Flow-thermal-structural study of aerodynamically heated leading edges

Dechaumphai, Pramote; Wieting, Allan R.; Thornton, Earl A.

doi:10.2514/3.26055

Cited by 112 publications

(37 citation statements)

References 11 publications

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“…The panel temperature was taken to be equal to the instantaneous flow temperature and was assumed to be uniform to allow for a lumpedcapacity approach to the solution of the heat transfer equations. Two related works used an explicit Taylor-Galerkin algorithm to solve the coupled fluid-thermal-structural equations to assess the impact of aerothermoelastic effects on leading edges [8] and panels [9]. These works employed an integrated finite element approach that solved the Navier-Stokes equations, energy equation, and quasi-static structural equations of motion in an integrated framework.…”

Section: Aerodynamicmentioning

confidence: 99%

Reduced-Order Aerothermoelastic Framework for Hypersonic Vehicle Control Simulation

Falkiewicz¹,

Cesnik²,

Crowell³

et al. 2011

AIAA Journal

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Hypersonic vehicle control system design and simulation require models that contain a low number of states. Modeling of hypersonic vehicles is complicated due to complex interactions between aerodynamic heating, heat transfer, structural dynamics, and aerodynamics. Although there exist techniques for analyzing the effects of each of the various disciplines, these methods often require solution of large systems of equations, which is infeasible within a control design and evaluation environment. This work presents an aerothermoelastic framework with reducedorder aerothermal, heat transfer, and structural dynamic models for time-domain simulation of hypersonic vehicles. Details of the reduced-order models are given, and a representative hypersonic vehicle control surface used for the study is described. The methodology is applied to a representative structure to provide insight into the importance of aerothermoelastic effects on vehicle performance. The effect of aerothermoelasticity on total lift and drag is found to result in up to an 8% change in lift and a 21% change in drag with respect to a rigid control surface for the four trajectories considered. An iterative routine is used to determine the angle of attack needed to match the lift of the deformed control surface to that of a rigid one at successive time instants. Application of the routine to different cruise trajectories shows a maximum departure from the initial angle of attack of 8%.corresponding to ith column of A C = correlation matrix c = modal coordinate of thermal proper orthogonal decomposition basis vector Cx = correlation model for kriging c p = specific heat at constant pressure d = structural modal coordinates, kriging sample point E = modulus of elasticity F = thermal load vector of full system in physical space f = generalized thermal load vector of reduced system in modal space F s = structural load vector of full system in physical space f s = generalized structural load vector of reduced system in modal space H i = coefficient matrices for integration of equations of motion h i = thickness of ith layer of thermal protection system K = thermal conductivity matrix of full system in physical space k = generalized thermal conductivity matrix of reduced system in modal space K G = geometric stiffness matrix K s = structural stiffness matrix k s = generalized stiffness matrix of reduced system in modal space k T = thermal conductivity of material K s = modified structural stiffness matrix L = aerodynamic lift, length M = thermal capacitance matrix of full system in physical space, Mach number m = generalized thermal capacitance matrix of reduced system in modal space M s = structural mass matrix of full system in physical space m s = generalized mass matrix of reduced system in modal space n = number of proper orthogonal decomposition snapshots n e = number of snapshots for kriging evaluation cases n k = number of kriging snapshots n s = number of structural parameters in reduced-order aerothermodynamic model n t = number of thermal parameters...

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Section: Aerodynamicmentioning

confidence: 99%

Reduced-Order Aerothermoelastic Framework for Hypersonic Vehicle Control Simulation

Falkiewicz¹,

Cesnik²,

Crowell³

et al. 2011

AIAA Journal

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“…Intersection of the compression shocks with the bow shock standing ahead of the blunt cowl lip generates severe shock interference heating [2][3][4][5], which may result in serious problems for the normal operation of an airbreathing hypersonic aircraft. Extremely high pressure and heat transfer rate gradients can occur in a small region where the interference pattern impinges onto the solid wall, and these may result in a large temperature gradient and then induce a significant thermal stress [6]. Although the bow shock is positioned ahead of the cowl lip alone, the severe aerodynamic heating can still cause a threat to the safety of the hypersonic vehicles.…”

Section: Introductionmentioning

confidence: 98%

Influence of Hypersonic Inlet Cowl Lip on Flowfield Structure and Thermal Load

Wang

Guo

2014

Journal of Propulsion and Power

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An asymmetrical hypersonic inlet cowl lip was numerically studied to explore a new approach of thermal protection. Comparative analyses of the flowfield structure and wall heat transfer with a symmetrical cowl lip were conducted. Results indicate that the peak heat transfer rate to the wall of the asymmetrical cowl lip decreases greatly when an impinging shock intersects a bow shock. Without the impinging shock, the location of the peak heat transfer rate occurs on the wall with a small radius, which does not coincide with the location of the air stagnation point; and the wall peak heat transfer rate of the asymmetrical cowl lip is higher than those of the symmetrical cowl lip, unless the flow attack angle to the wall with a large radius exceeds 12 deg. Nomenclature Ma= Mach number Q pw = peak value of the heat transfer rates to the wall, w∕m 2 Q w = heat transfer rates to the wall, w∕m 2 Q 0 = heat transfer rate at the air stagnation point at a Mach number of 8.03 without impinging oblique shock, w∕m 2 R as = radius of the large circular arc of an asymmetrical cowl lip, mm R s = radius of the symmetrical cowl lip, mm r as = radius of the small circular arc of an asymmetrical cowl lip, mm T = static temperature, K Y = ordinate of the cowl lip shape, m α = flow angle of attack, deg Φ = angular position on cylinder, deg

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“…7,8 Some researchers have developed the analysis of fluid-structure coupling. Tran et al 9 adopted the method of fluid-structure coupling to analyze the turbomachinery aeroelastic stability.…”

Section: Introductionmentioning

confidence: 99%

Multi-field coupling dynamic characteristics based on Kriging interpolation method

Yang

Yuan

Zhao

2016

Proceedings of the Institution of Mechanical Engineers, Part G:

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Multi-field coupling problems are taken more and more attention mainly because of the higher requirement of load, efficiency, and reliability in aero-engine operation. This research takes an aero-engine compressor as the research object, 3D flow field and structural models are established. For the method of cyclic symmetric, single-sector model is selected as the calculation domain. Considering the influence of former stator wakes, compressor flow field is simulated. The article analyzes the distribution law of unsteady aerodynamic load on rotor blade. Based on Kriging model, load transfer of aerodynamic pressure and temperature is achieved from flow field to blade structure. Then the effects of centrifugal force, aerodynamic pressure and temperature load are discussed on compressor vibration characteristic and structural strength. The results show dominant fluctuation frequencies of aerodynamic load on rotor blade are manly at frequency doubling of stator-rotor interaction, especially at one time frequency (1 Â f 0 ). Magnitude and pulsation amplitude on pressure surface are far greater than that on suction surface. Load transfer with Kriging model has a higher precision, it can meet the requirement of multi-field coupling dynamic calculation. In multi-field coupling interaction, temperature load makes the natural vibration frequencies decrease obviously, centrifugal force is the main source of deformation and stress. Bending stress induced by aerodynamic pressure and temperature load can counteract part of bending stress induced by centrifugal force. However, temperature load causes the maximum displacement of blade-disk system to increase.

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Flow-thermal-structural study of aerodynamically heated leading edges

Cited by 112 publications

References 11 publications

Reduced-Order Aerothermoelastic Framework for Hypersonic Vehicle Control Simulation

Reduced-Order Aerothermoelastic Framework for Hypersonic Vehicle Control Simulation

Influence of Hypersonic Inlet Cowl Lip on Flowfield Structure and Thermal Load

Multi-field coupling dynamic characteristics based on Kriging interpolation method

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