+publisher:"University of Illinois – Urbana-Champaign" +contributor:("Bragg, Michael B.")

University of Illinois – Urbana-Champaign

1. Diebold, Jeffrey Michael. The effects of turbulence on the measurements of five-hole probes.

Degree: PhD, Aerospace Engineering, 2016, University of Illinois – Urbana-Champaign

► The primary goals of this research were to quantify the effects of turbulence on the measurements of five-hole pressure probes (5HP) and to develop a… (more)

▼ The primary goals of this research were to quantify the effects of turbulence on the measurements of five-hole pressure probes (5HP) and to develop a model capable of predicting the response of a 5HP to turbulence. The five-hole pressure probe is a commonly used device in experimental fluid dynamics and aerodynamics. By measuring the pressure at the five pressure ports located on the tip of the probe it is possible to determine the total pressure, static pressure and the three components of velocity at a point in the flow. Previous research has demonstrated that the measurements of simple pressure probes such as Pitot probes are significantly influenced by the presence of turbulence. Turbulent velocity fluctuations contaminate the measurement of pressure due to the nonlinear relationship between pressure and velocity as well as the angular response characteristics of the probe. Despite our understanding of the effects of turbulence on Pitot and static pressure probes, relatively little is known about the influence of turbulence on five-hole probes. This study attempts to fill this gap in our knowledge by using advanced experimental techniques to quantify these turbulence-induced errors and by developing a novel method of predicting the response of a five-hole probe to turbulence. A few studies have attempted to quantify turbulence-induced errors in five-hole probe measurements but they were limited by their inability to accurately measure the total and static pressure in the turbulent flow. The current research utilizes a fast-response five-hole probe (FR5HP) in order to accurately quantify the effects of turbulence on different standard five-hole probes (Std5HP). The FR5HP is capable of measuring the instantaneous flowfield and unlike the Std5HP the FR5HP measurements are not contaminated by the turbulent velocity fluctuations. Measurements with the FR5HP and two different Std5HPs were acquired in the highly turbulent iii wakes of 2D and 3D cylinders in order to quantify the turbulence-induced errors in Std5HP measurements. The primary contribution of this work is the development and validation of a simulation method to predict the measurements of a Std5HP in an arbitrary turbulent flow. This simulation utilizes a statistical approach to estimating the pressure at each port on the tip of the probe. The angular response of the probe is modeled using experimental calibration data for each five-hole probe. The simulation method is validated against the experimental measurements of the Std5HPs, and then used to study the how the characteristics of the turbulent flowfield influence the measurements of the Std5HPs. It is shown that total pressure measured by a Std5HP is increased by axial velocity fluctuations but decreased by the transverse fluctuations. The static pressure was shown to be very sensitive to the transverse fluctuations while the axial fluctuations had a negligible effect. As with Pitot probes, the turbulence-induced errors in the Std5HPs measurements were dependent on both the properties of the… Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B%22%29&pagesize-30">Bragg, Michael B (advisor), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B%22%29&pagesize-30">Bragg, Michael B (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20%20S%22%29&pagesize-30">Elliott, Gregory S (committee member), Champaign%22%20%2Bcontributor%3A%28%22Selig%2C%20Michael%20S%22%29&pagesize-30">Selig, Michael S (committee member), Champaign%22%20%2Bcontributor%3A%28%22Chamorro%2C%20Leonardo%20P%22%29&pagesize-30">Chamorro, Leonardo P (committee member).

Subjects/Keywords: Multihole probes; Five-hole probes (5HP); Turbulence; Experimental fluid dynamics

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APA (6^th Edition):

Diebold, J. M. (2016). The effects of turbulence on the measurements of five-hole probes. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/97238

Chicago Manual of Style (16^th Edition):

Diebold, Jeffrey Michael. “The effects of turbulence on the measurements of five-hole probes.” 2016. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/97238.

MLA Handbook (7^th Edition):

Diebold, Jeffrey Michael. “The effects of turbulence on the measurements of five-hole probes.” 2016. Web. 25 Jan 2021.

Vancouver:

Diebold JM. The effects of turbulence on the measurements of five-hole probes. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2016. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/97238.

Council of Science Editors:

Diebold JM. The effects of turbulence on the measurements of five-hole probes. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2016. Available from: http://hdl.handle.net/2142/97238

2. Deters, Robert. Performance and slipstream characteristics of small-scale propellers at low Reynolds numbers.

Degree: PhD, 4048, 2014, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/49607

► The low Reynolds number effects of small-scale propellers were investigated. At the Reynolds numbers of interest (below 100,000), a decrease in lift and an increase… (more)

▼ The low Reynolds number effects of small-scale propellers were investigated. At the Reynolds numbers of interest (below 100,000), a decrease in lift and an increase in drag is common making it difficult to predict propeller performance characteristics. A propeller testing apparatus was built to test small scale propellers in static conditions and in an advancing flow. Twenty-seven off-the-shelf propellers, with diameters ranging from 2.25 in to 9 in, were tested in order to determine the general effects of low Reynolds numbers on small propellers. From these tests, increasing the Reynolds number for a propeller increases its efficiency by either increasing the thrust produced or decreasing the power. By doubling the Reynolds number of a propeller, it is not uncommon to increase the efficiency by more the 10%. Using off-the-shelf propellers limits the geometry available and finding propellers of the same geometry but of different scale is very difficult. To solve this problem, four propellers were design and built using a 3D printer. Two of the propellers were simple rectangular twisted blades of different chords. Another propeller was modeled after a full-scale propeller. The fourth propeller was created using inverse design to minimize power loss. Each propeller was built in a 5-in and 9-in diameter version in order to test a larger range of Reynolds numbers. A separate propeller blade and hub system was created to allow each propeller to be tested with different pitch angles and to test each propeller in a 2-, 3-, and 4-blade version. From the performance results of the 3D printed propellers, it was shown that propellers of different scale, but tested at the same Reynolds number, had about the same performance results. Finally, the slipstreams of different propellers were measured using a 7-hole probe. Propeller slipstreams can have a large effect on the aerodynamics of lifting surfaces downstream of the propeller. Small UAVs and MAVs flying in close proximity will also fly into the propeller slipstream of a neighbor. These slipstreams can produce relatively large gusts for very small and light aircraft, and the slipstreams can persist far downstream. Knowing the characteristics of propeller slipstreams will help in designing aircraft that can better handle close proximity flight. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Selig%2C%20Michael%20S.%22%29&pagesize-30">Selig, Michael S. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Selig%2C%20Michael%20S.%22%29&pagesize-30">Selig, Michael S. (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20S.%22%29&pagesize-30">Elliott, Gregory S. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Ragheb%2C%20Magdi%22%29&pagesize-30">Ragheb, Magdi (committee member).

Subjects/Keywords: propeller; slipstream; propeller performance; propeller efficiency; wind tunnel testing; 3D printing; Unmanned Aerial Vehicle (UAV); Micro air vehicle (MAV); low Reynolds number; 7-hole probe; pressure probe measurements; propeller testing; static tests; advancing flow tests; propeller wake; propeller thrust; propeller power; axial velocity; swirl; propeller wing interaction

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APA · Chicago · MLA · Vancouver · CSE | Export to Zotero / EndNote / Reference Manager

APA (6^th Edition):

Deters, R. (2014). Performance and slipstream characteristics of small-scale propellers at low Reynolds numbers. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/49607

Chicago Manual of Style (16^th Edition):

Deters, Robert. “Performance and slipstream characteristics of small-scale propellers at low Reynolds numbers.” 2014. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/49607.

MLA Handbook (7^th Edition):

Deters, Robert. “Performance and slipstream characteristics of small-scale propellers at low Reynolds numbers.” 2014. Web. 25 Jan 2021.

Vancouver:

Deters R. Performance and slipstream characteristics of small-scale propellers at low Reynolds numbers. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2014. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/49607.

Council of Science Editors:

Deters R. Performance and slipstream characteristics of small-scale propellers at low Reynolds numbers. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2014. Available from: http://hdl.handle.net/2142/49607

3. Camarinha Fujiwara, Gustavo. Design of 3D swept wing hybrid models for icing wind tunnel tests.

Degree: MS, 4048, 2015, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/72880

► The study of aircraft icing is critical to ensure the safety of any aircraft that might experience icing conditions in flight, including general, commercial, and… (more)

▼ The study of aircraft icing is critical to ensure the safety of any aircraft that might experience icing conditions in flight, including general, commercial, and military aviation. The certification of modern commercial transports requires manufacturers to demonstrate that these aircraft can safely operate during icing conditions through a set of flight tests, consistent with the standards set forth by the Federal Aviation Administration. This is often expensive and challenging to find the appropriate icing test conditions. Thus, both computational methods and icing wind tunnel experiments are utilized during the design and certification of aircraft ice-protection systems to provide a controlled and repeatable environment to mitigate risks, reduce costs, and validate the existing computational icing tools. However, the existing icing wind tunnel facilities cannot accommodate large wings such as those found on modern commercial aircraft without being dramatically scaled. Two methods of scaling exist. The first geometrically scales the entire geometry to fit inside the tunnel test section and then scales the icing conditions to obtain icing similitude. The second maintains the full-scale leading edge of the reference geometry and replaces the aft section with a truncated trailing edge that produces a similar flowfield around the leading edge with a significantly shorter chord, reducing model size and tunnel blockage. This type of model is referred to as a hybrid and its biggest advantage lies in the fact that it is designed to produce full-scale ice shapes, while reducing or even eliminating the need for icing scaling. While a design method for a straight, untapered hybrid wing is well documented and there is a broad set of experimental data available, the design of a swept, hybrid wing lacks both a design method and experimental data. This thesis established a design method for large hybrid swept wings that reproduce full-scale ice accretions through icing wind tunnel tests. The design method was broken down in two steps: 1) A 2D hybrid airfoil design, and 2) A 3D hybrid swept wing design. Multiple existing computational tools were employed and several parametric studies performed. It was shown, in 2D, that matching the stagnation point location on the leading edge of the hybrid airfoil had a first-order impact on matching the full-scale ice shape, while matching the suction peak magnitude and location had a second-order effect. The closer to the leading edge lift was generated for a given hybrid design, the less total load was required to reach the same stagnation point location. As an implication, more front-loaded airfoils required less lift than more aft-loaded ones to reach the same stagnation point location on a hybrid airfoil. More front load also increased the risk of flow separation near the leading edge, while more aft load increased the risk of separation near the trailing edge. Finally, higher hybrid scale factors were shown to increase the risk of flow separation. In 3D, sweep angle was… Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor).

Subjects/Keywords: Aerodynamics; aircraft icing; wind tunnel; icing; ice shape; Accretion; hybrid; sweep; swept; wing; computational fluid dynamics (CFD); aircraft certification; wing design

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APA (6^th Edition):

Camarinha Fujiwara, G. (2015). Design of 3D swept wing hybrid models for icing wind tunnel tests. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/72880

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Camarinha Fujiwara, Gustavo. “Design of 3D swept wing hybrid models for icing wind tunnel tests.” 2015. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/72880.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Camarinha Fujiwara, Gustavo. “Design of 3D swept wing hybrid models for icing wind tunnel tests.” 2015. Web. 25 Jan 2021.

Vancouver:

Camarinha Fujiwara G. Design of 3D swept wing hybrid models for icing wind tunnel tests. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2015. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/72880.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Camarinha Fujiwara G. Design of 3D swept wing hybrid models for icing wind tunnel tests. [Thesis]. University of Illinois – Urbana-Champaign; 2015. Available from: http://hdl.handle.net/2142/72880

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

4. Mortonson, Andrew. Use of hybrid airfoil design in icing wind tunnel tests of large scale swept wings.

Degree: MS, 4048, 2012, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/42351

► In-flight ice accretion is an important safety consideration for modern aircraft. The certification of commercial airliners for flight into known icing is difficult due to… (more)

Subjects/Keywords: truncated airfoil; airframe icing; hybrid airfoil; Icing Research Tunnel; Common Research Model (CRM)

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APA (6^th Edition):

Mortonson, A. (2012). Use of hybrid airfoil design in icing wind tunnel tests of large scale swept wings. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/42351

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Mortonson, Andrew. “Use of hybrid airfoil design in icing wind tunnel tests of large scale swept wings.” 2012. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/42351.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Mortonson, Andrew. “Use of hybrid airfoil design in icing wind tunnel tests of large scale swept wings.” 2012. Web. 25 Jan 2021.

Vancouver:

Mortonson A. Use of hybrid airfoil design in icing wind tunnel tests of large scale swept wings. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2012. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/42351.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Mortonson A. Use of hybrid airfoil design in icing wind tunnel tests of large scale swept wings. [Thesis]. University of Illinois – Urbana-Champaign; 2012. Available from: http://hdl.handle.net/2142/42351

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

5. Wiberg, Brock. Large-scale swept-wing ice accretion modeling in the NASA Glenn Icing Research Tunnel using LEWICE3D.

Degree: MS, 4048, 2014, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/46609

► The study of aircraft icing is necessary to ensure the safety of commercial, military, and general aviation aircraft. The certification of modern commercial transports requires… (more)

▼ The study of aircraft icing is necessary to ensure the safety of commercial, military, and general aviation aircraft. The certification of modern commercial transports requires manufacturers to demonstrate that these aircraft can safely operate during icing conditions consistent with the standards set forth by the Federal Aviation Administration (FAA). While some of these tests are performed on actual aircraft in flight, this is often very expensive and does not provide an adequately controlled matrix of test conditions. Computational tools are used throughout the design and certification of anti-ice systems. However, computational methods alone are not sufficient for aircraft certification. Icing wind tunnels are used for aircraft certification to reduce costs, provide a controlled test matrix of conditions, and validate computational icing tools. The size of aircraft models that can be tested in icing wind tunnels is limited by the size and capability of existing facilities. Large wings, such as those found on modern narrow and wide-body commercial transports, cannot fit in existing test sections without being dramatically scaled. Two methods of scaling exist. The first involves geometrically scaling a section of the reference wing to fit inside the tunnel test section and then scaling the icing conditions in order to maintain icing similitude. The second method maintains the full-scale leading edge of the reference geometry but replaces the aft section of the wing with a tail that is designed to produce similar flow around the leading edge but with a considerably shorter chord length, reducing model size and blockage. This type of model is called a hybrid and is used to generate full-scale ice shapes so that, in the simplest cases, no icing scaling is necessary. However, the methods can be combined so that the hybrid model design is used to maintain geometric similitude while icing scaling is employed to account for differences in pressure, velocity, or other conditions. Modern commercial transport aircraft have large, swept wings. While a broad set of experimental data exist in the literature for airfoil and straight wing icing, there is a distinct lack of data for large, swept wings. Such data is needed in order to better understand the 3D icing physics on swept wings and to allow computational tools to be developed and validated for 3D ice features such as scallops. In this thesis, computational tools were used to better understand the flow over a large-scale, swept-wing, hybrid model mounted vertically in the NASA Glenn Icing Research Tunnel (IRT). Fluent, a commercial CFD code, was used to calculate flows around the flapped-hybrid model in the IRT, mounted with the root at the floor and the tip at the ceiling of the test section. Inviscid analysis reveals that the upwash ahead of the model causes the local lift coefficient to increase significantly across the swept model due to the effect of the floor and ceiling. This change in spanwise loading is shown to move the attachment line location farther… Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor).

Subjects/Keywords: aircraft icing; icing; ice shape; Accretion; sweep; swept; wing; computational fluid dynamics (CFD); aircraft certification; hybrid; similitude

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APA (6^th Edition):

Wiberg, B. (2014). Large-scale swept-wing ice accretion modeling in the NASA Glenn Icing Research Tunnel using LEWICE3D. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/46609

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Wiberg, Brock. “Large-scale swept-wing ice accretion modeling in the NASA Glenn Icing Research Tunnel using LEWICE3D.” 2014. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/46609.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Wiberg, Brock. “Large-scale swept-wing ice accretion modeling in the NASA Glenn Icing Research Tunnel using LEWICE3D.” 2014. Web. 25 Jan 2021.

Vancouver:

Wiberg B. Large-scale swept-wing ice accretion modeling in the NASA Glenn Icing Research Tunnel using LEWICE3D. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2014. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/46609.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Wiberg B. Large-scale swept-wing ice accretion modeling in the NASA Glenn Icing Research Tunnel using LEWICE3D. [Thesis]. University of Illinois – Urbana-Champaign; 2014. Available from: http://hdl.handle.net/2142/46609

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

6. Monastero, Marianne. Validation of 3-D ice accretion documentation and replication method including pressure-sensitive paint.

Degree: MS, 4048, 2014, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/46651

► Accurate representation of ice accretions is important to the study and understanding of aircraft icing. For research and certification purposes, replicas of ice accretions generated… (more)

▼ Accurate representation of ice accretions is important to the study and understanding of aircraft icing. For research and certification purposes, replicas of ice accretions generated from icing wind tunnels are fabricated to perform aerodynamic tests in dry-air wind tunnels. The currently employed replication method consists of creating molds from original ice accretions and producing castings from the molds for wind tunnel testing. While this method reproduces the geometric features and aerodynamic effects of the original ice accretions well in the replicated ice shapes, it has several limitations. This method cannot scale the ice shapes to sizes other than the original and does not produce a digital record of the ice shape. Both of these capabilities are desirable in iced-aerodynamics research. To address these needs, NASA developed a methodology to obtain a digital record of ice accretions through the implementation of a laser scanner system. The resulting scan can be used in conjunction with rapid-prototype methods to generate ice shapes for wind tunnel testing. This work is a validation of the 3-D ice accretion measurement methodology where the ice shapes generated by both the currently-used and newly-developed methods from the same initial ice accretion are compared using force balance-derived aerodynamic performance, surface and wake pressures, and pressure-sensitive paint (PSP) data. The 3-D features of the tested ice shapes necessitated the use of a technique capable of obtaining high resolution data. The PSP technique allowed pressure coefficient data to be obtained over a larger area and at a greater resolution than is possible by only using the surface pressure tap method. The results discussed show the capability of the PSP technique, as implemented in the 3ft by 4ft subsonic wind tunnel at the University of Illinois, to resolve aerodynamic differences between ice shapes made from both the current and newly developed ice accretion replication methods. The same trends were observed in the PSP data as were found in the aerodynamic performance and pressure tap data, and the newly developed 3-D ice accretion measurement methodology produced ice shapes which aerodynamically agreed well with ice shapes generated from the mold and casting method. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor).

Subjects/Keywords: icing; Pressure-sensitive Paint; Aerodynamics; Wind Tunnel

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APA (6^th Edition):

Monastero, M. (2014). Validation of 3-D ice accretion documentation and replication method including pressure-sensitive paint. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/46651

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Monastero, Marianne. “Validation of 3-D ice accretion documentation and replication method including pressure-sensitive paint.” 2014. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/46651.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Monastero, Marianne. “Validation of 3-D ice accretion documentation and replication method including pressure-sensitive paint.” 2014. Web. 25 Jan 2021.

Vancouver:

Monastero M. Validation of 3-D ice accretion documentation and replication method including pressure-sensitive paint. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2014. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/46651.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Monastero M. Validation of 3-D ice accretion documentation and replication method including pressure-sensitive paint. [Thesis]. University of Illinois – Urbana-Champaign; 2014. Available from: http://hdl.handle.net/2142/46651

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

7. Bottalla, Joseph R. Performance of an airfoil with a power-saving, tab-assisted flap system.

Degree: MS, 4048, 2011, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/26028

► This study investigated the feasibility of reducing control surface input power with the use of a tab-assisted flap. Wind tunnel tests were conducted at the… (more)

▼ This study investigated the feasibility of reducing control surface input power with the use of a tab-assisted flap. Wind tunnel tests were conducted at the University of Illinois at Urbana-Champaign (UIUC) on a NACA 3415 airfoil model with a flap including a trim tab. Measurements were taken for two configurations: a baseline fixed tab case where tab deflection was zero and a tabbed case where multiple flap and tab angle combinations were tested. Hinge moment measurements were taken for both the flap and tab for comparison between the two cases. In addition; lift, drag and moment measurements along with surface pressures were acquired to aid in the analysis of the concept and provide flow diagnostics. The data were compared to computational results which compared well with the exception of flap and tab deflection cases where large regions of unsteady separated flow were present. All data were taken at a Reynolds number of 1.8 million and Mach number of 0.18. To analyze the power-savings capability of a tab-assisted flap, several studies were conducted: a generalized tab performance study evaluating hinge moment reduction and a quasidynamic study using the static data to calculate work savings for two simulated flap deflections. The generalized tab performance study revealed large hinge moment reductions for each of the flap deflections when using the tab to actuate the flap. These reductions came at the cost of increased drag, reduced lift and loss of flap effectiveness. The quasi-dynamic study produced significantly large work savings for both simulated flap deflection schedules. Even though this study ignored the effect on lift and drag, as well as the unsteady aerodynamics and control surface inertia, it suggests the large power-savings potential of a tab-assisted flap. A flow visualization analysis was performed to further assess the loss of flap effectiveness observed in the tab performance calculations as well as the non-linear behavior in the lift, drag and hinge moment data. The results of the analysis showed the complex flowfield behavior in cases where the flap and tab deflections were of opposite sign and larger magnitude and identified the cause of behavior in other cases where data non-linearities existed. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor).

Subjects/Keywords: Wind tunnel; Aerodynamics; Airfoil; Flap; Tab; Performance

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APA (6^th Edition):

Bottalla, J. R. (2011). Performance of an airfoil with a power-saving, tab-assisted flap system. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/26028

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Bottalla, Joseph R. “Performance of an airfoil with a power-saving, tab-assisted flap system.” 2011. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/26028.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Bottalla, Joseph R. “Performance of an airfoil with a power-saving, tab-assisted flap system.” 2011. Web. 25 Jan 2021.

Vancouver:

Bottalla JR. Performance of an airfoil with a power-saving, tab-assisted flap system. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2011. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/26028.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Bottalla JR. Performance of an airfoil with a power-saving, tab-assisted flap system. [Thesis]. University of Illinois – Urbana-Champaign; 2011. Available from: http://hdl.handle.net/2142/26028

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

8. Diebold, Jeffrey. Aerodynamics of a swept wing with leading-edge ice at low Reynolds number.

Degree: MS, 4048, 2012, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/34279

► An experimental study of the aerodynamics of a swept wing with ice at low Reynolds number has been performed. The goal of this work was… (more)

▼ An experimental study of the aerodynamics of a swept wing with ice at low Reynolds number has been performed. The goal of this work was to demonstrate the use of various experimental techniques applied to understanding the aerodynamic effects of a leading-edge ice simulation on a highly swept, high-aspect ratio wing. The swept wing model was a modified version of the NASA Common Research Model, designed to represent a typical wide body commercial airliner. The modified geometry of the model used in this study included a ΛLE = 35º, AR = 8.3 and λ=0.296. The experimental techniques used were force balance measurements, surface pressure measurements, surface oil flow visualization and 5-hole probe wake surveys. Tests were performed at Reynolds numbers of 3x105, 6x105 and 7.8x105 and corresponding Mach numbers of 0.08, 0.15 and 0.2. Force balance results show that the ice shape had a significant effect on performance. The stalling angle of attack and maximum lift coefficient were reduced while the drag was increased throughout the entire range of angles of attack tested. A large leading-edge vortex behind the ice shape was observed in the oil flow, and the pressure measurements showed this vortex had a significant effect on the pressure field over the wing. From the 5-hole wake survey results it was seen that the ice shape increased the profile drag while the induced drag was relatively unaffected. Using the oil flow, the evolution of the leading-edge vortex was observed and features seen in the oil flow were related to features observed in the wake. The flowfield of the iced wing contained several similarities to the flowfield of an airfoil with horn ice; however, there were several important differences due to the three-dimensional nature of the swept wing flowfield. The spanwise distribution of lift and drag were also investigated. By comparing the distributions on the clean and iced wing it was possible to determine that the ice had the largest impact on the aerodynamics of the outboard sections. It was also shown that features observed in the surface oil flow and the wake can be correlated to certain features in the lift and drag distributions. Finally, the effect of the Reynolds number was investigated. Over the range of Reynolds numbers tested, which was not representative of flight, it was observed that the Reynolds number had a reduced influence on the iced wing. This trend was observed in the performance and flowfield results. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor).

Subjects/Keywords: aircraft icing; swept wing icing; swept wing aerodynamics

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APA (6^th Edition):

Diebold, J. (2012). Aerodynamics of a swept wing with leading-edge ice at low Reynolds number. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/34279

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Diebold, Jeffrey. “Aerodynamics of a swept wing with leading-edge ice at low Reynolds number.” 2012. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/34279.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Diebold, Jeffrey. “Aerodynamics of a swept wing with leading-edge ice at low Reynolds number.” 2012. Web. 25 Jan 2021.

Vancouver:

Diebold J. Aerodynamics of a swept wing with leading-edge ice at low Reynolds number. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2012. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/34279.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Diebold J. Aerodynamics of a swept wing with leading-edge ice at low Reynolds number. [Thesis]. University of Illinois – Urbana-Champaign; 2012. Available from: http://hdl.handle.net/2142/34279

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

9. Hortensius, Ruben. An experimental study of the aft guide vanes of an engine bypass nacelle for low-boom supersonic flight.

Degree: MS, 4048, 2012, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/29820

► Due to the large disturbance created by the sonic boom, supersonic flight is strictly controlled by the FAA. One way in which to minimize the… (more)

▼ Due to the large disturbance created by the sonic boom, supersonic flight is strictly controlled by the FAA. One way in which to minimize the sonic boom is through shape-tailoring of the aircraft body and of the propulsion system. To this end, a new supersonic engine concept has been proposed, wherein a core turbofan engine, which has a non-axisymmetric external profile due to a protruding gearbox, has been circularized. A new, secondary, bypass with a highly complex internal geometry is created during this process. The high-flow nacelle bypass geometry includes a forward and aft fairing to direct the flow around the gearbox, a set of thin forward guide vanes, and a set of thick, strut-like aft guide vanes. The aft guide vanes, which also serve structural purposes, are used to direct the flow such that the exhaust is a uniform, nearly-full annular cross-section, and to choke and then accelerate the flow to supersonic freestream conditions upon exit. A supersonic wind tunnel facility at the University of Illinois was modified and used to simulate the flow through the aft bypass at approximately 6% scale. In order to aid in understanding the effect of the aft vanes, two models, one with and one without guide vanes, are studied. Due to facility limitations, the design operating condition could not be achieved; a series of off-design operating conditions are tested instead. Radial pressure surveys are conducted at several azimuthal stations at the inlet to the aft bypass in order to establish in-flow conditions. Static pressure taps on the model surface provide insight into the nature of the flow through the bypass on a per channel basis. An isentropic-case comparison, an estimate of total pressure losses, and mass flow rate calculations were performed. Pressure data were supplemented with schlieren imagery and surface oil flow visualization. Results indicate the flow through the aft bypass is highly three-dimensional and contains a large amount of flow separation in the off-design conditions tested. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20S.%22%29&pagesize-30">Elliott, Gregory S. (advisor).

Subjects/Keywords: Gulfstream; Rolls-Royce; supersonic engine; bypass; high-flow bypass nacelle; Guide Vanes; oil flow visualization; schlieren; supersonic business jet; quiet supersonic business jet; flow through vanes; choked flow; sonic boom; annular flow; wind tunnel testing; experimental testing

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APA (6^th Edition):

Hortensius, R. (2012). An experimental study of the aft guide vanes of an engine bypass nacelle for low-boom supersonic flight. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/29820

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Hortensius, Ruben. “An experimental study of the aft guide vanes of an engine bypass nacelle for low-boom supersonic flight.” 2012. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/29820.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Hortensius, Ruben. “An experimental study of the aft guide vanes of an engine bypass nacelle for low-boom supersonic flight.” 2012. Web. 25 Jan 2021.

Vancouver:

Hortensius R. An experimental study of the aft guide vanes of an engine bypass nacelle for low-boom supersonic flight. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2012. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/29820.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Hortensius R. An experimental study of the aft guide vanes of an engine bypass nacelle for low-boom supersonic flight. [Thesis]. University of Illinois – Urbana-Champaign; 2012. Available from: http://hdl.handle.net/2142/29820

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

10. Herrera, Arthur A. An experimental study of the forward guide vanes of an engine bypass nacelle for low-boom supersonic flight.

Degree: MS, 4048, 2011, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/26159

► Current FAA regulations limit the speed of civilian passenger jets to subsonic or transonic speeds due to the effects of the sonic boom produced by… (more)

▼ Current FAA regulations limit the speed of civilian passenger jets to subsonic or transonic speeds due to the effects of the sonic boom produced by aircraft during supersonic flight. An engine nacelle design has been proposed that removes low-quality flow from the engine core and reduces the signature of the sonic boom caused by the external protuberances of a traditional engine housing. This new concept incorporates an outer nacelle surrounding the asymmetric engine surface, which creates a highly-complex, secondary bypass flow. Due to the complexity of the flow within this region of the engine, an experimental study has been conducted on the integration of guide vanes within the subsonic portion of the bypass region as the flow is diverted around a partial annular blockage. A wind tunnel facility at the University of Illinois at Urbana-Champaign accommodates an approximately 1/6th scale model that simulates the three-dimensional flowfield around the engine components. In order to observe the influence of the guide vanes on the overall flow quality, tests were also conducted on a model without forward vanes. Pressure data were collected upstream and downstream of the guide vanes at several axial locations, with high resolution in both the azimuthal and radial directions. In addition to flow speed, flow direction was also analyzed via a five-hole multi-directional probe and surface flow visualization techniques. Experimental flow analysis in this study was conducted to support computational models and to provide insight into techniques that may further improve the flow characteristics within the bypass flow region. Results from the study indicate upstream flow uniformity due to the presence of the guide vanes as well as highly-complex flow features downstream of the vanes. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20S.%22%29&pagesize-30">Elliott, Gregory S. (advisor).

Subjects/Keywords: Supersonic Business Jet (SSBJ); Sonic Boom; Engine Nacelle; Annular Flow; Annular Blockage; Guide Vanes; Wind Tunnel Testing; Experimental Testing

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APA (6^th Edition):

Herrera, A. A. (2011). An experimental study of the forward guide vanes of an engine bypass nacelle for low-boom supersonic flight. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/26159

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Herrera, Arthur A. “An experimental study of the forward guide vanes of an engine bypass nacelle for low-boom supersonic flight.” 2011. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/26159.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Herrera, Arthur A. “An experimental study of the forward guide vanes of an engine bypass nacelle for low-boom supersonic flight.” 2011. Web. 25 Jan 2021.

Vancouver:

Herrera AA. An experimental study of the forward guide vanes of an engine bypass nacelle for low-boom supersonic flight. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2011. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/26159.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Herrera AA. An experimental study of the forward guide vanes of an engine bypass nacelle for low-boom supersonic flight. [Thesis]. University of Illinois – Urbana-Champaign; 2011. Available from: http://hdl.handle.net/2142/26159

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

11. Ansell, Phillip J. Unsteady modes in the flowfield about an airfoil with a leading-edge horn-ice shape.

Degree: PhD, 4048, 2014, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/46620

► An analysis of unsteady modes present in the flowfield of an airfoil with a leading-edge horn-ice shape was performed in the current study. An NACA… (more)

▼ An analysis of unsteady modes present in the flowfield of an airfoil with a leading-edge horn-ice shape was performed in the current study. An NACA 0012 airfoil was tested in a subsonic wind tunnel at Re = 1.8 × 106. In addition to the clean configuration, the airfoil model was also tested with a set of boundary-layer trips, a two-dimensional extrusion of a horn-ice shape casting, and an array of simulated icing configurations created using simple geometries. Time-averaged and unsteady static pressure measurements were acquired about the airfoil surface, along with unsteady wake velocity and surface hot-film array measurements. Additionally, surface and off-body flow visualization techniques were used to visualize the airfoil flowfield. A technique was also developed to determine the unsteady shear-layer reattachment location of the ice-induced laminar separation bubble downstream of the horn-ice shape using the surface hot-film array measurements. The maximum amount of unsteadiness in the iced-airfoil flowfield was observed to increase with increasing angle of attack. For a fixed angle of attack prior to stall, a change in the feature height of the simulated ice shape led to a change in the distribution of flowfield unsteadiness, but did not change the maximum levels of unsteadiness present in the flowfield. The iced-airfoil flowfield unsteadiness was primarily associated with three different frequencies. The first was represented by an increase in spectral energy across a broad-band frequency range, and was observed just upstream of shear-layer reattachment as well as downstream of shear-layer reattachment. This increase in spectral energy was caused by the regular mode of unsteadiness due to vortical motion in the separated shear layer and vortex shedding from the separation bubble. The average Strouhal number of this regular mode corresponded to StL = 0.60, and the average vortex convection velocity was observed to be 0.45U∞. These values were highly consistent with those reported elsewhere in the literature. The other two frequencies were much lower and were observed as narrow-band peaks in the spectral content of the acquired measurements that were primarily present in the region covered by the ice-induced separation bubble. The first was attributed to the shear-layer flapping phenomenon and was particularly dominant in the upstream portion of the separation bubble. The Strouhal number associated with this shear-layer flapping mode corresponded to Sth = 0.0185, which was consistent with those reported in studies of separation bubbles about canonical geometries. The second frequency was lower than that of shear-layer flapping and was associated with a low-frequency mode of unsteadiness that can occur prior to static stall for airfoils of thin-airfoil stall type. This low-frequency mode was characterized by a low-frequency oscillation of the airfoil circulation, and it was clearly identified in the spectral content of the iced-airfoil lift coefficient. The resulting values of Strouhal number… Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20S.%22%29&pagesize-30">Elliott, Gregory S. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Selig%2C%20Michael%20S.%22%29&pagesize-30">Selig, Michael S. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Christensen%2C%20Kenneth%20T.%22%29&pagesize-30">Christensen, Kenneth T. (committee member).

Subjects/Keywords: airfoil; icing; ice accretion; flowfield unsteadiness; shear-layer flapping; vortex shedding; aircraft; separation bubble; separated flow; airfoil circulation; Strouhal number; shear-layer reattachment; stall; unsteady flow

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APA · Chicago · MLA · Vancouver · CSE | Export to Zotero / EndNote / Reference Manager

APA (6^th Edition):

Ansell, P. J. (2014). Unsteady modes in the flowfield about an airfoil with a leading-edge horn-ice shape. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/46620

Chicago Manual of Style (16^th Edition):

Ansell, Phillip J. “Unsteady modes in the flowfield about an airfoil with a leading-edge horn-ice shape.” 2014. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/46620.

MLA Handbook (7^th Edition):

Ansell, Phillip J. “Unsteady modes in the flowfield about an airfoil with a leading-edge horn-ice shape.” 2014. Web. 25 Jan 2021.

Vancouver:

Ansell PJ. Unsteady modes in the flowfield about an airfoil with a leading-edge horn-ice shape. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2014. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/46620.

Council of Science Editors:

Ansell PJ. Unsteady modes in the flowfield about an airfoil with a leading-edge horn-ice shape. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2014. Available from: http://hdl.handle.net/2142/46620

University of Illinois – Urbana-Champaign

12. Ansell, Phillip J. Flight envelope protection using flap hinge moment measurement.

Degree: MS, 4048, 2010, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/16147

► An experimental investigation of the sensitivity of flap hinge moment to airfoil surface contamination was conducted at the University of Illinois at Urbana-Champaign Aerodynamics Research… (more)

▼ An experimental investigation of the sensitivity of flap hinge moment to airfoil surface contamination was conducted at the University of Illinois at Urbana-Champaign Aerodynamics Research Lab. Tests were conducted on two airfoil models, an NACA 3415 and an NACA 23012, at Reynolds numbers of 1.8 × 106 and 1.0 × 106. The effects of six different simulated contamination configurations on the performance characteristics of both airfoils were tested. These configurations consisted of glaze ice, rime ice, two severities of distributed leading-edge roughness, three-dimensional leading-edge damage, and three-dimensional upper-surface damage. Additionally, the effects of flap deflection and trim tab deflection on the unsteady hinge moment were studied. Results from this study found that large increases in Ch.StDev often occurred at the same angle of attack as Cl,max. By correlating regions of separated flow observed in Cp distributions and fluorescent-oil flow visualizations to Ch,StDev at discrete angles of attack, it was determined that regions of boundary-layer separation were the primary driver for large increases in unsteadiness in the hinge moment. It was also found that the unsteady hinge moment had negligible dependence on trim tab deflection. The response of Ch,StDev was dependent on the stalling characteristics of the airfoil model. Of all of the contamination configurations tested, the two simulated ice cases had the largest effect on the performance of the airfoils. For the distributed leading-edge roughness cases, the larger roughness elements had a larger effect on the performance than the smaller roughness elements, but the Ch,StDev response of both roughness cases were comparable. While the 3D simulated damage cases did not significantly affect the lifting characteristics of either model, the magnitude of the Ch,StDev response of the 3D simulated damage case was comparable to the 2D contamination cases. Additionally, the large increase in Ch,StDev occurred prior to stall due to localized regions of separated flow that resulted from the simulated damage. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor).

Subjects/Keywords: hinge moment; envelope protection; icing; simulated icing; roughness; simulated damage; contamination; Aerodynamics; aeronautics; stall; NACA 3415; NACA 23012; glaze ice; rime ice; trim tab; Flap; control surface

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APA · Chicago · MLA · Vancouver · CSE | Export to Zotero / EndNote / Reference Manager

APA (6^th Edition):

Ansell, P. J. (2010). Flight envelope protection using flap hinge moment measurement. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/16147

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Ansell, Phillip J. “Flight envelope protection using flap hinge moment measurement.” 2010. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/16147.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Ansell, Phillip J. “Flight envelope protection using flap hinge moment measurement.” 2010. Web. 25 Jan 2021.

Vancouver:

Ansell PJ. Flight envelope protection using flap hinge moment measurement. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2010. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/16147.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Ansell PJ. Flight envelope protection using flap hinge moment measurement. [Thesis]. University of Illinois – Urbana-Champaign; 2010. Available from: http://hdl.handle.net/2142/16147

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

13. Herges, Thomas. The effects of micro-vortex generators on normal shock wave/boundary layer interactions.

Degree: PhD, 4048, 2014, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/46573

► Shock wave/boundary-layer interactions (SWBLIs) are complex flow phenomena that are important in the design and performance of internal supersonic and transonic flow fields such as… (more)

▼ Shock wave/boundary-layer interactions (SWBLIs) are complex flow phenomena that are important in the design and performance of internal supersonic and transonic flow fields such as engine inlets. This investigation was undertaken to study the effects of passive flow control devices on normal shock wave/boundary layer interactions in an effort to gain insight into the physics that govern these complex interactions. The work concentrates on analyzing the effects of vortex generators (VGs) as a flow control method by contributing a greater understanding of the flowfield generated by these devices and characterizing their effects on the SWBLI. The vortex generators are utilized with the goal of improving boundary layer health (i.e., reducing/increasing the boundary-layer incompressible shape factor/skin friction coefficient) through a SWBLI, increasing pressure recovery, and reducing flow distortion at the aerodynamic interface plane while adding minimal drag to the system. The investigation encompasses experiments in both small-scale and large-scale inlet testing, allowing multiple test beds for improving the characterization and understanding of vortex generators. Small-scale facility experiments implemented instantaneous schlieren photography, surface oil-flow visualization, pressure-sensitive paint, and particle image velocimetry to characterize the effects of an array of microramps on a normal shock wave/boundary-layer interaction. These diagnostics measured the time-averaged and instantaneous flow organization in the vicinity of the microramps and SWBLI. The results reveal that a microramp produces a complex vortex structure in its wake with two primary counter-rotating vortices surrounded by a train of Kelvin- Helmholtz (K-H) vortices. A streamwise velocity deficit is observed in the region of the primary vortices in addition to an induced upwash/downwash which persists through the normal shock with reduced strength. The microramp flow control also increased the spanwise-averaged skin-friction coefficient and reduced the spanwise-averaged incompressible shape factor, thereby improving the health of the boundary layer. The velocity in the near-wall region appears to be the best indicator of microramp effectiveness at controlling SWBLIs. Continued analysis of additional micro-vortex generator designs in the small-scale facility revealed reduced separation within a subsonic diffuser downstream of the normal shock wave/boundary layer interaction. The resulting attached flow within the diffuser from the micro-vortex generator control devices reduces shock wave position and pressure RMS fluctuations within the diffuser along with increased pressure recovery through the shock and at the entrance of the diffuser. The largest effect was observed by the micro-vortex generators that produce the strongest streamwise vortices. High-speed pressure measurements also indicated that the vortex generators shift the energy of the pressure fluctuations to higher frequencies. Implementation of micro-vortex generators into a… Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20S.%22%29&pagesize-30">Elliott, Gregory S. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Dutton%2C%20J.%20Craig%22%29&pagesize-30">Dutton, J. Craig (advisor), Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20S.%22%29&pagesize-30">Elliott, Gregory S. (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Dutton%2C%20J.%20Craig%22%29&pagesize-30">Dutton, J. Craig (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Christensen%2C%20Kenneth%20T.%22%29&pagesize-30">Christensen, Kenneth T. (committee member).

Subjects/Keywords: shock wave boundary layer interactions (SWBLI); normal shock wave; boundary layer; supersonic flow; vortex generators; SWBLI control; shock wave unsteadiness; supersonic inlet; supersonic inlet control; inlet buzz; optical diagnostic techniques; camera housing; schlieren imaging; surface oil flow visualization; pressure sensitive paint; particle image velocimetry; shear stress measurements; format figure references in Microsoft Word

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APA · Chicago · MLA · Vancouver · CSE | Export to Zotero / EndNote / Reference Manager

APA (6^th Edition):

Herges, T. (2014). The effects of micro-vortex generators on normal shock wave/boundary layer interactions. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/46573

Chicago Manual of Style (16^th Edition):

Herges, Thomas. “The effects of micro-vortex generators on normal shock wave/boundary layer interactions.” 2014. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/46573.

MLA Handbook (7^th Edition):

Herges, Thomas. “The effects of micro-vortex generators on normal shock wave/boundary layer interactions.” 2014. Web. 25 Jan 2021.

Vancouver:

Herges T. The effects of micro-vortex generators on normal shock wave/boundary layer interactions. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2014. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/46573.

Council of Science Editors:

Herges T. The effects of micro-vortex generators on normal shock wave/boundary layer interactions. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2014. Available from: http://hdl.handle.net/2142/46573

14. Uhlig, Daniel. Micro air vehicle motion tracking and aerodynamic modeling.

Degree: PhD, 4048, 2014, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/49560

► Aerodynamic performance of small-scale fixed-wing flight is not well understood, and flight data are needed to gain a better understanding of the aerodynamics of micro… (more)

▼ Aerodynamic performance of small-scale fixed-wing flight is not well understood, and flight data are needed to gain a better understanding of the aerodynamics of micro air vehicles (MAVs) flying at Reynolds numbers between 10,000 and 30,000. Experimental studies have shown the aerodynamic effects of low Reynolds number flow on wings and airfoils, but the amount of work that has been conducted is not extensive and mostly limited to tests in wind and water tunnels. In addition to wind and water tunnel testing, flight characteristics of aircraft can be gathered through flight testing. The small size and low weight of MAVs prevent the use of conventional on-board instrumentation systems, but motion tracking systems that use off-board triangulation can capture flight trajectories (position and attitude) of MAVs with minimal onboard instrumentation. Because captured motion trajectories include minute noise that depends on the aircraft size, the trajectory results were verified in this work using repeatability tests. From the captured glide trajectories, the aerodynamic characteristics of five unpowered aircraft were determined. Test results for the five MAVs showed the forces and moments acting on the aircraft throughout the test flights. In addition, the airspeed, angle of attack, and sideslip angle were also determined from the trajectories. Results for low angles of attack (less than approximately 20 deg) showed the lift, drag, and moment coefficients during nominal gliding flight. For the lift curve, the results showed a linear curve until stall that was generally less than finite wing predictions. The drag curve was well described by a polar. The moment coefficients during the gliding flights were used to determine longitudinal and lateral stability derivatives. The neutral point, weather-vane stability and the dihedral effect showed some variation with different trim speeds (different angles of attack). In the gliding flights, the aerodynamic characteristics exhibited quasi-steady effects caused by small variations in the angle of attack. The quasi-steady effects, or small unsteady effects, caused variations in the aerodynamic characteristics (particularly incrementing the lift curve), and the magnitude of the influence depended on the angle-of-attack rate. In addition to nominal gliding flight, MAVs in general are capable of flying over a wide flight envelope including agile maneuvers such as perching, hovering, deep stall and maneuvering in confined spaces. From the captured motion trajectories, the aerodynamic characteristics during the numerous unsteady flights were gathered without the complexity required for unsteady wind tunnel tests. Experimental results for the MAVs show large flight envelopes that included high angles of attack (on the order of 90 deg) and high angular rates, and the aerodynamic coefficients had dynamic stall hysteresis loops and large values. From the large number of unsteady high angle-of-attack flights, an aerodynamic modeling method was developed… Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Selig%2C%20Michael%20S.%22%29&pagesize-30">Selig, Michael S. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Selig%2C%20Michael%20S.%22%29&pagesize-30">Selig, Michael S. (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Chung%2C%20Soon-Jo%22%29&pagesize-30">Chung, Soon-Jo (committee member), Champaign%22%20%2Bcontributor%3A%28%22Christensen%2C%20Kenneth%20T.%22%29&pagesize-30">Christensen, Kenneth T. (committee member).

Subjects/Keywords: unsteady aerodynamics; Micro air vehicle (MAV); flight dynamics; motion tracking; Vicon; modeling; separation parameter; MAV performance

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APA (6^th Edition):

Uhlig, D. (2014). Micro air vehicle motion tracking and aerodynamic modeling. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/49560

Chicago Manual of Style (16^th Edition):

Uhlig, Daniel. “Micro air vehicle motion tracking and aerodynamic modeling.” 2014. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/49560.

MLA Handbook (7^th Edition):

Uhlig, Daniel. “Micro air vehicle motion tracking and aerodynamic modeling.” 2014. Web. 25 Jan 2021.

Vancouver:

Uhlig D. Micro air vehicle motion tracking and aerodynamic modeling. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2014. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/49560.

Council of Science Editors:

Uhlig D. Micro air vehicle motion tracking and aerodynamic modeling. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2014. Available from: http://hdl.handle.net/2142/49560

15. Merrett, Craig G. Aero-servo-viscoelasticity theory: lifting surfaces, plates, velocity transients, flutter, and instability.

Degree: PhD, 4048, 2011, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/24148

► Modern flight vehicles are fabricated from composite materials resulting in flexible structures that behave differently from the more traditional elastic metal structures. Composite materials offer… (more)

▼ Modern flight vehicles are fabricated from composite materials resulting in flexible structures that behave differently from the more traditional elastic metal structures. Composite materials offer a number of advantages compared to metals, such as improved strength to mass ratio, and intentional material property anisotropy. Flexible aircraft structures date from the Wright brothers' first aircraft with fabric covered wooden frames. The flexibility of the structure was used to warp the lifting surface for flight control, a concept that has reappeared as aircraft morphing. These early structures occasionally exhibited undesirable characteristics during flight such as interactions between the empennage and the aft fuselage, or control problems with the elevators. The research to discover the cause and correction of these undesirable characteristics formed the first foray into the field of aeroelasticity. Aeroelasticity is the intersection and interaction between aerodynamics, elasticity, and inertia or dynamics. Aeroelasticity is well suited for metal aircraft, but requires expansion to improve its applicability to composite vehicles. The first is a change from elasticity to viscoelasticity to more accurately capture the solid mechanics of the composite material. The second change is to include control systems. While the inclusion of control systems in aeroelasticity lead to aero-servo-elasticity, more control possibilities exist for a viscoelastic composite material. As an example, during the lay-up of carbon-epoxy plies, piezoelectric control patches are inserted between different plies to give a variety of control options. The expanded field is called aero-servo-viscoelasticity. The phenomena of interest in aero-servo-viscoelasticity are best classified according to the type of structure considered, either a lifting surface or a panel, and the type of dynamic stability present. For both types of structures, the governing equations are integral-partial differential equations. The spatial component of the governing equations is eliminated using a series expansion of basis functions and by applying Galerkin's method. The number of terms in the series expansion affects the convergence of the spatial component, and convergence is best determined by the von Koch rules that previously appeared for column buckling problems. After elimination of the spatial component, an ordinary integral-differential equation in time remains. The dynamic stability of elastic and viscoelastic problems is assessed using the determinant of the governing system of equations and the time component of the solution in the form exp (lambda t). The determinant is in terms of lambda where the values of lambda are the latent roots of the aero-servo-viscoelastic system. The real component of lambda dictates the stability of the system. If all the real components are negative, the system is stable. If at least one real component is zero and all others are negative, the system is neutrally stable. If one or more real components are positive, the… Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Hilton%2C%20Harry%20H.%22%29&pagesize-30">Hilton, Harry H. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Hilton%2C%20Harry%20H.%22%29&pagesize-30">Hilton, Harry H. (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Ostoja-Starzewski%2C%20Martin%22%29&pagesize-30">Ostoja-Starzewski, Martin (committee member), Champaign%22%20%2Bcontributor%3A%28%22Bodony%2C%20Daniel%20J.%22%29&pagesize-30">Bodony, Daniel J. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (committee member).

Subjects/Keywords: Aeroelasticity; aero-servo-elasticity; viscoelasticity; aero-viscoelasticity; velocity transients; flutter; panel flutter; dynamic stability

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APA (6^th Edition):

Merrett, C. G. (2011). Aero-servo-viscoelasticity theory: lifting surfaces, plates, velocity transients, flutter, and instability. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/24148

Chicago Manual of Style (16^th Edition):

Merrett, Craig G. “Aero-servo-viscoelasticity theory: lifting surfaces, plates, velocity transients, flutter, and instability.” 2011. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/24148.

MLA Handbook (7^th Edition):

Merrett, Craig G. “Aero-servo-viscoelasticity theory: lifting surfaces, plates, velocity transients, flutter, and instability.” 2011. Web. 25 Jan 2021.

Vancouver:

Merrett CG. Aero-servo-viscoelasticity theory: lifting surfaces, plates, velocity transients, flutter, and instability. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2011. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/24148.

Council of Science Editors:

Merrett CG. Aero-servo-viscoelasticity theory: lifting surfaces, plates, velocity transients, flutter, and instability. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2011. Available from: http://hdl.handle.net/2142/24148

University of Illinois – Urbana-Champaign

16. Yeong, Yong Han. Wind tunnel testing of a nacelle bypass concept for a quiet supersonic aircraft.

Degree: MS, 4048, 2010, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/14621

► Sonic boom attenuation is a considerable design challenge to enable civilian aircraft to operate at supersonic flight conditions. One technology proposed by Gulfstream Aerospace Corporation… (more)

▼ Sonic boom attenuation is a considerable design challenge to enable civilian aircraft to operate at supersonic flight conditions. One technology proposed by Gulfstream Aerospace Corporation for the production of low-noise supersonic aircraft is the high-flow nacelle bypass concept in which an outer nacelle surface is used to encircle the asymmetric external engine protuberances of a traditional turbine engine. Although this bypass flow may reduce the overall sonic boom signature of the vehicle, the engine gearbox and protuberances create a highly complex 3-D flow in the annular bypass region. To better understand the 3-D flow features, an approximately 1/6th engine model was tested in a newly constructed 11.1 inch diameter axisymmetric test section of a subsonic wind tunnel at the University of Illinois at Urbana Champaign. By rotating the test section, pressure measurements were obtained over a range of circumferential angles and radial positions. The pressure measurements were used to create planar maps of nondimensionalized total and dynamic pressure upstream and downstream of the bypass model. Wind tunnel testing was performed on the empty axisymmetric wind tunnel followed by model configurations of increasing complexity until a full test configuration of the engine model with the gearbox fairing and crane beam mounts was achieved. Results show significant pressure loss behind the gearbox fairing further characterized using surface flow visualization. Due to the blockage created by the gearbox fairing mounted at the underside of the model, results also show increased flow velocity in the upper section of the bypass duct. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20S.%22%29&pagesize-30">Elliott, Gregory S. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Elliott%2C%20Gregory%20S.%22%29&pagesize-30">Elliott, Gregory S. (Committee Chair).

Subjects/Keywords: Sonic Boom; Supersonic; Nacelle Bypass

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APA (6^th Edition):

Yeong, Y. H. (2010). Wind tunnel testing of a nacelle bypass concept for a quiet supersonic aircraft. (Thesis). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/14621

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Chicago Manual of Style (16^th Edition):

Yeong, Yong Han. “Wind tunnel testing of a nacelle bypass concept for a quiet supersonic aircraft.” 2010. Thesis, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/14621.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

MLA Handbook (7^th Edition):

Yeong, Yong Han. “Wind tunnel testing of a nacelle bypass concept for a quiet supersonic aircraft.” 2010. Web. 25 Jan 2021.

Vancouver:

Yeong YH. Wind tunnel testing of a nacelle bypass concept for a quiet supersonic aircraft. [Internet] [Thesis]. University of Illinois – Urbana-Champaign; 2010. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/14621.

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

Council of Science Editors:

Yeong YH. Wind tunnel testing of a nacelle bypass concept for a quiet supersonic aircraft. [Thesis]. University of Illinois – Urbana-Champaign; 2010. Available from: http://hdl.handle.net/2142/14621

Note: this citation may be lacking information needed for this citation format:
Not specified: Masters Thesis or Doctoral Dissertation

University of Illinois – Urbana-Champaign

17. Busch, Greg T. Experimental study of full-scale iced-airfoil aerodynamic performance using sub-scale simulations.

Degree: PhD, 4048, 2010, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/14714

► Determining the aerodynamic effects of ice accretion on aircraft surfaces is an important step in aircraft design and certification. The goal of this work was… (more)

▼ Determining the aerodynamic effects of ice accretion on aircraft surfaces is an important step in aircraft design and certification. The goal of this work was to develop a complete sub-scale wind tunnel simulation methodology based on knowledge of the detailed iced-airfoil flowfield that allows the accurate measurement of aerodynamic penalties associated with the accretion of ice on an airfoil and to validate this methodology using full-scale iced-airfoil performance data obtained at near-flight Reynolds numbers. In earlier work, several classifications of ice shape were developed based on key aerodynamic features in the iced-airfoil flowfield: ice roughness, streamwise ice, horn ice, and tall and short spanwise-ridge ice. Castings of each of these classifications were acquired on a full-scale NACA 23012 airfoil model and the aero- dynamic performance of each was measured at a Reynolds number of 12.0 x 10⁶ and a Mach number = 0.20. In the current study, sub-scale simple-geometry and 2-D smooth simulations of each of these castings were constructed based on knowledge of iced-airfoil flowfields. The effects of each simulation on the aerodynamic performance of an 18-inch chord NACA 23012 airfoil model was measured in the University of Illinois 3 x 4 ft. wind tunnel at a Reynolds number of 1.8 x 10⁶ and a Mach number of 0.18 and compared with that measured for the corresponding full-scale casting at high Reynolds number. Geometrically-scaled simulations of the horn-ice and tall spanwise-ridge ice castings modeled Cl,max to within 2% and Cd,min to within 15%. Good qualitative agreement in the Cp distributions suggests that important geometric features such as horn and ridge height, surface location, and angle with respect to the airfoil chordline were appropriately modeled. Geometrically-scaled simulations of the ice roughness, streamwise ice, and short-ridge ice tended to have conservative Cl,max and Cd. The aerodynamic performance of simulations of these types of accretion was found to be sensitive to roughness height and concentration. Scaled roughness heights smaller than those found on the casting were necessary to improve simulation accuracy, resulting in Cl,max and Cd,min within 3% and 5% of the casting, respectively. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Loth%2C%20Eric%22%29&pagesize-30">Loth, Eric (committee member), Champaign%22%20%2Bcontributor%3A%28%22Selig%2C%20Michael%20S.%22%29&pagesize-30">Selig, Michael S. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Christensen%2C%20Kenneth%20T.%22%29&pagesize-30">Christensen, Kenneth T. (committee member).

Subjects/Keywords: ice accretion; iced airfoil; sub-scale simulation

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APA (6^th Edition):

Busch, G. T. (2010). Experimental study of full-scale iced-airfoil aerodynamic performance using sub-scale simulations. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/14714

Chicago Manual of Style (16^th Edition):

Busch, Greg T. “Experimental study of full-scale iced-airfoil aerodynamic performance using sub-scale simulations.” 2010. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/14714.

MLA Handbook (7^th Edition):

Busch, Greg T. “Experimental study of full-scale iced-airfoil aerodynamic performance using sub-scale simulations.” 2010. Web. 25 Jan 2021.

Vancouver:

Busch GT. Experimental study of full-scale iced-airfoil aerodynamic performance using sub-scale simulations. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2010. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/14714.

Council of Science Editors:

Busch GT. Experimental study of full-scale iced-airfoil aerodynamic performance using sub-scale simulations. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2010. Available from: http://hdl.handle.net/2142/14714

University of Illinois – Urbana-Champaign

18. Campbell, Scot E. Multi-scale path planning for reduced environmental impact of aviation.

Degree: PhD, 4048, 2010, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/16034

► A future air traffic management system capable of rerouting aircraft trajectories in real-time in response to transient and evolving events would result in increased aircraft… (more)

▼ A future air traffic management system capable of rerouting aircraft trajectories in real-time in response to transient and evolving events would result in increased aircraft efficiency, better utilization of the airspace, and decreased environmental impact. Mixed-integer linear programming (MILP) is used within a receding horizon framework to form aircraft trajectories which mitigate persistent contrail formation, avoid areas of convective weather, and seek a minimum fuel solution. Areas conducive to persistent contrail formation and areas of convective weather occur at disparate temporal and spatial scales, and thereby require the receding horizon controller to be adaptable to multi-scale events. In response, a novel adaptable receding horizon controller was developed to account for multi-scale disturbances, as well as generate trajectories using both a penalty function approach for obstacle penetration and hard obstacle avoidance constraints. A realistic aircraft fuel burn model based on aircraft data and engine performance simulations is used to form the cost function in the MILP optimization. The performance of the receding horizon algorithm is tested through simulation. A scalability analysis of the algorithm is conducted to ensure the tractability of the path planner. The adaptable receding horizon algorithm is shown to successfully negotiate multi-scale environments with performance exceeding static receding horizon solutions. The path planner is applied to realistic scenarios involving real atmospheric data. A single flight example for persistent contrail mitigation shows that fuel burn increases 1.48% when approximately 50% of persistent contrails are avoided, but 6.19% when 100% of persistent contrails are avoided. Persistent contrail mitigating trajectories are generated for multiple days of data, and the research shows that 58% of persistent contrails are avoided with a 0.48% increase in fuel consumption when averaged over a year. Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (advisor), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Neogi%2C%20Natasha%20A.%22%29&pagesize-30">Neogi, Natasha A. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Wuebbles%2C%20Donald%20J.%22%29&pagesize-30">Wuebbles, Donald J. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Coverstone%2C%20Victoria%20L.%22%29&pagesize-30">Coverstone, Victoria L. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Langbort%2C%20Cedric%22%29&pagesize-30">Langbort, Cedric (committee member).

Subjects/Keywords: contrail; aviation; environment; path planning; Mixed-integer linear programming (MILP); optimization; trajectory; air traffic management

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APA (6^th Edition):

Campbell, S. E. (2010). Multi-scale path planning for reduced environmental impact of aviation. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/16034

Chicago Manual of Style (16^th Edition):

Campbell, Scot E. “Multi-scale path planning for reduced environmental impact of aviation.” 2010. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/16034.

MLA Handbook (7^th Edition):

Campbell, Scot E. “Multi-scale path planning for reduced environmental impact of aviation.” 2010. Web. 25 Jan 2021.

Vancouver:

Campbell SE. Multi-scale path planning for reduced environmental impact of aviation. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2010. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/16034.

Council of Science Editors:

Campbell SE. Multi-scale path planning for reduced environmental impact of aviation. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2010. Available from: http://hdl.handle.net/2142/16034

University of Illinois – Urbana-Champaign

19. Lee, Sang. Large eddy simulation of shock boundary layer interaction control using micro-vortex generators.

Degree: PhD, 4048, 2010, University of Illinois – Urbana-Champaign

URL: http://hdl.handle.net/2142/14579

► The performance of supersonic engine inlets and external aerodynamic surfaces can be critically affected by shock wave / boundary layer interactions (SBLIs), whose severe adverse… (more)

▼ The performance of supersonic engine inlets and external aerodynamic surfaces can be critically affected by shock wave / boundary layer interactions (SBLIs), whose severe adverse pressure gradients can cause boundary layer separation. Currently such problems are avoided primarily through the use of boundary layer bleed/suction which can be a source of significant performance degradation. This study investigates a novel type of flow control device called micro-vortex generators (µVGs) which may offer similar control benefits without the bleed penalties. µVGs have the ability to alter the near-wall structure of compressible turbulent boundary layers to provide increased mixing of high speed fluid which improves the boundary layer health when subjected to flow disturbance. Due to their small size,µVGs are embedded in the boundary layer which provide reduced drag compared to the traditional vortex generators while they are cost-effective, physically robust and do not require a power source. To examine the potential of µVGs, a detailed experimental and computational study of micro-ramps in a supersonic boundary layer at Mach 3 subjected to an oblique shock was undertaken. The experiments employed a flat plate boundary layer with an impinging oblique shock with downstream total pressure measurements. The moderate Reynolds number of 3,800 based on displacement thickness allowed the computations to use Large Eddy Simulations without the subgrid stress model (LES-nSGS). The LES predictions indicated that the shock changes the structure of the turbulent eddies and the primary vortices generated from the micro-ramp. Furthermore, they generally reproduced the experimentally obtained mean velocity profiles, unlike similarly-resolved RANS computations. The experiments and the LES results indicate that the micro-ramps, whose height is h≈0.5δ, can significantly reduce boundary layer thickness and improve downstream boundary layer health as measured by the incompressible shape factor, H. Regions directly behind the ramp centerline tended to have increased boundary layer thickness indicating the significant three-dimensionality of the flow field. Compared to baseline sizes, smaller micro-ramps yielded improved total pressure recovery. Moving the smaller ramps closer to the shock interaction also reduced the displacement thickness and the separated area. This effect is attributed to decreased wave drag and the closer proximity of the vortex pairs to the wall. In the second part of the study, various types of µVGs are investigated including micro-ramps and micro-vanes. The results showed that vortices generated from µVGs can partially eliminate shock induced flow separation and can continue to entrain high momentum flux for boundary layer recovery downstream. The micro-ramps resulted in thinner downstream displacement thickness in comparison to the micro-vanes. However, the strength of the streamwise vorticity for the micro-ramps decayed faster due to dissipation especially after the shock interaction. In addition, the close… Advisors/Committee Members: Champaign%22%20%2Bcontributor%3A%28%22Loth%2C%20Eric%22%29&pagesize-30">Loth, Eric (advisor), Champaign%22%20%2Bcontributor%3A%28%22Loth%2C%20Eric%22%29&pagesize-30">Loth, Eric (Committee Chair), Champaign%22%20%2Bcontributor%3A%28%22Bragg%2C%20Michael%20B.%22%29&pagesize-30">Bragg, Michael B. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Christensen%2C%20Kenneth%20T.%22%29&pagesize-30">Christensen, Kenneth T. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Lee%2C%20Ki%20D.%22%29&pagesize-30">Lee, Ki D. (committee member), Champaign%22%20%2Bcontributor%3A%28%22Pantano-Rubino%2C%20Carlos%20A.%22%29&pagesize-30">Pantano-Rubino, Carlos A. (committee member).

Subjects/Keywords: supersonic; vortex generator; shock boundary layer interaction; Large Eddy Simulation (LES)

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APA · Chicago · MLA · Vancouver · CSE | Export to Zotero / EndNote / Reference Manager

APA (6^th Edition):

Lee, S. (2010). Large eddy simulation of shock boundary layer interaction control using micro-vortex generators. (Doctoral Dissertation). University of Illinois – Urbana-Champaign. Retrieved from http://hdl.handle.net/2142/14579

Chicago Manual of Style (16^th Edition):

Lee, Sang. “Large eddy simulation of shock boundary layer interaction control using micro-vortex generators.” 2010. Doctoral Dissertation, University of Illinois – Urbana-Champaign. Accessed January 25, 2021. http://hdl.handle.net/2142/14579.

MLA Handbook (7^th Edition):

Lee, Sang. “Large eddy simulation of shock boundary layer interaction control using micro-vortex generators.” 2010. Web. 25 Jan 2021.

Vancouver:

Lee S. Large eddy simulation of shock boundary layer interaction control using micro-vortex generators. [Internet] [Doctoral dissertation]. University of Illinois – Urbana-Champaign; 2010. [cited 2021 Jan 25]. Available from: http://hdl.handle.net/2142/14579.

Council of Science Editors:

Lee S. Large eddy simulation of shock boundary layer interaction control using micro-vortex generators. [Doctoral Dissertation]. University of Illinois – Urbana-Champaign; 2010. Available from: http://hdl.handle.net/2142/14579

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