Aerothermal Loads Research Articles

Analytical Study of Articulating Turbine Rotor Blade Concept for Improved Off-Design Performance of Gas Turbine Engines

Gas turbine engines are generally optimized to operate at nearly a fixed speed with fixed blade geometries for the design operating condition. When the operating condition of the engine changes, the flow incidence angles may not be optimum with the blade geometry resulting in reduced off-design performance. Articulating the pitch angle of turbine blades in coordination with adjustable nozzle vanes can improve performance by maintaining flow incidence angles within the optimum range at all operating conditions of a gas turbine engine. Maintaining flow incidence angles within the optimum range can prevent the likelihood of flow separation in the blade passage and also reduce the thermal stresses developed due to aerothermal loads for variable speed gas turbine engine applications. U.S. Army Research Laboratory (ARL) has partnered with University of California San Diego and Iowa State University Collaborators to conduct high fidelity stator–rotor interaction analysis for evaluating the aerodynamic efficiency benefits of articulating turbine blade concept. The flow patterns are compared between the baseline fixed geometry blades and articulating conceptual blades. The computational fluid dynamics (CFD) studies were performed using a stabilized finite element method developed by the Iowa State University and University of California San Diego researchers. The results from the simulations together with viable smart material-based technologies for turbine blade actuations are presented in this paper.

Impact of Retrorocket Plumes on Upper-Stage Aerothermodynamics During Stage Separation

In this paper, an extensive investigation of the separation process of the first two stages of a carrier rocket that employs solid rocket motors for the lower stage is presented. As the reference vehicle, the VEGA rocket is used. The effect of the plumes of first-stage retrorockets on upper-stage aerodynamics and aerothermal loads is analyzed by means of wind-tunnel testing in the hypersonic wind tunnel H2K of DLR, German Aerospace Center. Aerodynamic coefficients are determined by force measurements. In addition, pressure distributions on the upper-stage surface and schlieren images for flow visualization are recorded. Infrared thermography measurements are conducted to determine the effect on aerothermal loads. Different flow conditions are achieved by variation of Reynolds number, retrorocket injection pressure ratio, and angle of attack. Results showed extensive flow separation around almost the entire upper stage by the retrorocket plumes already at low injection pressure ratios. During angle-of-attack sweeps, sudden changes in the flow structure occurred accompanied by strong changes in aerodynamic forces at values of . This behavior was found to be influenced by hysteresis effects.

Stability and Bifurcation of a Nonlinear Aero-thermo-elastic Panel in Supersonic Flow

Stability and bifurcation of a nonlinear supersonic panel under aerothermal loads are analyzed numerically in the present study. In the structural model, von Karman’s large deformation theory is taken into account for the geometric nonlinearity of the panel. In light of Hamilton’s principle, the governing equation of motion of a twodimensional aero-thermo-elastic panel is established. Coupling with the panel vibration, aerodynamic pressure is evaluated by first order supersonic piston theory and aerothermal load is approximated by quasi-steady theory of thermal stress. By transforming the partial differential equation to a series of ordinary differential equations via Galerkin method, fixed points and their stabilities of the system are studied using nonlinear dynamic theory. The complex dynamic responses regions are discussed with temperature loads as a bifurcation parameter. The results show that the thermal stress has a significant influence in the stability of the panel. The panel system undergoes Hopf bifurcation, period doubling, quasi-period and chaos with the increase of the temperature.

Aerothermal Load and Drag Force Analysis of the Electromagnetically Launched Projectiles Under Rarefied Gas Conditions

Electromagnetically launched projectiles fly with hypersonic speeds in different regions of the Earth's atmosphere. Because of their hypersonic speeds, these projectiles are designed to withstand extreme thermal loads. Drag forces should also be considered to maximize the operational range. To reduce the thermal loads and drag forces, the geometric shape of these projectiles should be carefully designed. We can utilize either experimental or numerical methods to calculate these heating effects and drag forces. Numerical methods are more economic in terms of monetary cost and time. However, we cannot use the same numerical method in different regions of the Earth's atmosphere. In this paper, we used direct simulation Monte Carlo method to calculate thermal loads and drag forces of four different projectile geometries in the rarefied part of the atmosphere. Simulation results show that thermal loads and drag forces vary considerably depending on the projectile geometry.

New Concepts of Deployable De-Orbit and Re-Entry Systems for CubeSat Miniaturized Satellites

In the present work new concepts of de-orbit and re-entry modules for standard CubeSats are presented. On the one hand, these modules can be very useful to contrast the growing phenomenon of the Space debris and, on the other hand, to recover Cubesat payload and components after the end of the mission. The concepts are mainly based on deployable, umbrella-like, structures, useful to perform de-orbit and re-entry operations taking advantage from a substantial reduction of the ballistic coefficient. In the de-orbit configuration the deployable structure works only as an aerobrake. In this case the satellite lifetime may be strongly reduced, making it possible to match the CubeSat maximum decay requirement of 25 years, even for altitudes larger than 700 km. In the Earth re-entry configuration the ballistic parameter reduction produces also a sensible decrease of the aerothermal loads on the system, giving the opportunity to recover payload, potential data and satellite components. In this case the umbrella-like structure has both the function to protect the capsule and to control de-orbit and re-entry trajectories, in order to safely reach the desired landing site. The structure can be deployed by means of preloaded springs or other actuators and, once the umbrella is opened, the reference surface can be adjusted to control the trajectory with a kinematic mechanism driven by an electro-mechanical actuator.

Flow characteristics of hypersonic inlets with different cowl-lip blunting methods

Under hypersonic flight conditions, the sharp cowl-lip leading edges have to be blunted because of the severe aerodynamic heating. This paper proposes four cowl-lip blunting methods and studies the corresponding flow characteristics and performances of the generic hypersonic inlets by numerical simulation under the design conditions of a flight Mach number of 6 and an altitude of 26 km. The results show that the local shock interference patterns in the vicinity of the blunted cowl-lips have a substantial influence on the flow characteristics of the hypersonic inlets even though the blunting radius is very small, which contribute to a pronounced degradation of the inlet performance. The Equal Length blunting Manner (ELM) is the most optimal in that a nearly even reflection of the ramp shock produces an approximately straight and weak cowl reflection shock. The minimal total pressure loss, the lowest cowl drag, maximum mass-capture and the minimal aeroheating are achieved for the hypersonic inlet. For the other blunting manners, the ramp shock cannot reflect evenly and produces more curved cowl reflection shock. The Type V shock interference pattern occurs for the Cross Section Cutting blunting Manner (CSCM) and the strongest cowl reflection shock gives rise to the largest flow loss and drag. The cowl-lip blunted by the other two blunting manners is subjected to the shock interference pattern that transits with an increase in the blunting radius. Accordingly, the peak heat flux does not fall monotonously with the blunting radius increasing. Moreover, the cowl-lip surface suffers from severe aerothermal load when the shear layer or the supersonic jet impinges on the wall.

Research on the Sufficient Condition for the Existence of Periodic Solution of FGM Subjected to Aero-Thermal Load

In this paper, the average equations are given through using the multi-scale approach method. By using the Melinkov function, the nonsingular linear transformation and the Poincaré map, the sufficient condition for existence of periodic solution of the nonlinear dynamical system about the FGM subjected to aero-thermal load is derived.

Hypersonic high altitude aerothermodynamics of a space re-entry vehicle

The analysis of the rarefaction effects in the prediction of the main aerothermal loads of a space re-entry vehicle is presented. It is well known that the Navier–Stokes equations fail in rarefied regimes and therefore other approaches must be used. In the present paper different configurations have been simulated by using the Direct Simulation Monte Carlo method. Moreover, slip flow boundary conditions have been implemented in a Navier–Stokes code in order to extend the validity of such a continuum method in the transitional flow regime. Finally, bridging-formulae for the high altitude aerodynamics of winged bodies have also been used. For the tuning of the methodologies, two simple geometries have been analysed, specifically designed to study the phenomenon of shock wave boundary layer interaction: the first one is a hollow cylinder flare, for which experimental data are also available; the second one is the geometry of a test article that was designed and tested at the Italian Aerospace Research Centre. The other two configurations that have been taken into account are an experimental winged re-entry vehicle and a capsule, for which global aerodynamic coefficients and local wall heating have been determined using different computational approaches. The Navier–Stokes code with slip flow boundary conditions has shown good predictive capabilities of the size of the recirculation bubble compared with the experimental results in the hollow cylinder flare test case; however, for the winged vehicle and capsule case, the CFD results are not fully satisfactory and the Monte Carlo method remains the most reliable approach, together with the bridging formulae, that provide good results for the global aerodynamic coefficients.

Heat transfer in a multi-layered thermal protection system under aerodynamic heating

Future generation reusable re-entry vehicles must be capable of sustaining consistent repeated aero-thermal loads without damage or deterioration. This means that such structures should be able to withstand high temperatures engendered by aero-thermal re-entry fluxes due to the establishment of a hypersonic regime over the body. Thermal Protection Systems (TPS) are used to maintain a reusable launch vehicle structural temperature within acceptable limits during re-entry flights; that is internal temperature should not be higher than the limit required by the internal structure. TPS are usually composed of several layers of different materials. Heat transfer through a multilayer insulation during atmospheric re-entry involves combined modes of heat transfer: heat conduction and radiation through the solid, heat radiation to the outer space etc. In the framework of the TPS design activities, a procedure based on one-dimensional analytical solutions of transient non-linear analysis has been developed in order to estimate the temperature variation with time and space of a multi-layered body subjected to aerodynamic heating inside a radiating space. Furthermore a semi-analytical procedure, able to take also into account non-linear thermal properties, has been conceived. Since internal temperature values of TPS of re-entry vehicles cannot exceed certain values, these procedures allow to quickly evaluate these temperature values and the preliminary size layer thicknesses without resorting to numerical solutions.

Coupled Flow-Thermal-Structural Analysis of Hypersonic Aerodynamically Heated Cylindrical Leading Edge

:A new flow-thermal-structural coupling approach based on coupled heat transfer simulation is applied to the analysis of a hypersonic aerodynamically heated leading edge. The three-dimensional unsteady flow-thermal-structural coupling is studied for Mach Number 6.47 flow over a cylinder for which the flow behavior and aero-thermal loads are calibrated by experimental data, and the numerical results of surface pressure, cold-wall heating-rate, hot-wall heating-rate and temperature distributions are all well consistent with experimental data. Interactions between the flow, structural heat transfer and deformation as well as the requirement of grid resolution for k – ε – R turbulence model are also studied.

Dynamics of Shock Dispersion and Interactions in Supersonic Freestreams with Counterflowing Jets

This study describes an active flow control concept that uses counterflowing jets to significantly modify external flowfields and strongly disperse the shock waves of supersonic and hypersonic vehicles to reduce aerothermal loads and wave drag. The potential aerothermal and aerodynamic benefits of the concepts were investigated by conducting experiments on a 2.6%-scale Apollo capsule model in Mach 3.48 and 4.0 freestreams in a trisonic blowdown wind tunnel, as well as pretest computational fluid dynamics analyses of the flowfields, with and without counterflowing jets. The model employed three sonic and two supersonic (with design Mach numbers of 2.44 and 2.94) jet nozzles with exit diameters ranging from 0.25 to 0.5 in. The schlieren images were consistent with the pretest computational fluid dynamics predictions, showing a long penetration mode jet interaction at low jet flow rates of 0.05 and 0.1 Ib m /s, whereas a short penetration mode jet was revealed at higher flow rates. The long penetration mode jet appeared to be almost fully expanded and was unsteady, with the bow shock becoming so dispersed that it was no longer discernible. High-speed camera schlieren data revealed the bow shock to be dispersed into striations of compression waves, which suddenly coalesced to a weaker bow shock with a larger standoff distance as the flow rate reached a critical value. Heat transfer results showed a significant reduction in heat flux, even giving negative heat flux for some short penetration mode interactions, indicating that the flow wetting the model had a cooling effect, instead of heating, which could significantly impact thermal protection system requirements and design. The findings suggest that high-speed vehicle design and performance can benefit from the application ofcounterflowing jets as an active flow control.

Nonlinear flutter of aerothermally buckled composite shells with damping treatments

Finite element analyses of the aerothermoelastic behavior of aerothermally buckled cylindrical composite shells with various damping treatments were implemented using the theory of layerwise displacement field. The arc-length method and iterative nonlinear scheme were taken up to estimate the post-buckling deformation of composite shells due to aerothermal loads. The complex modulus was adopted to take into account the effects of viscoelastic damping. The post-buckling behaviors and aeroelastic characteristics, under thermal loads, of several damped composite shells were investigated with various damping treatments including free layer and constrained layer and sandwiched damping layers. The results show that the flutter boundary of the cylindrical composite panels with viscoelastic layers can be remarkably affected by the aerothermally buckled shapes and damping treatments. Further, the proper damping treatments, coupled with thermal loads, can improve aerothermoelastic characteristics, resulting in the increase of the flutter boundary.

Entry Trajectory Model with Thermomechanical Breakup

a most probable cutplane. The detailed geometric and inertial modules have been coded to generate fragments of a general shape and size. The iteration over debris fragments includes complete thermal ablation and mechanical disintegration by centrifugal stresses. As soon as new debris is generated, its state vector and geometry is fed back to the dynamics and aerothermal loads modules, and simulation is performed over the next time interval. Only the largest surviving fragment is tracked for long-term simulation, whereas the other fragments are tracked for a prespecified duration. When a fragment is either completely ablated/disintegrated or reaches ground level, its simulation is ended. Hence, the scheme results in a complete simulation of the entry trajectory with en route debris shedding events until either ground impact or complete ablation and mechanical disintegration. Sample trajectory simulations for an entry capsule representative of deboost from a circular (or elliptical) orbit are considered. A dispersion analysis based on variations in the entry conditions has been conducted and its effect on the probable debris field is calculated. The speed and flight-path angle have the largest effect on downrange dispersion, whereas the heading angle affects cross-range dispersion the most. Increase in the object’s entry mass (assumed nominally distributed) has a negligible effect on the debris field, whereas a 10% decrease in the mass causes an appreciable variation in debris dispersion.

Aerothermodynamic Study of Ultrahigh-Temperature Ceramic Winglet for Atmospheric Reentry Test

This paper deals with the aerothermodynamic analysis of an advanced concept of hot structure to be investigated at atmospheric reentry conditions. The paper gives a general description of the hot structure architecture and performs an aerothermodynamic analysis to optimize the winglet configuration, based on an ultrahigh-temperature ceramics leading edge able to withstand very high temperatures. Three-dimensional fluid-dynamic computations are carried out to evaluate the aerothermal loads on the winglet. The physical model includes viscous effects, real gas properties, nonequilibrium chemical reactions and surface catalytic effects. The numerical model has been validated by experimental results of two- and three-dimensional hypersonic flowfields. A thermal model of the structure has been implemented to predict the temperature of the winglet during reentry. The results are discussed with reference to the effects of the thermal interaction with the capsule skin, the relevance of thermal conduction inside the structure, and the transient methodology. The effect of different assumptions on the ultrahigh-temperature ceramics catalytic properties is discussed.

Structural stability of functionally graded panels subjected to aero-thermal loads

Abstract Static and dynamic stabilities of functionally graded (FG) panels which are subjected to combined thermal and aerodynamic loads are investigated. The volume fractions of constituent materials composing the FG panels are assumed to be given by a simple power law distribution. Material properties of the FG panels are obtained by a linear rule of mixture. Panels are considered as rectangular plates which are based on the first-order shear deformation theory. The von Karman strain–displacement relation is used to account for geometric nonlinearity, which is caused by a large deformation. The first-order piston theory is used to simulate supersonic aerodynamic loads acting on the panels. Equations of motion are derived by the principle of virtual work and numerical solutions are obtained by a finite element method. The Newton–Raphson method is applied to get solutions of the nonlinear governing equations. Flutter boundaries are defined by linear flutter analysis and the Guyan reduction is applied to reduce degree of freedom for eigenvalue analysis. The influence of the material constitution, asymmetric characteristics of FG panels on thermal buckling and flutter characteristics are examined. Static and dynamic stability margins of FG panels are defined for various volume fractions.

4D Carbon-Carbon Composite Processing and its Properties in Comparison with Other Carbon-Carbon Preforms

The re-entry structure of a long range missile will experience severe aero thermal load during flight. The nose tip of the missile shall be capable of withstanding these severe thermal stresses arising out of thermal shock in all the directions. The material has to be so designed as to exhibit symmetric erosion and should have the lowest erosion rate with respect to all directions. So multidirectional reinforced carbon-carbon composite is an ideal choice for this application. To obtain strength in 2, 3 or 4 directions we weave fibers in different directions. Though carbon-carbon composite shows anisotropic (direction dependent) property, 4D carbon-carbon composite shows the same properties (strength, erosion rate, etc) in all 4 directions. The process involved for making 4D carbon-carbon is very tedious as well as very expensive. The properties like density, tensile strength, specific modulus, etc are studied and are very much comparable with those of the other preforms.

Projectile Nosetip Thermal Management for Railgun Launch to Space

A major issue in achieving direct launch to space using electromagnetic railguns from the surface of the earth is that we are at the bottom of a gravity well, requiring very high launch velocities to achieve orbit insertion. The atmosphere is most dense at the earth's surface-this, combined with high velocity, leads to very stressing thermal conditions for the launch aerobody. To minimize aerodynamic losses, a typical hypersonic flight body is a blunt slender cone. The thermal loads are greatest on the nosetip of this cone, and for small bodies (<10 kg) the thermal loads may exceed the capabilities of present passive thermal protection systems, such as dense, ablative carbon phenolics. Even for launch from a high-altitude airborne platform, aerothermal loads can be considerable unless rocket assist is provided. This paper discusses aspects of the launch package and characteristics of the railgun needed to achieve orbit insertion velocities

A Research Program to Study Airborne Launch to Space

Researchers supported by a multiuniversity research initiative award from the Air Force Office of Scientific Research are studying the technical issues involved in an airborne electromagnetic launch to space of small payload masses. The payload mass under consideration (1-10 kg) is much smaller than in earlier studies of launch to space, where payloads of ges1000 kg required a substantial and expensive ground-based launch facility. The lower mass now being evaluated allows all components-launcher and power supplies-to be scaled down, so that power and energy ratings and component masses are considerably reduced. To offset the aerothermal heating of the small projectile as it transits the atmosphere at >7 km/s, the entire launch system may be mounted on a large cargo aircraft and airlifted to a high altitude, where lower air density reduces the aerothermal loads to a feasible level. Such an electromagnetic launcher and pulsed power supply system could fit in a large aircraft, such as a C-5B or A-380F

Aerothermodynamic study of UHTC-based thermal protection systems

Computational Fluid Dynamics (CFD) simulations are coupled to a thermal analysis model to investigate the thermal response of the new Ultra High Temperature Ceramics (UHTC) being considered for Thermal Protection Systems (TPS) of future reusable re-entry vehicles. The numerical methodology has been applied to the Sub-orbital Re-entry Test (SRT), mission foreseen in the frame of the Italian unmanned space program. The numerical prediction of the aerodynamic heating of the “sharp” nose of the vehicle has been performed coupling the solution of the external aerodynamic flow field with the thermal field in the TPS; different methodologies have been applied, considering the effects of surface catalysis and laminar-turbulent transition. The results indicate that, for a correct prediction of the aerothermal loads on a UHTC material, coupled time-dependent simulations for both external aerodynamic flow and internal thermal field are essential. The numerical investigations also show that surface catalysis is of negligible importance for a sharp configuration in free flight, but it must be taken into account while performing ground tests in high enthalpy facilities. Furthermore, for the considered “sharp” configuration, laminar-turbulence transition effects are critical for TPS design, since they may increase TPS surface temperature of, at least, 400 K.

Computational Aerothermodynamics in Aeroassist Applications

Aeroassisted planetary entry uses atmospheric drag to decelerate spacecraft from super-orbital to orbital or sub- orbital velocities. Numerical simulation of flow fields surrounding these spacecraft during hypersonic atmospheric entry is required to define aerothermal loads. The severe compression in the shock layer in front of the vehicle and subsequent, rapid expansion into the wake are characterized by high temperature, thermo-chemical nonequilibrium processes. Implicit algorithms required for efficient, stable computation of the governing equations involving disparate time scales of convection, diffusion, chemical reactions, and thermal relaxation are discussed. Robust point-implicit strategies are utilized in the initialization phase; less robust but more efficient line-implicit strategies are applied in the endgame. Applications to ballutes (balloon-like decelerators) in the atmospheres of Venus, Mars, Titan, Saturn, and Neptune and a Mars Sample Return Orbiter (MSRO) are featured. Examples are discussed where time-accurate simulation is required to achieve a steady-state solution.