Browsing by Author "Grossman, Bernard M."
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- An actively cooled floating element skin friction balance for direct measurement in high enthalpy supersonic flowsChadwick, Kenneth Michael (Virginia Tech, 1992-12-14)An investigation was conducted to design instruments to directly measure skin friction along the chamber walls of supersonic combustor models. Measurements were made in a combustor at the General Applied Science Laboratory (GASL) and in the Direct Connect Arcjet Facility (DCAF) supersonic combustor at the NASA AMES Research Center. Flow conditions in the high enthalpy combustor models ranged from total pressures of 275-800 psia (1900-5550 kPa) and total temperatures from 5800-8400 R (3222-4667 K). This gives enthalpies in the range of 1700-3300 BTU/Ibm (3950-7660 KJ/kg) and simulated flight Mach number from 9 to 13. A direct force measurement device was used to measure the small tangential shear force resulting from the flow passing over a non-intrusive floating element. The floating head is mounted to a stiff cantilever beam arrangement with deflection due to the shear force on the order of 0.0005 in (0.0125 mm). This small deflection allows the balance to be a non-nulling type. Several measurements were conducted in cold supersonic flows to verify the concept and establish accuracy and repeatability. This balance design includes actively controlled cooling of the floating sensor head temperature through an internal cooling system to eliminate nonuniform temperature effects between the head and the surrounding chamber wall. This enabled the device to be suitable for shear force measurement in very hot flows. The key to this device is the use of a quartz tube cantilever with strain gages bonded at orthogonal positions directly on the surface at the base. A symmetric fluid flow was developed inside the quartz tube to provide cooling to the backside of the floating head. Bench tests showed that this did not influence the force measurement. Numerical heat transfer calculations were conducted for design feasibility and analysis, and to determine the effectiveness of the active cooling of the floating head. Analysis of the measurement uncertainty in cold supersonic flow tests show that uncertainty under 8% is achievable, but variations in the balance cooling during a particular test raised uncertainty up to 20% in these very hot flows during the early tests. Improvements to the strain gages and balance cooling reduced uncertainty for the later tests to under 15%.
- Adaptive finite element simulation of incompressible viscous flowFithen, Robert Miller (Virginia Tech, 1993-08-05)A finite element method is employed for solving two- and three-dimensional incompressible flows. The formulation is based on a segregated solution method. In this segregated formulation, the velocities and pressures are uncoupled and the equations for each are solved one after the other. This segregated solution method is numerically compared to the penalty method and to previous reported data to determine its validity. Next an iterative solution method which employs an element by - element data structure of the finite element method is developed. Two types of iterative methods are used. For a symmetric stiffness matrix, the conjugate gradient method is used. For an unsymmetric stiffness matrix, the bi-conjugate gradient method is used. Both iterative solution methods make use of a diagonal preconditioning method (Jacobi preconditioning). Several problems are solved using this segregated method. In two-dimensions, flow over a backward facing step and flow in a cavity are investigated. In three-dimensions, the problems include flow in a cavity at Reynolds number 100 and 1000, and flow in a curved duct. The simulation compares very well with previously reported data, where available.
- Advances In Computational Fluid Dynamics: Turbulent Separated Flows And Transonic Potential FlowsNeel, Reece E. (Virginia Tech, 1997-06-06)Computational solutions are presented for flows ranging from incompressible viscous flows to inviscid transonic flows. The viscous flow problems are solved using the incompressible Navier-Stokes equations while the inviscid solutions are attained using the full potential equation. Results for the viscous flow problems focus on turbulence modeling when separation is present. The main focus for the inviscid results is the development of an unstructured solution algorithm. The subject dealing with turbulence modeling for separated flows is discussed first. Two different test cases are presented. The first flow is a low-speed converging-diverging duct with a rapid expansion, creating a large separated flow region. The second case is the flow around a stationary hydrofoil subject to small, oscillating hydrofoils. Both cases are computed first in a steady state environment, and then with unsteady flow conditions imposed. A special characteristic of the two problems being studied is the presence of strong adverse pressure gradients leading to flow detachment and separation. For the flows with separation, numerical solutions are obtained by solving the incompressible Navier-Stokes equations. These equations are solved in a time accurate manner using the method of artificial compressibility. The algorithm used is a finite volume, upwind differencing scheme based on flux-difference splitting of the convective terms. The Johnson and King turbulence model is employed for modeling the turbulent flow. Modifications to the Johnson and King turbulence model are also suggested. These changes to the model focus mainly on the normal stress production of energy and the strong adverse pressure gradient associated with separating flows. The performance of the Johnson and King model and its modifications, along with the Baldwin-Lomax model, are presented in the results. The modifications had an impact on moving the flow detachment location further downstream, and increased the sensitivity of the boundary layer profile to unsteady flow conditions. Following this discussion is the numerical solution of the full potential equation. The full potential equation assumes inviscid, irrotational flow and can be applied to problems where viscous effects are small compared to the inviscid flow field and weak normal shocks. The development of a code is presented which solves the full potential equation in a finite volume, cell centered formulation. The unique feature about this code is that solutions are attained on unstructured grids. Solutions are computed in either two or three dimensions. The grid has the flexibility of being made up of tetrahedra, hexahedra, or prisms. The flow regime spans from low subsonic speeds up to transonic flows. For transonic problems, the density is upwinded using a density biasing technique. If lift is being produced, the Kutta-Joukowski condition is enforced for circulation. An implicit algorithm is employed based upon the Generalized Minimum Residual method. To accelerate convergence, the Generalized Minimum Residual method is preconditioned. These and other problems associated with solving the full potential equation on an unstructured mesh are discussed. Results are presented for subsonic and transonic flows over bumps, airfoils, and wings to demonstrate the unstructured algorithm presented here.
- Aerodynamic Design Sensitivities on an Unstructured Mesh Using the Navier-Stokes Equations and a Discrete Adjoint FormulationNielsen, Eric John (Virginia Tech, 1998-11-16)A discrete adjoint method is developed and demonstrated for aerodynamic design optimization on unstructured grids. The governing equations are the three-dimensional Reynolds-averaged Navier-Stokes equations coupled with a one-equation turbulence model. A discussion of the numerical implementation of the flow and adjoint equations is presented. Both compressible and incompressible solvers are differentiated, and the accuracy of the sensitivity derivatives is verified by comparing with gradients obtained using finite differences and a complex-variable approach. Several simplifying approximations to the complete linearization of the residual are also presented. A first-order approximation to the dependent variables is implemented in the adjoint and design equations, and the effect of a "frozen" eddy viscosity and neglecting mesh sensitivity terms is also examined. The resulting derivatives from these approximations are all shown to be inaccurate and often of incorrect sign. However, a partially-converged adjoint solution is shown to be sufficient for computing accurate sensitivity derivatives, yielding a potentially large cost savings in the design process. The convergence rate of the adjoint solver is compared to that of the flow solver. For inviscid adjoint solutions, the cost is roughly one to four times that of a flow solution, whereas for turbulent computations, this ratio can reach as high as ten. Sample optimizations are performed for inviscid and turbulent transonic flows over an ONERA M6 wing, and drag reductions are demonstrated.
- Aerodynamic Modeling Using Computational Fluid Dynamics and Sensitivity EquationsLimache, Alejandro Cesar (Virginia Tech, 2000-04-10)A mathematical model for the determination of the aerodynamic forces acting on an aircraft is presented. The mathematical model is based on the generalization of the idea of aerodynamically steady motions. One important use of these results is the determination of steady (time-invariant) aerodynamic forces and moments. Such aerodynamic forces can be determined using computer simulation by determining numerically the associated steady flows around the aircraft when it is moving along such generalized steady trajectories. The method required the extension of standard (inertial) CFD formulations to general non-inertial reference frames. Generalized Navier-Stokes and Euler equations have been derived. The formulation is valid for all ranges of Mach numbers including transonic flow. The method was implemented numerically for the planar case using the generalized Euler equations. The developed computer codes can be used to obtain numerical flow solutions for airfoils moving in general steady motions (i.e. circular motions). From these numerical solutions it is possible to determine the variation of the lift, drag and pitching moment with respect to the pitch rate at different Mach numbers and angles of attack. One of the advantages of the mathematical model developed here is that the aerodynamic forces become well-defined functions of the motion variables (including angular rates). In particular, the stability derivatives are associated with partial derivatives of these functions. These stability derivatives can be computed using finite differences or the sensitivity equation method.
- Aerodynamic pitch-up of cranked arrow wings: estimation, trim, and configuration designBenoliel, Alexander M. (Virginia Tech, 1994-05-06)Low aspect ratio, highly-swept cranked arrow wing planforms are often proposed for high-speed civil transports. These wing planforms offer low supersonic drag without suffering greatly from low lift/drag ratios in low-speed flight. They can, however, suffer from pitch-up at modest angles of attack (as low as 5°) during low-speed flight due to leading edge vortex influence, flow separation and vortex breakdown. The work presented here describes an investigation conducted to study past research on the longitudinal aerodynamic characteristics of highly-swept cranked wing planforms, the development of a new method to estimate pitch-up of these configurations, and the applications of this new method to the analysis of tail designs for trim at high lift coefficients. The survey of past research placed emphasis on 1) understanding the problem of pitch-up, 2) ascertaining the effects of leading and trailing edge flaps, and 3) determining the benefits and shortfalls of tail, tailless, and canard configurations. The estimation method used a vortex lattice method to calculate the inviscid flow solution. Then, the results were adjusted to account for flow separation on the outboard wing section by imposing a limit on the equivalent 2-D sectional lift coefficient. The new method offered a means of making low cost estimates of the nonlinear pitching moment characteristics of slender, cranked arrow wing configurations with increased accuracy compared to conventional linear methods. Numerous comparisons with data are included. The new method was applied to analyze the trim requirement of slender wing designs generated by an aircraft configuration optimization and design program. The effects of trailing edge flaps and horizontal tail on the trimmed lift coefficient was demonstrated. Finally, recommendations were made to the application of this new method to multidisciplinary design optimization methods.
- Aircraft Multidisciplinary Design Optimization using Design of Experiments Theory and Response Surface Modeling MethodsGiunta, Anthony A. (Virginia Tech, 1997-05-01)Design engineers often employ numerical optimization techniques to assist in the evaluation and comparison of new aircraft configurations. While the use of numerical optimization methods is largely successful, the presence of numerical noise in realistic engineering optimization problems often inhibits the use of many gradient-based optimization techniques. Numerical noise causes inaccurate gradient calculations which in turn slows or prevents convergence during optimization. The problems created by numerical noise are particularly acute in aircraft design applications where a single aerodynamic or structural analysis of a realistic aircraft configuration may require tens of CPU hours on a supercomputer. The computational expense of the analyses coupled with the convergence difficulties created by numerical noise are significant obstacles to performing aircraft multidisciplinary design optimization. To address these issues, a procedure has been developed to create two types of noise-free mathematical models for use in aircraft optimization studies. These two methods use elements of statistical analysis and the overall procedure for using the methods is made computationally affordable by the application of parallel computing techniques. The first modeling method, which has been the primary focus of this work, employs classical statistical techniques in response surface modeling and least squares surface fitting to yield polynomial approximation models. The second method, in which only a preliminary investigation has been performed, uses Bayesian statistics and an adaptation of the Kriging process in Geostatistics to create exponential function-based interpolating models. The particular application of this research involves modeling the subsonic and supersonic aerodynamic performance of high-speed civil transport (HSCT) aircraft configurations. The aerodynamic models created using the two methods outlined above are employed in HSCT optimization studies so that the detrimental effects of numerical noise are reduced or eliminated during optimization. Results from sample HSCT optimization studies involving five and ten variables are presented here to demonstrate the utility of the two modeling methods.
- Analysis of the vortical flow around a 60 degree delta wing with vortex flapSung, Bongzoo (Virginia Polytechnic Institute and State University, 1985)Subsonic wind tunnel investigations were conducted on a 60° swept, flat plate, delta wing with a leading edge vortex flap. The pressure distributions were measured over a range of angles of attack starting from zero to 40° in 5° interval and flap deflection angles from zero to 45° with 5° increments at a Reynolds number of about 2.14 x 10‘ based on the root chord. The flow visualization experiments were performed from zero degree to the stall angle, with ten different flap deflection angles at the same Reynolds number. The mean flow field was measured at angles of attack l0° and 15° with the flap deflection angles of l0° and 30° at a Reynolds number of about 1.50 x 10°. The experimental results shows that the leading edge vortex flap is an effective means to control the vortex flow over a delta wing. The optimum flap deflection angles were found where the primary vortex was confined to the leading edge vortex flap, thus producing a thrust on the flap. It was found that flap deflection could be used to restore a vortex flow from burst vortex condition.
- Analysis of Two-point Turbulence Measurements for AeroacousticsWenger, Christian W. (Virginia Tech, 1998-12-18)Simultaneous two-point three-component four-sensor hot-wire velocity measurements taken in three flows of aeroacoustic interest are here analyzed. The analyses provide information on the turbulence structure of the flows as it would be encountered by hypothetical noise producing blades passing through the flows. Two-point measurements taken in the first flow, a lifting wake from a rectangular NACA 0012 half wing, are used to calculate space-time correlation functions and 'pointwise' wave number frequency spectra. Two upwash spectra, calculated for locations in the region of the wake that is roughly homogenous in the spanwise direction, are direct estimates of the full wave number frequency spectra at their locations. As such, they are used to perform aeroacoustic calculations, and the results are compared to results achieved using the von Kármán isotropic spectrum. Amiet's approximation, where the wave number frequency spectra can be represented by the correlation length scales is found to hold reasonably well for the measured spectra. The two-point measurements in the second flow, a vortex/blade-tip interaction, are analyzed to provide information useful to researchers of blade-wake interaction noise produced by helicopter rotors. Space-time correlation functions and wave number frequency spectra are calculated for five cuts through the region of interaction. The correlation functions provide information concerning the turbulence length scales found in the interaction region. The spectra are compared to the von Kármán isotropic spectrum and found to be greatly different. However, the spectra do bear some resemblance to spectra calculated in the spanwise homogenous region of the lifting wake. The two-point measurements taken in the third flow, the wake from a fan cascade, are analyzed to provide information of use to modelers of broadband noise produced through rotor wake/stator interactions. In particular, space-time correlation functions are calculated for a grid of two-point measurements, which allows the estimation of the turbulence structure as seen by a passing stator blade. Space-time correlation functions and wave number frequency spectra are calculated for various stator configurations. The implications of engine operating speed and stator configuration for broadband noise production are discussed. [Vita removed March 2, 2012. GMc]
- An Application of Anti-Optimization in the Process of Validating Aerodynamic CodesCruz, Juan Ramón (Virginia Tech, 2003-04-04)An investigation was conducted to assess the usefulness of anti-optimization in the process of validating of aerodynamic codes. Anti-optimization is defined here as the intentional search for regions where the computational and experimental results disagree. Maximizing such disagreements can be a useful tool in uncovering errors and/or weaknesses in both analyses and experiments. The codes chosen for this investigation were an airfoil code and a lifting line code used together as an analysis to predict three-dimensional wing aerodynamic coefficients. The parameter of interest was the maximum lift coefficient of the three-dimensional wing, CL max. The test domain encompassed Mach numbers from 0.3 to 0.8, and Reynolds numbers from 25,000 to 250,000. A simple rectangular wing was designed for the experiment. A wind tunnel model of this wing was built and tested in the NASA Langley Transonic Dynamics Tunnel. Selection of the test conditions (i.e., Mach and Reynolds numbers) were made by applying the techniques of response surface methodology and considerations involving the predicted experimental uncertainty. The test was planned and executed in two phases. In the first phase runs were conducted at the pre-planned test conditions. Based on these results additional runs were conducted in areas where significant differences in CL max were observed between the computational results and the experiment — in essence applying the concept of anti-optimization. These additional runs were used to verify the differences in CL max and assess the extent of the region where these differences occurred. The results of the experiment showed that the analysis was capable of predicting CL max to within 0.05 over most of the test domain. The application of anti-optimization succeeded in identifying a region where the computational and experimental values of CL max differed by more than 0.05, demonstrating the usefulness of anti-optimization in process of validating aerodynamic codes. This region was centered at a Mach number of 0.55 and a Reynolds number of 34,000. Including considerations of the uncertainties in the computational and experimental results confirmed that the disagreement was real and not an artifact of the uncertainties.
- Assessment of a Leading Edge Fillet for Decreasing Vane Endwall Temperatures in a Gas Turbine EngineLethander, Andrew Tait (Virginia Tech, 2003-12-01)The objective of this investigation was to improve the thermal environment for a turbine vane through reduction of passage secondary flows. This was accomplished by modifying the vane/endwall junction to include a leading edge fillet. The problem approach was to integrate optimization methods with computational fluid dynamics to optimize the fillet design. The resulting leading edge fillet was then tested in a large-scale, low speed cascade to verify thermal performance. A combustor simulator located upstream of the cascade was used to generate realistic inlet conditions for the turbine vane. Both computational and experimental results underscore the importance of properly modeling the inlet conditions to the turbine. Results of the computational optimization process indicate that significant reductions in adiabatic wall temperature can be achieved with a leading edge fillet. While the intent of the initial fillet design was to improve the thermal environment for the vane endwall, computational results also indicate thermal benefit to the vane surfaces. Flow and thermal field results show that a fillet can enhance coolant effectiveness, prevent formation of the leading edge horseshoe vortex, and preclude full development of a passage vortex. In experimental testing, four cascade inlet conditions were investigated to evaluate the effectiveness of the fillet in reducing endwall temperature levels. Two tested conditions featured a flush combustor/cascade interface, while the remaining two included coolant injection through a backward-facing slot. With the flush interface, fillet thermal performance was evaluated for two inlet total pressure profiles. For the design profile, the fillet had a positive impact on the endwall temperature distribution as well as on the passage thermal field. For the off-design profile, the fillet was observed to have a slightly detrimental impact on the endwall adiabatic temperature distribution; however, passage thermal field results indicate a thermal benefit for the vane suction surface. With the backward-facing slot, thermal tests were conducted for two slot coolant flow rates. For both slot flow rates, the fillet improved endwall thermal protection and prevented coolant lift-off. While increasing the flow rate of slot coolant enhanced endwall effectiveness, fillet thermal performance was similar for the two slot flow rates.
- A CFD/CSD Interaction Methodology for Aircraft WingsBhardwaj, Manoj K. (Virginia Tech, 1997-09-15)With advanced subsonic transports and military aircraft operating in the transonic regime, it is becoming important to determine the effects of the coupling between aerodynamic loads and elastic forces. Since aeroelastic effects can contribute significantly to the design of these aircraft, there is a strong need in the aerospace industry to predict these aero-structure interactions computationally. To perform static aeroelastic analysis in the transonic regime, high fidelity computational fluid dynamics (CFD) analysis tools must be used in conjunction with high fidelity computational structural dynamics (CSD)analysis tools due to the nonlinear behavior of the aerodynamics in the transonic regime. There is also a need to be able to use a wide variety of CFD and CSD tools to predict these aeroelastic effects in the transonic regime. Because source codes are not always available, it is necessary to couple the CFD and CSD codes without alteration of the source codes. In this study, an aeroelastic coupling procedure is developed which will perform static aeroelastic analysis using any CFD and CSD code with little code integration. The aeroelastic coupling procedure is demonstrated on an F/A-18 Stabilator using NASTD (an in-house McDonnell Douglas CFD code)and NASTRAN. In addition, the Aeroelastic Research Wing (ARW-2) is used for demonstration of the aeroelastic coupling procedure by using ENSAERO (NASA Ames Research Center CFD code) and a finite element wing-box code (developed as a part of this research). The results obtained from the present study are compared with those available from an experimental study conducted at NASA Langley Research Center and a study conducted at NASA Ames Research Center using ENSAERO and modal superposition. The results compare well with experimental data. In addition, parallel computing power is used to investigate parallel static aeroelastic analysis because obtaining an aeroelastic solution using CFD/CSD methods is computationally intensive. A parallel finite element wing-box code is developed and coupled with an existing parallel Euler code to perform static aeroelastic analysis. A typical wing-body configuration is used to investigate the applicability of parallel computing to this analysis. Performance of the parallel aeroelastic analysis is shown to be poor; however with advances being made in the arena of parallel computing, there is definitely a need to continue research in this area.
- Clean Wing Airframe Noise Modeling for Multidisciplinary Design and OptimizationHosder, Serhat (Virginia Tech, 2004-07-29)A new noise metric has been developed that may be used for optimization problems involving aerodynamic noise from a clean wing. The modeling approach uses a classical trailing edge noise theory as the starting point. The final form of the noise metric includes characteristic velocity and length scales that are obtained from three-dimensional, steady, RANS simulations with a two- equation k-omega turbulence model. The noise metric is not the absolute value of the noise intensity, but an accurate relative noise measure as shown in the validation studies. One of the unique features of the new noise metric is the modeling of the length scale, which is directly related to the turbulent structure of the flow at the trailing edge. The proposed noise metric model has been formulated so that it can capture the effect of different design variables on the clean wing airframe noise such as the aircraft speed, lift coefficient, and wing geometry. It can also capture three-dimensional effects which become important at high lift coefficients, since the characteristic velocity and the length scales are allowed to vary along the span of the wing. Noise metric validation was performed with seven test cases that were selected from a two-dimensional NACA 0012 experimental database. The agreement between the experiment and the predictions obtained with the new noise metric was very good at various speeds, angles of attack, and Reynolds Number, which showed that the noise metric is capable of capturing the variations in the trailing edge noise as a relative noise measure when different flow conditions and parameters are changed. Parametric studies were performed to investigate the effect of different design variables on the noise metric. Two-dimensional parametric studies were done using two symmetric NACA four-digit airfoils (NACA 0012 and NACA 0009) and two supercritical (SC(2)-0710 and SC(2)-0714) airfoils. The three-dimensional studies were performed with two versions of a conventional transport wing at realistic approach conditions. The twist distribution of the baseline wing was changed to obtain a modified wing which was used to investigate the effect of the twist on the trailing edge noise. An example study with NACA 0012 and NACA 0009 airfoils demonstrated a reduction in the trailing edge noise by decreasing the thickness ratio and the lift coefficient, while increasing the chord length to keep the same lift at a constant speed. Both two- and three-dimensional studies demonstrated that the trailing edge noise remains almost constant at low lift coefficients and gets larger at higher lift values. The increase in the noise metric can be dramatic when there is separation on the wing. Three-dimensional effects observed in the wing cases indicate the importance of calculating the noise metric with a characteristic velocity and length scale that vary along the span. The twist change does not have a significant effect on the noise at low lift coefficients, however it may give significant noise reduction at higher lift values. The results obtained in this study show the importance of the lift coefficient on the airframe noise of a clean wing and favors having a larger wing area to reduce the lift coefficient for minimizing the noise. The results also point to the fact that the noise reduction studies should be performed in a multidisciplinary design and optimization framework, since many of the parameters that change the trailing edge noise also affect the other aircraft design requirements. It's hoped that the noise metric developed here can aid in such multidisciplinary design and optimization studies.
- A Coarse Grained Parallel Variable-Complexity Multidisciplinary Optimization ParadigmBurgee, Susan L.; Giunta, Anthony A.; Balabanov, Vladimir; Grossman, Bernard M.; Mason, William H.; Narducci, Robert; Haftka, Raphael T.; Watson, Layne T. (Department of Computer Science, Virginia Polytechnic Institute & State University, 1995-10-01)Modern aerospace vehicle design requires the interaction of multiple discipines, traditionally processed in a sequential order. Multidisciplinary optimization (MDO), a formal methodology for the integration of these disciplines, is evolving towards methods capable of replacing the traditional sequential methodology of aerospace vehicle design by concurrent algorithms, with both an overall gain in product performance and a decrease in design time. A parallel MDO paradigm using variable-complexity modeling and multipoint response surface approximations is presented here for the particular instance of the design of a high speed civil transport (HSCT). This paradigm interleaves the disciplines at one level of complexity, and processes them hierarchically at another level of complexity, achieving parallelism within disciplines, rather than across disciplines. A master-slave paradigm manages a coarse grained parallelism of the analysis and optimization codes required by the disciplines showing reasonable speedups and efficiencies on an Intel Paragon.
- A comparison of flux-splitting algorithms for the Euler equations with equilibrium air chemistryGarrett, Joseph Lee (Virginia Tech, 1989-01-05)The use of flux-splitting techniques on the Euler equations is considered for high Mach number, high temperature flows in which the fluid is assumed to be inviscid air in equilibrium. Three different versions of real gas extensions to the Steger-Warming and Van Leer flux-vector splitting, and four different versions of real gas extensions to the Roe flux-difference splitting, are compared with regard to general applicability and ease of implementation in existing perfect gas g algorithms. Test computations are performed for the M = 5, high temperature flow over a 10-degree wedge and the M = 24.5 flow over a blunt body. Although there were minor differences between the computed results for the three types of flux-splitting algorithms considered, little variation is observed between different versions of the same algorithm.
- Complex Equilibrium of Laterally Curved WakesBereketab, Semere (Virginia Tech, 1999-10-25)Turbulent wakes generated from an aircraft or submarine vehicles has been of main interest to researchers due to the broad band noise associated with such wakes. One such case is the noise generated by spiral vortices shed of from one blade interacting with another oncoming blade of helicopter rotor. Consequently, researchers have been trying to understand the basic physics and evolution of such wakes. Although there has been numerous studies done on plane wakes, there has been little research being done on laterally curved wakes. Single and two-point velocity measurements were taken on a plane and laterally curved turbulent wakes to understand the evolution and effect of lateral curvature into the far wake region. The analyses provide useful information in modeling curved or spiral wakes such as turbulence field surrounding tip vortices shed from a wing. In order to achieve our objectives, the Virginia Tech 3’ x 2’ subsonic wind tunnel was used to take velocity measurements of toroidal ring model and a straight cylinder as a control case. Velocity measurements were done using four sensor hot-wire anemometers, to obtain all mean velocity, Reynolds stress, triple product components of the turbulence field. Single point, spectra and two-point measurements of the wakes were performed throughout the development into the far wake region. The single point results reveal the universality of the mean axial velocity, however the Reynolds stresses and triple products were not universal illustrating that the turbulence field has its own length and velocity scales different from that of the mean flow. The effect of lateral curvature is mainly evidenced in the early development of the curved ring wake. The turbulent energy budget reveals similar trend for both wakes and plane wake achieves approximate equilibrium. The spectra result reveals for the plane wake that self-preservation is achieved for all scales of motion, while the ring wake does not achieve such a state. While the longitudinal correlations of both wakes are similar in form, in general difference in form and orientation prevailed over all indicating the difference in the turbulent structure of both wakes. Linear stochastic estimation reveals the presence of spanwise and double-roller eddy structures in the plane wake and only spanwise eddies were detected for the ring wake.
- Compound Aircraft Transport: Wingtip-Docked Flight Compared to Formation FlightMagill, Samantha Anne (Virginia Tech, 1998-06-29)Compound Aircraft Transport (CAT) flight involves two or more aircraft using the resources of each other; a symbiotic relationship exists consisting of a host, the mothership aircraft and a parasite, the hitchhiker aircraft. Wingtip-docked flight is just as its name implies; the two aircraft are connected wingtip-to-wingtip. Formation flight describes multiple aircraft or flying objects that maintain a pattern or shape in the air. There are large aerodynamic advantages in CAT flight. The aforementioned wingtip-docked flight increases total span of the aircraft system, and formation flight utilizes the upwash from the trailing wingtip vortex of the lead aircraft (mothership) to reduce the energy necessary to achieve and/or maintain a specific flight goal for the hitchhiker and the system. The Stability Wind Tunnel (6 X 6 X 24 foot test section) at Virginia Tech, computational aerodynamic analysis with the vortex lattice method (VLM), and a desktop aircraft model were used to answer questions of the best location for a hitchhiker aircraft and analyze stability of the CAT system. Wind tunnel tests implemented a 1/32 scale F-84E model (hitchhiker) and an outboard wing portion representing a B-36 (mothership). These models were chosen to simulate flight tests of an actual wingtip-docked project, Tom Tom, in the 1950s. That project was terminated after a devastating accident that demonstrated a possible "flapping" motion instability. The wind tunnel test included a broad range of hitchhiker locations: varying spanwise gap distance, longitudinal or streamwise distance, and vertical location (above or below wing) with respect to a B-36-like wing. The data showed very little change in the aerodynamic forces of the mothership, and possibilities of large benefits in lift and drag for the hitchhiker when located slightly aft and inboard with respect to the mothership. Three CAT flight configurations were highlighted: wingtip-docked, close formation, and towed formation. The wingtip-docked configuration had a 20-40 percent performance benefit for the hitchhiker compared to solo flight. The close formation configuration had performance benefits for the hitchhiker approximately 10 times that of solo flight, and the towed formation was approximately 8 times better than solo flight. The VLM analysis completed and reenforced the experimental wind tunnel data. A modified VLM program (VLM CAT) incorporated multiple aircraft in various locations as well as additional calculations for induced drag. VLM CAT results clearly followed the trends seen in the wind tunnel data, but since VLM did not model the fuselage, has assumptions like a flat wake, and is an inviscid computation it did not predict the large benefits or excursions as seen in the wind tunnel data. Increases in performance for the hitchhiker in VLM CAT were on the order of 3 to 4 times that of the hitchhiker in solo flight, while the wind tunnel study saw up to 10 times that of solo flight. VLM CAT is a valuable tool in supplying quick analysis of position and planform effects in CAT flight. Modifications to a desktop F-16 dynamic simulation have been developed to investigate the stability of wingtip-docked flight. These modifications analyze the stability issues linked with sideslip angle as seen by the Tom Tom Project test pilot, when he entered docking maneuvers with 5 degrees yaw to simulate a ``tired pilot". The wingtip-docked system was determined to have an unstable aperiodic mode for sideslip angle greater than 0.0 degrees and an unstable oscillatory mode for sideslip angle greater than 2.0 degrees. There is a small range of sideslip angle that is a stable oscillatory mode, sideslip angle between 0.0 and 2.0 degrees. The variables, altitude and speed, yield little effect on the stability of the system. The sensitivity analysis was indeterminate in distinguishing a state driving the instability, but the analysis was conclusive in verifying the lateral-longitudinal (roll-pitch) coupled motion observed by test pilots in wingtip-docked flight experiments. The parameter with the largest influence on the instability was the change in pitch angular acceleration with respect to roll angle. The aerodynamic results presented in this study have determined some important parameters in the location of a hitchhiker with respect to a mothership. The largest aerodynamic benefits are seen when the hitchhiker wingtip is slightly aft, inboard and below the wingtip of the mothership. In addition, the stability analysis has identified an instability in the CAT system in terms of sideslip angle, and that the wingtip-docked hitchhiker is coupled in lateral and longitudinal motion, which does concur with the divergent "flapping" motion about the hinged rotational axis experienced by the Tom Tom Project test pilot.
- Compressible turbulence in a high-speed high Reynolds number mixing layerBowersox, Rodney (Virginia Tech, 1992-09-07)Compressible turbulence in a high-speed, high Reynolds number, supersonic free shear layer was studied. A two-dimensional free mixing layer was chosen to study turbulence rather than a wall bounded flow due to the experimental fact that the effects of compressibility become significant at lower Mach numbers. The mixing layer was generated by supersonic injection of air (Ms = 1.8, Pts = 0.5 atm. Tts= 295K. and Re/m = 7x10⁶) through a rearward facing tangential slot, into a supersonic free stream (M∞ = 4.0, Pt∞ = 12.5 atm, Tt∞ = 290K, and Re/m = 70x10⁶). Flow visualization was accomplished by nanosecond Shadowgraph photography. The overall flow structure was documented with the Shadowgraph and conventional mean flow probes (Pitot pressure, cone-static pressure, and thermocouple probes). The turbulent structure of the flow field was also clearly depicted in the Shadowgraphs. Image processing techniques were developed in order to determine root-mean-square index of refraction (density) fluctuation levels from the Shadowgraph plates. Multiple overheat normal and cross-wire techniques were developed and/or improved for this study. The present research concentrated on the Reynolds averaged form of the Navier-Stokes equations. where the effects of compressibility are manifested through "apparent mass" terms (i.e. p′u′i). These terms appear in all of the Reynolds averaged Navier-Stokes equations (continuity, momentum, and energy). A new turbulence transformation, coupled with innovative experimental methods. allowed the full compressible Reynolds shear stress (the typical incompressible term, pu′iu′j as well as the apparent mass terms) to be directly measured. The full compressible heat flux and apparent mass terms were also estimated from the cross-wire results. Profiles were obtained at four downstream stations which were strategically located to map different levels of development of the shear flow. The first station was very close to the injector, about one free stream boundary layer thickness downstream (x/δ∞ ≈ 1), hence, it is in the initial region. The second station was located at x/δ∞ ≈ 28, which was near the beginning of the fully developed zone. The third station, x/δ∞ = 83, was just prior the shear layer and floor boundary layer merging. The last station was positioned just aft of the layer merging, x/δ∞ = 106. Reynolds averaging of the compressible Navier-Stokes equations implies that the compressible turbulence affects all of the governing equations. It was found, experimentally, that the effects of compressibility on turbulence were more than significant accounting for about 75% of the total level of the Reynolds shear stress formulation for the present study (i.e. the apparent mass term multiplied by the axial velocity was about 3-4 times the typical incompressible shear term). For the present mean adiabatic flow, the compressible turbulence accounted for 100% of the turbulent heat flux. The apparent mass in the continuity equation was, by definition, only due to compressibility. These results led to the development of anew Compressible Apparent Mass Mixing Length Extension (CAMMLE) model that accounts for compressible turbulence in all of the governing equations (i.e. the turbulence terms in the continuity, momentum, and energy were all consistently formulated). The CAMMLE formulation is a generalization of the Situ-Schetz compressible mixing length formulation, which was developed to account for the apparent mass terms in the momentum equation. A total of seven turbulence models were experimentally evaluated, the CAMMLE model, the Prandtl incompressible and the Situ-Schetz compressible mixing length models, the Prandtl and Bradshaw turbulent kinetic energy (TKE) formulations, and two compressible TKE extensions that are based upon a newly defined compressible TKE formulation. The measured turbulence data was used to assess the various models, where the measured mean flow profiles were used in the model formulations. The incompressible formulations were generally successful in representing the measured incompressible part of the Reynolds shear stress. However, this term only accounted for about 25% of the total shear stress level. All of the compressible extensions provided accurate estimates of the full compressible Reynolds shear stress. In addition, the newly developed CAMMLE model was also successful in representing the apparent mass terms in the continuity equation. The CAMMLE model was also the only formulation to accurately predict the measured compressible turbulent heat flux in the energy equation. The CAMMLE, Situ-Schetz, and Prandtl incompressible mixing length models were all incorporated in to a 3-D finite volume Navier-Stokes code (GASP 2.0). The numerical simulations indicated that the new compressible apparent mass mixing length extension performed very well. The CFD results also enlightened a misuse with all of the current compressible turbulence models. With the exception of the new apparent mass formulation, all existing turbulence models neglect the compressible turbulence effects on the continuity equation and treat the energy equation in an ad hoc effective eddy viscosity and thermal conductivity fashion. The numerical and theoretical studies indicated that this led to poor prediction of the mixing layer width for cases where the free stream Mach number was significantly higher than the injection Mach number.
- Computational Simulations of a Non-body of Revolution Ellipsoidal Model Utilizing RANSSomero, John Ryan (Virginia Tech, 2010-08-06)The ability of Reynolds Averaged Navier Stokes (RANS) models to predict the characteristics of a non-Body of Revolution (non-BOR) Ellipsoidal model is studied to establish the feasibility of utilizing RANS as a non-BOR concept design tool. Data unable to be obtained experimentally, such as streamwise and spanwise pressure gradients and yaw turn boundary layer characteristics, are also established. A range of conditions are studied including ahead, pitched up, steady 10 and 15 degree yaw turns, and unsteady 10 and 15 degree yaw turns. Simulation results show good agreement for ahead and pitched forces and moments. Straight ahead skin friction values also showed good agreement, providing even improved agreement over an LES model which utilized wall functions. Yaw turn conditions also showed good agreement for roll angles up to 10 degrees. Steady maneuvering forces and moments showed good agreement up to 10 degrees roll and separation calculations also showed good agreement up to 10 degrees roll. Unsteady maneuvering characteristics showed mixed results, with the normal force and pitching moment trends generally agreeing with experimental data, whereas the unsteady rolling moment did not tend to follow experimental trends. Two primary conditions, the change in curvature between the mid-body and elliptical ends and the accuracy of modeling of 3D flows with RANS, are discussed as sources of discrepancies between the experimental data and steady simulations greater than 10 degrees roll and unsteady rolling simulations.
- Dense gas effects in a converging-diverging nozzleFry, Richard N. (Virginia Tech, 1992-12-15)Analytical and computational models for the solution of steady inviscid flows in a converging-diverging nozzle are presented for a general fluid. The main emphasis is placed on Bethe-Zel'dovich-Thompson fluids, i.e., those having specific heats so large that the fundamental derivative of gasdynamic is negative over a finite range of pressures and temperatures. Three general classes of flow are delineated which include two nonclassical types in addition to the usual classical flows; the latter are qualitatively similar to those of a perfect gas. The nonclassical flows are characterized by isentropes containing as many as three sonic points. Numerical solutions depicting finite strength expansion shocks, steady flows with shock waves standing upstream of the nozzle throat, and steady flows containing as many as three shock waves are presented. Nonclassical flows having arbitrarily large exit Mach numbers can be obtained only if a sonic expansion shock is formed in the nozzle.