Concepts. [Masters Thesis]. Virginia Tech; 2005. Available from: http://hdl.handle.net/10919/31256
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Open Access Theses and Dissertations