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Aero-Thermodynamics for Conceptual Design
Abstract
A software tool for the prediction of the aero -thermodynamic environments of conceptual aerospace configurations is presented. The vehic le geometry is defined using unstructured, triangulated surface meshes. For subsonic Mach numbers a fast, unstructured, multi -pole panel code is coupled with a streamline tracing formulation to define the viscous surface solution. For supersonic and hypers onic Mach numbers, various independent panel methods are coupled with the streamline tracing formulation, an attachment line detection method , and stagnation -attachment line heating models to define the viscous aero -thermal environment.
... Additionally, nonlinear vortex lift was estimated using empirical correlations. The second tool is CBAERO [12], which includes supersonic/hypersonic panel methods and was used for Mach numbers of three and greater. ...
... Estimates for the aerodynamic properties at Mach numbers of three and greater were generated using CBAERO [12]. CBAERO includes methods for both subsonic and supersonic/hypersonic aerodynamic analysis. ...
... This process was done by curve fitting both the engineering analysis and CFD results. The lift coefficients were fit as a function of angle of attack using the relationship (12) and drag coefficients were fit as a function of lift coefficient using the relationship ...
... In the hypersonic vehicle design, engineering-level lowfidelity tools rather than high-fidelity CFD codes have been widely used for the calculation of aerodynamic forces and moments of hypersonic configurations due to their fast turn-around time and acceptable accuracy for early design phases [3][4][5][6]. The engineering approach only requires surface mesh on a configuration, and, therefore, volume mesh generation is not needed. ...
... Hypersonic engineering codes using local surface inclination methods have been developed and widely used for hypersonic vehicle design such as reentry spacecrafts and hypersonic cruise vehicles [4][5][6]. The HABP (Hypersonic Arbitrary Body Program) [4] has been widely used in related industries for decades, and it adopts quadrangular structural surface meshes. ...
... The HABP (Hypersonic Arbitrary Body Program) [4] has been widely used in related industries for decades, and it adopts quadrangular structural surface meshes. Recently developed hypersonic engineering codes [5,6] are based on unstructured triangular surface meshes for easy handling of complex geometries and more efficient surface mesh generation. In Ref. [6] the FAAT code was developed implementing three distinct local surface inclination methods. ...
An efficient aerodynamic analysis was conducted for rapid prediction of the aerodynamic performance of air-breathing hypersonic vehicles in the early design phase. Surface pressure calculation on compression surfaces was performed using local surface inclination methods such as modified Newtonian, tangent cone, and tangent wedge methods. The Prandtl–Meyer expansion method was adopted for surface pressure calculation on expansion surfaces. Sharma’s three-dimensional correction is applied to the Prandtl–Meyer expansion if needed. The reference temperature method was used for the prediction of the local skin friction coefficient. An efficient approximate streamline tracing method was suggested to calculate the reference length for each local surface cell with just one percent of the CPU time required for the original naive stream tracing method. An internal flow passage for air-breathing propulsion was also considered under the isentropic assumption for prediction of aerodynamic performance of a whole vehicle. Comparisons are made for aerodynamic coefficients calculated by the present code and a high fidelity CFD code for a X-51A-like configuration at flight conditions of Mach 5 and 6, and good agreements were obtained. The developed code is very efficient and possesses enough accuracy for engineering prediction of the aerodynamic performance of hypersonic air-breathing vehicles.
... Similar stagnation heat flux prediction methods also include the Kemp-Riddell [15,16] method, the Lees method [17], the Scala method [18], and the Romig method [19]. Besides predicting stagnation heat flux, Kinney [20], Hamilton et al. [21], and Lee and Wurster [22] predicted the heat flux distribution over hypersonic vehicles efficiently by solving the Euler equation outside the boundary layer and performing viscous correction inside the boundary layer to avoid solving full NS equations. By using the Newton cooling formula, McNamara et al. [23] could obtain the heat flux distribution with different wall temperatures immediately. ...
... represents the total energy per unit volume of fluid, and V V ⋅ n is the inversion velocity on the interface of the control body. The linear shear stress tensor τ ij could be expressed as 20) and the linear heat flux term is obtained by the Fourier's law as ...
The efficient and accurate prediction of the aeroheating performance of hypersonic vehicles is a challenging task in the thermal protection system structure design process, which is greatly affected by grid distribution, numerical schemes, and iterative steps. From the inspiration of the theoretical analysis and machine learning strategy, a new wall heat flux prediction framework is proposed first by establishing the relationship between the wall heat flux and the flow variables at an extreme temperature point (ETP) in the normal direction of the corresponding wall grid cell, which is named the machine learning (ML)-ETP method. In the training phase, the flow properties and their gradients at the ETP and the distance from the ETP normal to the wall are employed as feature values, and the accurate wall heat flux predicted by the converged fine grid is regarded as the tag value. With the assistance of the trained regression model, the heat flux of the same configuration with a coarse grid in the wall-normal direction could be predicted accurately and efficiently. Moreover, test cases of different configurations and inflow conditions with a coarse grid are also carried out to assess the model’s generalization performance. All comparison results demonstrate that the ML-ETP strategy could predict wall heat flux more rapidly and accurately than the traditional numerical method due to its nonstrict grid distribution requirements. The improvement of the predictive capability of the coarse-graining model could make the ML-ETP method an effective tool in hypersonic engineering applications, especially for unsteady ablation simulations or aerothermal optimizations.
... In the 21st century, the Ames Center developed the Advanced Engineering Environment (AEE) [3] to connect data from different disciplines, which could automate the multidisciplinary design and analysis of spacecraft. And based on AEE, NASA developed MISSION software [4] for orbit design and CBAero software [5] to generate the aerothermal database for aerothermal analysis, which played a huge role in promoting the design of thermal protection of early aircraft. ...
... In the automatic material selection module of the engineering software, the maximum temperature of the outer surface of the structure at each structural point is calculated according to formula (4), and then the material of each design point will be selected by formula (5). The extreme temperature of the material and other material properties are included in the input material library. ...
During the flight mission of hypersonic aircraft, severe aerodynamic heating will occur on the surface, so thermal protection system (TPS) is required to protect the load-bearing structure of the aircraft. The present paper develops an engineering software for automatic optimization of the thickness of tile-type TPS for reusable aircraft. For requirements on TPS of reusable aircraft in the reentry stage, the method of heat flow-time curve enveloping, automatic material selection, and one-dimensional unsteady heat transfer calculation for multilayer plates under thermal load conditions had been researched, an interactive engineering software had been developed. The software improves the calculation accuracy and calculation efficiency of TPS thickness optimization, and it is suitable for rapid design in the conceptual design stage of the aircraft. Finally, by an example, the function of the software is verified.
... Thus, the key for getting the heat transfer and, as a consequence, the aerodynamics correct is getting the air chemistry right. Unfortunately, as mentioned above, only limited experimental data with a high degree of uncertainty is available at the high temperatures and conditions (e.g., behind the shock) of interest, and because of this, the assumption of chemical equilibrium is often times made in regions, which in reality are very far from equilibrium (e.g., [3][4][5][6][7][8][9]). Even if in some cases the chemical equilibrium approximation is justified such as when the wall is catalytic and the Lewis number near unity [4] (itself an approximation to simplify the governing equations), this approximation is in general not valid. ...
... Clearly, despite the improved fit provided by the two-temperature model, knowledge of these distributions at each instant of time is beyond the scope of standard engineering design models and codes (e.g., [3][4][5][6][7][8]22]). These models are, however, part and parcel of high-fidelity computational fluid dynamics (CFD) models such as, for example, found in [23,24]. ...
... Vehicle properties are calculated using CBAERO, a modified-Newtonian solver that has been shown to generate accurate aerodynamic coefficients for geometries in hypersonic flow fields [22]. CBAERO solves the Fay-Riddell equations [23] to obtain wall temperature and heat-flux maps across a geometry surface. ...
This investigation presents recent efforts to develop a computational toolset and analysis framework capable of modeling the path-dependent change of vehicle shape due to ablation during high-ablation atmospheric entry maneuvers. The framework uses Engineering Sketch Pad (ESP) to model geometry deformations based on recession rates computed using Configuration Based Aerodynamics (CBAERO) and Fully Implicit Ablation and Thermal Analysis Program (FIAT). Shape change due to recession is handled using free-form deformation in ESP with spline deformations calculated using a design parameter matching algorithm provided by Parametric Legacy Unstructured Geometry System (PLUGS). This method enables modeling shape change without the risk of mesh tangling or other issues invoked by direct deformation of a surface mesh. Shape deformation is calculated using flow conditions computed along an input trajectory. The toolset is demonstrated via a notional proof of concept using sample low- and mid-lift to drag ratio (L/D) vehicle geometries. Results indicate that the analysis framework can model the pseudo-arbitrary shapes that develop as a result of geometry recession due to ablation. Further, a preliminary analysis shows that geometry recession due to ablation has a stronger effect on a mid-L/D entry shape than on a low-L/D blunted entry capsule.
... The Taylor Maccoll equation is further modified [13] to accommodate its use for analysing this equation for the shockcone interaction between the super-hypersonic wings and conical flow of various shock angles which further helped in identifying the influence of wave drag and the skin friction drag between the interaction of shockcone and the boundary layer of the configuration overall improving its application in studying the velocity potential at these speeds [14,15]. The recent use of this for various advances in hypersonic and supersonic modelling is investigated [10,[16][17][18][19]. The particular use of the vortex lattice method (VLM) being quite popular low-order method hinges on specific influence terms [20], which dictate the formation and movement of horseshoe vortices across the aircraft's surface. ...
In the realm of supersonic design, obtaining data for numerous supersonic configurations amidst intricate flow conditions proves time-consuming due to the excessive costs associated with high-fidelity computational demands. Running iterative simulations over an extended period is often impractical or entails substantial expenses. This inherent challenge necessitates the adoption of low-order potential solvers with reasonable accuracy to generate datasets. In support of this objective, This study addresses the high computational costs of obtaining data for supersonic configurations by developing a low-order solver that combines the Taylor-Maccoll hypervelocity method (TMHM) with the supersonic vortex lattice method. This approach aims to provide accurate drag predictions in supersonic flows while minimizing computational demands. By integrating TMHM to calculate wave drag and skin friction drag and enhancing the vortex lattice method to handle shockwave impacts through panel matching, the solver achieves improved accuracy in lift and drag computations. Validation against experimental data shows a 20% reduction in drag prediction error compared to traditional vortex lattice methods, with a 2.01% error for low-shock angles. The method achieves accuracy rates between 90% and 95% across various configurations, including a 90% accuracy for delta wings, 85% for positive dihedral wings, and 95% for large sweptback angle designs, as confirmed by comparisons with high-fidelity CFD data.
... The dataset includes 4940 total steady-state simulations, each at different operating conditions. From the NASA website [53]: "CBAERO is a software tool for the prediction of the conceptual aero-thermodynamic environments of aerospace configurations. The vehicle geometry is defined using unstructured, triangulated surface meshes. ...
High-speed flight vehicles, which travel much faster than the speed of sound, are crucial for national defense and space exploration. However, accurately predicting their behavior under numerous, varied flight conditions is a challenge and often prohibitively expensive. The proposed approach involves creating smarter, more efficient machine learning models (also known as surrogate models or meta models) that can fuse data generated from a variety of fidelity levels -- to include engineering methods, simulation, wind tunnel, and flight test data -- to make more accurate predictions. These models are able to move the bulk of the computation from high performance computing (HPC) to single user machines (laptop, desktop, etc.). The project builds upon previous work but introduces code improvements and an informed perspective on the direction of the field. The new surrogate modeling framework is now modular and, by design, broadly applicable to many modeling problems. The new framework also has a more robust automatic hyperparameter tuning capability and abstracts away most of the pre- and post-processing tasks. The Gaussian process regression and deep neural network-based models included in the presented framework were able to model two datasets with high accuracy (R^2>0.99). The primary conclusion is that the framework is effective and has been delivered to the Air Force for integration into real-world projects. For future work, significant and immediate investment in continued research is crucial. The author recommends further testing and refining modeling methods that explicitly incorporate physical laws and are robust enough to handle simulation and test data from varying resolutions and sources, including coarse meshes, fine meshes, unstructured meshes, and limited experimental test points.
... Sharp edges on a body are not preferred since they create associated shock waves that, despite having less drag, cause less heat to dissipate into the environment (due to a thin shock layer). Conversely, a blunt shape body is recommended since it creates a disconnected shock wave termed as bow shock [2]. More area in the atmosphere is thus provided for the dispersion of heat. ...
It is important to understand the aerothermodynamics besides aerodynamics to design a high speed vehicle. A body travelling at high speed could generate enough heat to cause it to burst apart. Thus, a blunt nose design is used for a simple re-entry vehicle in order to achieve a detached shockwave (bow shock) and so dissipate more heat into the surrounding atmosphere. For blunt bodies versus sharp-edged, the ballistic coefficient is lower. The vehicle slows down higher in the atmosphere as a result, which lessens the aero-thermal loads it encounters. A basic design of existing Orion-based crew exploration vehicle travelling in a low hypersonic speed (Mach 6) is considered owing to the feasibility of available computational support. Computational fluid dynamics (CFD) study consisting of both qualitative and quantitative analyses has been done to further reduce the aerodynamic heating of the capsule. Ansys Fluent was utilized to capture the heat flux over the surface, velocity profiles, pressure, and temperature distribution surrounding the capsule. To reduce the aerodynamic heating while preserving other parameters (such as lift and drag) and a suitable ballistic coefficient, a variety of design enhancements can be made. Concave heat shield, a revolutionary design advancement that reduces aerodynamic heating, was therefore investigated in this work. A significant decrease in the aerodynamic heating is observed for the proposed design provided with a slight increment in drag which in turn is favourable.
... These codes have incorporated advanced methodologies, such as the axisymmetric analogue, Lees' formulations, and DeJarnette, and Davis's approximate techniques for streamline distribution [10,11]. Industry codes like HABP [12], MINIVER [13], AEROHEAT [14], INCHES [15], CBAERO [16], and HATLAP [17] have enhanced the fidelity of calculations for surface streamlines, pressure, and heating formulations, often comparing favorably with more complex simulations, such as viscous shock layer models and comprehensive CFD simulations. Notably, AEROHEAT includes a unique equivalent boundary layer method that involves a direct solution of a boundary layer specific to each streamline [10]. ...
In this research, a streamlined numerical approach designed for the quick estimation of temperature profiles across the finite thickness of a hemispherical dome subjected to aerodynamic heating is introduced. Hemispherical domes, with their advantageous aerodynamic, structural, and optical properties, are frequently utilized in the front sections of objects traveling at supersonic velocities, including missiles or vehicles. The proposed method relies on one-dimensional analyses of fluid dynamics and flow characteristics to approximate the local heat flux across the exterior surface of the dome. By calculating these local heat flux values, it is also possible to predict the temperature variations within the thickness of the dome by employing the finite difference technique, to solve the heat conduction equation in spherical coordinates. This process is iterated over successive time intervals, to simulate the entire flight duration. Unlike traditional Computational Fluid Dynamics (CFD) simulations, the proposed strategy offers the benefits of significantly lower computational time and resource demands. The primary objective of this work is to provide an efficient numerical tool for evaluating aerodynamic heating impact and temperature gradients on hemispherical domes under specific conditions. The effectiveness of the proposed method will be validated by comparing the temperature profiles derived for a standard flight scenario against those obtained from 2-D axisymmetric transient CFD simulations performed using ANSYS-Fluent 2022 R2.
... non-constant) singularity elements on the panels. Owing to their low computational cost, lifting-line and lifting-surface methods remain in use for conceptual design and analysis frameworks for fixed-wing aircraft [24,25], high-Mach number systems [26], wind turbines [27], propellers [28], rotorcraft comprehensive analyses [29][30][31][32] and real-time flight simulation [33]. Incorporating additional physics and design considerations remain an open research area, such as aerostructural optimisaton [34] and unsteady aerodynamics [35,36]. ...
The accuracy of several numerical schemes for solving the lifting-line equation is investigated. Circulation is represented on discrete elements using polynomials of varying degree, and a novel scheme is introduced based on a discontinuous representation that permits arbitrary polynomial degrees to be used. Satisfying the Helmholtz theorems at inter-element boundaries penalises the discontinuities in the circulation distribution, which helps ensure the solution converges towards the correct, continuous behaviour as the number of elements increases. It is found that the singular vorticity at the wing tips drives the leading-order error of the solution. With constant panel widths, numerical schemes exhibit suboptimal accuracy irrespective of the basis degree; however, driving the width of the tip panel to zero at a rate faster than the domain average enables improved accuracy to be recovered for the quadratic-strength elements. In all cases considered, higher-order circulation elements exhibit higher accuracy than their lower-order counterparts for the same total degrees of freedom in the solution. It is also found that the discontinuous quadratic elements are more accurate than their continuous counterparts while also being more flexible for geometric representation.
... These estimations are performed to determine the seriousness of the problem and identify the most critical conditions of flight. Many engineering estimation tools have been developed to compute aerodynamic heating such as AEROHEAT (Quinn and Gong, 1990), MINIVER (Louderback, 2013), TPATH (Quinn and Gong, 2000), INCHES (Hamilton et al., 1993), LATCH (Hamilton et al., 1994), HABP (Smyth and Loo, 1981), CBAERO (Kinney, 2004) and HATLAP (Jain and Hayes, 2004). ...
Validation of one-dimensional aerodynamic heating and ablation prediction program, AeroheataBS to calculate transient skin temperatures and heat fluxes for high-speed vehicles has been performed. In the tool shock relations, flat plate convective heating expressions, Eckert's reference temperature method and modified Newtonian flow theory are utilized to compute local heat transfer coefficients. Corresponding governing equations are discretized explicitly and numerically solved. Time varying flight conditions including velocity, altitude and angle of attack serve as input to the program. In order to examine the accuracy of aerodynamic heating capabilities of AeroheataBS, calculated temperature histories are compared with flight data of the X-15 research vehicle, a modified von-Karman nose shaped body, cone-cylinder-flare configuration and results of conjugate computational fluid dynamics studies. Comparative studies show that computed values are in good agreement with the reference data and prove that methodology established in AeroheataBS is appropriate for estimating aerodynamic heating and structural thermal response.
... Above Mach 10, the Configuration-Based Aerodynamics (CBAERO) program was used. CBAERO simulations of the HSRV use modified Newtonian methods for pressure loads and ignore viscous effects [4]. For Mach 10 down to Mach 2, Cart3D was used. ...
... The specific heat ratio, γ, was obtained from Reference 26. Figure 21 shows the resulting lift and drag coefficients for each case after drag skirt jettison. These results are plotted against values from CBAERO, a Modified-Newtonian inviscid flow solver [27], for validation. Both software packages produce similar coefficients, although Cart3D predicts lower values for CD and less variability in CL than CBAERO. ...
A new approach is proposed of adaptively selecting emulators for emulator embedded neural networks. Emulators typically take the form of physics-based low-fidelity models. If querying a low-fidelity model incurs significant computational costs, an emulator can still be constructed by training a surrogate model, such as a Gaussian process model or neural network, on data gathered from the low-fidelity model. These emulators are embedded into the neural network architecture and play an important role in boosting neural network accuracy with limited training data. However, in some practical situations finding a suitable low-fidelity emulator model is challenging. Use of an improper emulator can fail to improve learning performance, and in some cases no viable emulator is available. To address this technical gap, this study proposes using the decomposed functions of the actual physics-based model as emulators. Global sensitivity analysis is performed to compute the Sobol indices of the decomposed functions, which reveal their contribution ranks to the system response of interest. Based on the contribution ranks, the decomposed functions are embedded as emulators in an adaptive and iterative process. The proposed method is demonstrated with fundamental analytical examples. Additionally, a representative hypersonic vehicle design problem is included as a practical engineering example.
View Video Presentation: https://doi.org/10.2514/6.2022-2288.vid The Hypersonic foldable Aeroshell for THermal protection using ORigami (HATHOR) is a rigid mechanically deployable heat shield, a versatile and promising alternative to traditional rigid aeroshells for next generation space missions. Given its unconventional (faceted) deployed shape, a new analytical tool is required for its aerothermal analysis. HEAT-3D is an engineering level code for the estimation of preliminary heat transfer rates and heat loads experienced by a 3D heat shield during entry. An extensive validation of heating and temperature estimates from HEAT-3D is presented and an accuracy of 10% is demonstrated for laminar flow at zero incidence. HEAT-3D is employed for the TPS sizing of a 2.65 m technology demonstrator of HATHOR that has been designed and built at Imperial College London. Reusable and ablative TPS options are considered and a 14 mm layer of the ablator SLA-561V is chosen for this purpose.
View Video Presentation: https://doi.org/10.2514/6.2022-1304.vid Recent advances in highly efficient algorithms and high-performance computing allow the construction of an integrated design framework wherein the traditionally segregated disciplines of airframe design and trajectory design are coupled together in order to undertake the design and optimization of hypersonic vehicles as integrated systems. The particular interest in this paper is the potential approach to incorporating high-fidelity aerodynamic models in the hypersonic trajectory optimization problem, incrementally varying the geometric parameters of the vehicle to observe induced performance variations, and employing Bayesian optimization and machine learning algorithms to optimize the vehicle geometry for specified mission profiles. First, the exigency for considering high-fidelity aerodynamic models is justified. Then energy-based problem formulations for hypersonic trajectory optimization are introduced. A panel method based on the modified Newtonian flow theory and Eckert’s reference model is used to produce high-fidelity aerodynamic force and heating coefficients, based on which a pseudospectral optimal control package is used to solve for optimal trajectories. Finally, an iterative procedure employing the Bayesian optimization and machine learning is established to successively search for the geometry that enables the optimal mission performance. Preliminary results demonstrate the feasibility and advantage of the developed approach.
View Video Presentation: https://doi.org/10.2514/6.2022-0023.vid Simulation approaches for hypersonic vehicles have been around for decades, with many of the early semi-empirical methods being drawn from theoretical and experimental knowledge developed in the 1950s and 1960s. Some of those approaches have stood the test of time and can do a reasonable job predicting pressures at high Mach numbers, but the prediction of other parameters, such as heat transfer, have always been challenging. Over time, CFD has become the simulation approach most typically used for hypersonic vehicles, but typical CFD codes developed for lower Mach regimes require the inclusion of multiple physical phenomena that are not currently modeled appropriately, or are not included at all, in many simulation codes. These physical aspects include: combustion physics, turbulence and transition, fluid-thermal-structural interactions, non-equilibrium chemistry, shock-boundary layer interactions, and ablation. Without the inclusion of these aspects of hypersonic flight (either through physics-based prediction or appropriate modeling), simulation of hypersonic vehicles is deficient and cannot properly provide full support to the development and acquisition of hypersonic vehicles. In order to improve the national capability in hypersonic vehicle RDT&E, a Hypersonic Vehicle Simulation Institute has been established by the DoD High Performance Computing Modernization Program at the US Air Force Academy. The institute is following a multi-pronged, multi-year approach to address these shortcomings in hypersonic vehicle simulation. Details about the progress will be included, such as the various turbulence and transition modeling projects that are taking place, and the various validation experiments that are being supported. How these projects are interacting with various government labs and the aerospace industry will also be discussed.
This paper introduces the Hypersonic Engineering model for rapid AeroThermal analysis of 3D bodies (HEAT-3D). This robust tool has been developed for the rapid heat transfer analysis of generic (i.e. not axisymmetric) 3D bodies. A streamline tracing algorithm is used in conjunction with modified Newtonian theory (to determine the inviscid solution), coupled with a range of empirical engineering relationships to predict the boundary layer flow and estimate laminar and turbulent convective heat transfer rates. HEAT-3D has been validated against existing experimental results and it is shown to approximate heat transfer rates to within 20%, in excellent agreement with significantly more complex and computationally expensive CFD codes. HEAT-3D has been employed during the conceptual design phase of the Hypersonic foldable Aeroshell for THermal protection using ORigami (HATHOR), a mechanically deployable heat shield architecture that consists of rigid Thermal Protection System (TPS) panels fitted between retractable ribs and folded in an origami pattern when stowed. The sensitivity of heat transfer rates to the detailed aeroshell geometry is discussed, with particular attention to the seamless integration of the nose cap, ribs and TPS panels, and design recommendations are made.
Recent advances in highly efficient algorithms and high-performance computing allow us to build integrated design framework where the traditionally separate disciplinary models are coupled together, so as to improve the design optimization of hypersonic vehicles as integrated systems. Our particular interest in this paper is the potential approach to incorporating high-fidelity aerothermodynamic models in the hypersonic trajectory optimization problems. First, the necessity and motivation of considering high-fidelity aerodynamics are justified. Then, both the time-based and energy-based problem formulations for hypersonic trajectory optimization are introduced. Different from the conventional design approaches in the literature, two high-fidelity aerodynamic models are built and integrated in the trajectory optimization process. One is a panel method based on the modified Newtonian flow theory and Eckert's reference enthalpy method, and the other is a CFD model based on the Reynolds-Averaged Navier-Stokes equations. A pseudospectral optimal control package is used to solve the considered problem, and preliminary results demonstrate the feasibility of the developed approach.
Previous Mars entry, descent, and landing systems have utilized bank-angle maneuvers to steer during the hypersonic phase of entry. An alternate solution for hypersonic steering is a set of independently-articulated aerodynamic flaps. Deflecting these flaps results in trim at non-zero angles of attack and sideslip angles, enabling the vehicle to generate lift in arbitrary directions. This study investigates configuration options for flap-based steering systems for Mars entry applications. Configurations are parametrically studied by varying the number, size, geometry, and locations of flaps. Results from a computational aerodynamic trim solver indicate that the aerodynamic control authority of an entry vehicle can be increased by using a high number of flaps, and by using flaps with large areas and moment arms. Flap-steering vehicles also exhibit range capabilities with markedly different shapes than bank-angle steering vehicles with equivalent maximum lift-to-drag ratios. Results also show that as the number of flaps increases, a flap-steering vehicle behaves increasingly similar to a bank-angle steering vehicle with the same maximum lift-to-drag ratio.
Fast aerodynamic estimation is one of the key technologies for performance evaluation and optimization design in the preliminary design phase of the aircraft. It requires a balance between the calculation accuracy and calculation speed. This study aimed to establish a fast aerodynamic estimation combined model in the wide-speed domain and design fast aerodynamic estimation software based on the mechanistic theory and engineering model. First, a combined calculation model for pressure along the streamline was established based on the shock wave theory, expansion wave theory, modified Newton's theory, and so forth. Further, the combined calculation model for friction along the streamline was constructed based on the van Driest II theory and plate boundary layer theory. Second, the HL-20 lift body of the Langley Research Center was taken as the research object, and the results of the model were compared with high-precision computational fluid dynamics (CFD) data for verification. The results showed that the established model was in good agreement with CFD calculation results. Finally, combined wide-speed domain aerodynamic model and a rapid estimation platform (aerodynamic force preliminary evaluation, AFPE) independent of the calculation object was developed based on the streamline tracking technology and compared with the experimental data in the open literature. The results showed that AFPE had an excellent reproducibility of test data and an acceptable calculation speed in working ranges with relatively large Mach numbers, providing tool support for hypersonic aircraft high-speed aerodynamic estimation.
Hypersonic vehicles with Mach numbers greater than five exhibit complex shock wave interactions, accompanied by locally enhanced heat flux near the shock-affected surface. To investigate the unclear interstage thermal environment in a Two-Stage-To-Orbit system, a model composed of a wide-speed-range vehicle and a reusable rocket is employed in an aerothermodynamic study of the Two-Stage-To-Orbit system considering the interstage interactions performed at freestream Mach number = 6. The innovations of this paper are the complex model, which is highly similar to the real situation, and the study of the thermal environment between stages near different reflection positions, whereby the shock wave interference that causes different forms of heat flux is divided into two categories. The thermal design of the Two-Stage-To-Orbit system should especially consider the head of the wide-speed-range vehicle, interstage distance, and the swept angle of canted fins and wings, as well as their width. This paper adopts the second-order AUSMDV (a variant of the Advection Upstream Splitting Method) scheme for spatial discretization and Lower-Upper Symmetric Gauss–Seidel method for time discretization in formulating the Reynolds-averaged Navier–Stokes equations, which are then solved by the finite volume method and Menter’s shear stress transport k-ω turbulence model. Numerical analysis sheds light on the complex shock structure and the reasons for the considerable increase in wall temperature and aerothermal loads at critical parts of various reflected shock positions. The incident shock wave generated by the head apex of the rocket constantly reflects and gradually weakens between the two stages, which brings about shock wave/boundary-layer interaction and gives rise to elevated heat flux. The moderate incident shock and the weak reflected shock are primarily responsible for the lack of boundary-layer separation, although the expansion caused by the convex curvature of the head of the wide-speed-range vehicle also contributes. A non-reflected part of the incident shock wave interacts with the windward surface of the canted fins and wings, producing shock wave/boundary-layer interaction and a V-shaped heat flux distribution that extends to the leading edges of the wings. As the angle of attack decreases, the intensity of the incident shock and the reflected shock gradually weakens, leading to reduced heat flux on the windward surface of the wide-speed-range vehicle. The range over which the incident shock impinges the walls also narrows, corresponding to the smaller heat flux distribution. These aerothermal studies comprehensively determine the law of thermal changes, which is significant for the design of thermal protection in Two-Stage-To-Orbit systems.
An Earth smallsat mission architecture is developed to demonstrate the feasibility of aerocapture. The mission concept proposes to transfer from a geosynchronous transfer orbit rideshare trajectory to a low Earth orbit. Single-event drag modulation is used as a straightforward means of achieving the control required during the maneuver. Low- and high-fidelity guidance algorithm choices are considered. Numeric trajectory simulations and Monte Carlo uncertainty analyses are performed to show the robustness of the system to day-of-flight environments and uncertainties. Related investigations are described for Mars mission concepts to show the relevance of the proposed flight test to potential future applications.
Small satellites may provide a low-cost platform for targeted science investigations in the Mars system. With current technology, small satellites require ride shares with larger orbiters to capture into orbit, limiting the range of orbits available to small satellite mission designers. Successful development of an independent orbit insertion capability for small satellites, using aerocapture, would allow small satellite mission designers to choose the orbit most appropriate for a science investigation while enabling small satellite ride shares on any mission to Mars. A generic small satellite drag-modulation aerocapture system is assessed for use at Mars across a range of approach trajectories and destinations in the Mars system. Analyses include assessment of the sensitivity of the entry corridor size over different atmospheric conditions, a comparison of velocity-trigger and numerical predictor-corrector guidance schemes for drag modulation, and aerocapture flight performance assessment via Monte Carlo techniques. A special focus is placed on four baseline missions: a low-altitude Mars mapping orbit, Phobos and Deimos flyby/rendezvous, and areosynchronous orbit. Results indicate that aerocapture may decrease the orbit insertion system mass fraction by 30% or more with respect to fully propulsive options.
Rapid evaluation of aerodynamic forces is a key technology in conceptual and preliminary performance assessment and design optimization for supersonic/hypersonic vehicles, which is expected to achieve a balance between accuracy and time. In this paper, the preliminary evaluation platform of aerodynamic forces is built based on the physical theories and engineering methods. This platform computes the local velocity vector by using the streamline tracing technology, and extracts the pressure and friction coefficients by calling methods in method database. Furthermore, an example based on HL-20 is brought out. A good agreement is obtained by comparing the calculated results via CFD with the experimental. Also, a good angle of attack compatibility of streamline tracing technology is concluded. This platform has a computation time less than that via CFD and an flight range greater than that via wind tunnel. Finally, the simple flight dynamics simulation is taken out and the synergetic process is verified. © 2018, Editorial Board of Journal of Northwestern Polytechnical University. All right reserved.
The paper deals with the aerodynamic performance analysis of three reusable and unmanned flying laboratories designed to perform a return flight from low Earth orbit to provide experimental data in the framework of re-entry technologies. Several design approaches, ranging from low-order methods to computational fluid dynamics analyses, have been addressed in this work. In particular, vehicles aerodynamic performances for a wide range of free stream flow conditions, from subsonic to hypersonic regime, including reacting and non-reacting flow and different angles of attack have been provided and in some cases compared. Computational fluid dynamics results confirm that real gas effects seem to be fundamental for the assessment of the concept aerodynamics, especially concerning pitching moment evaluation.
This paper extends a previous modified axisymmetric analog method to predict heating rates on hypersonic vehicles in conjunction with inviscid computational fluid dynamic (CFD) codes which can provide more accurate inviscid solutions and are suited for complex configurations. The major problem is the heating anomalies encountered in the stagnation region, as the quality of the heating solution is very sensitive to the quality of the inviscid solution in the high-gradient stagnation region. To overcome this problem, a hybrid approach is developed to eliminate noise in the inviscid solution in the near-stagnation region by recalculating a noise-free inviscid solution in that region using an engineering method. As a result, there is no need to spend much effort to compute a high-quality inviscid solution in the near-stagnation region when solving the inviscid solution using an inviscid CFD code, thus significantly reducing run times. The proposed method is applied to several typical hypersonic vehicles and compared with existing approaches to validate its effectiveness. The results show that the proposed method can predict surface heating rates on complex configurations with reasonable accuracy but requires much shorter computational times.
This chapter deals with the aerodynamic performance analysis of three reusable and unmanned aerial vehicles conceived as flying laboratories to perform experimental flights in low Earth orbit (LEO). Each vehicle concept is an orbital re-entry vehicle (ORV), with re-entry energy of the order of 25 MJ/kg. The chapter shows the vehicle configuration of ORV-WSB, ORV-WBB, and ORV-SB. From the perspective of aerospace vehicle designers, the emphasis is on aerodynamic performance, stability, control, vehicle surface pressure, and shear and heating loads. The considered vehicle concepts feature a compact wing-body configuration equipped with a rounded-edge delta-like fuselage cross section, a delta wing, and a V-tail. The chapter also talks about rarefied and transitional regimes for ORV concepts. The ORV re-entry trajectory is reported in the Mach-Reynolds numbers map together with iso-Knudsen curves, which bound the different flow regimes, according to the Bird regime classification.
A novel method is described of directly calculating the force on N in
the gravitational N-body problem that grows only as N log N. The
technique uses a tree-structured hierarchical subdivision of space into
cubic cells, each of which is recursively divided into eight subcells
whenever more than one particle is found to occupy the same cell. This
tree is constructed anew at every time step, avoiding ambiguity and
tangling. Advantages over potential-solving codes include accurate local
interactions, freedom from geometrical assumptions and restrictions, and
applicability to a wide class of systems, including planetary, stellar,
galactic, and cosmological ones. Advantages over previous hierarchical
tree-codes include simplicity and the possibility of rigorous analysis
of error.
Laminar and turbulent heating-rate calculations from an 'engineering' code and laminar calculations from a 'benchmark' Navier-Stokes code are compared with experimental wind-tunnel data obtained on several candidate configurations for the X-33 Phase 2 flight vehicle. The experimental data were obtained at a Mach number of 6 and a freestream Reynolds number ranging from 1 to 8 x 10(exp 6)/ft. Comparisons are presented along the windward symmetry plane and in a circumferential direction around the body at several axial stations at angles of attack from 20 to 40 deg. The experimental results include both laminar and turbulent How. For the highest angle of attack some of the measured heating data exhibited a "non-laminar" behavior which caused the heating to increase above the laminar level long before 'classical' transition to turbulent flow was observed. This trend was not observed at the lower angles of attack. When the flow was laminar, both codes predicted the heating along the windward symmetry plane reasonably well but under-predicted the heating in the chine region. When the flow was turbulent the LATCH code accurately predicted the measured heating rates. Both codes were used to calculate heating rates over the X-33 vehicle at the peak heating point on the design trajectory and they were found to be in very good agreement over most of the vehicle windward surface.