Geometry definition and grid generation for a complete fighter aircraft

Recent advances in computing power and numerical solution procedures have enabled computational fluid dynamicists to attempt increasingly difficult problems. In particular, efforts are focusing on computations of complex three-dimensional flow fields about realistic aerodynamic bodies. To perform such computations, a very accurate and detailed description of the surface geometry must be provided, and a three-dimensional grid must be generated in the space around the body. The geometry must be supplied in a format compatible with the grid generation requirements, and must be verified to be free of inconsistencies. This paper presents a procedure for performing the geometry definition of a fighter aircraft that makes use of a commercial computer-aided design/computer-aided manufacturing system. Furthermore, visual representations of the geometry are generated using a computer graphics system for verification of the body definition. Finally, the three-dimensional grids for fighter-like aircraft are generated by means of an efficient new parabolic grid generation method. This method exhibits good control of grid quality

Component-Based Aerodynamic Shape Optimization using Overset Meshes

Advances in computational power allow the increase in the fidelity level of analysis tools used in conceptual aircraft design and optimization. These tools not only give more accurate assessments of aircraft efficiency, but also provide insights to improve the performance of next-generation aircraft. Aerodynamic shape optimization involves the inclusion of aerodynamic analysis tools in optimization frameworks to maximize the aerodynamic efficiency of an aircraft configuration via modifications of its outer mold line. When using CFD-based aerodynamic shape optimization, generating high-quality structured meshes for complex aircraft configurations becomes challenging, especially near junctions. Furthermore, mesh deformation procedures frequently generate negative volume cells when applied to these structured meshes during optimization. Complex geometries can be accurately modeled using overset meshes, whereby multiple high-quality structured meshes corresponding to different aircraft components overlap to model the complete aircraft configuration. However, from the standpoint of geometry manipulation, most methods operate on the entire geometry rather than on separate components, which diminishes the advantages of overset meshes. Tracking intersections among multiple components is a key challenge in the implementation of component-based geometry manipulation methods. The mesh nodes should also be updated in accordance to the intersection curves. This thesis addresses this issue by introducing of a geometry module that operates on individual components and uses triangulated surfaces to automatically compute intersections during optimization. A modified hyperbolic mesh marching algorithm is used to regenerate the overset meshes near intersections. The reverse-mode automatic differentiation is used to compute partial derivatives across this geometry module, so that it fits into an optimization framework that uses a hybrid adjoint method (ADjoint) to efficiently compute gradients for a large number of design variables. Particularities of the automatic differentiation of the geometry module are detailed in this thesis. By using these automatically updated meshes and the corresponding derivatives, the aerodynamic shape of the DLR-F6 geometry is optimized while allowing changes in the wing-fuselage intersection. Sixteen design variables control the fuselage shape and 128 design variables determine the wing surface. Under transonic flight conditions, the optimization reduces drag by 16 counts (5%) compared with the baseline design. This approach is also used to minimize drag of the PADRI 2017 strut-braced wing benchmark for a fixed lift constraint at transonic flight conditions. The drag of the optimized configuration is 15% lower than the baseline due to reduction of shocks and separation in the wing-strut junction region. This result is an example where high-fidelity modeling is required to quantify the benefits of a new aircraft configuration and address potential issues during the conceptual design. The methodologies developed in this work give additional flexibility for geometry and mesh manipulation tools used in aerodynamic shape optimization frameworks. This extends the applicability of design optimization tools to provide insights to more complex cases involving multiple components, including unconventional aircraft configurations.PHDAerospace EngineeringUniversity of Michigan, Horace H. Rackham School of Graduate Studieshttps://deepblue.lib.umich.edu/bitstream/2027.42/146042/1/neysecco_1.pd

Modeling and sensitivity analysis of aircraft geometry for multidisciplinary optimization problems

Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2012.This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.Cataloged from student-submitted PDF version of thesis.Includes bibliographical references (p. 415-421).A new geometry management paradigm for aircraft design utilizes Computer Aided Design (CAD) systems as the source for consistent geometry models across design phases and analysis tools. Yet various challenges inhibit the widespread application of CAD models in aircraft conceptual design because current CAD platforms are not designed for automated shape optimization. In particular, CAD models built with conventional methods can perform poorly in automated design frameworks and their associated CAD systems do not provide shape sensitivities. This thesis aims to remedy these concerns by bridging the computational geometry tools in CAD with aerospace design needs. A methodology for constructing CAD models is presented using concepts of multifidelity/multidisciplinary geometry and design motion. A formalized definition of design intent emerges from this approach that enables CAD models with parameterization flexibility, shape malleability and regeneration robustness for automated design settings. Analytic shape sensitivities are also presented to apply CAD models in gradient-based shape optimization. The parameterization and sensitivities for sketches, extrude, revolve and sweep features are given for mechanical design; shape sensitivities for B-spline curves and surfaces are also presented for airfoil and wing design. Furthermore, analytic methods modeling the sensitivity of intersection edges and nodes in a boundary representation (BRep) are given. Comparisons between analytic and finite-difference gradients show excellent agreement, however an error associated with the finite-difference gradient is found to exist if linearizing the support points of B-spline curves/surfaces and regenerating with a geometry kernel. This important outcome highlights a limitation of the finite-difference method when used on CAD models containing these entities. Finally, various example design problems are shown which highlight the application of the methods presented in the thesis. These include mechanical part design, inverse/forward design of airfoils and wings, and a multidisciplinary design space study. Gradient-based optimization is used in each design problem to compare the impact of analytic and finite-difference geometry gradients on the final designs obtained. With each of these contributions, the application of CAD-by David Sergio Lazzara.Ph.D

Modelling and Animation using Partial Differential Equations. Geometric modelling and computer animation of virtual characters using elliptic partial differential equations.

This work addresses various applications pertaining to the design, modelling and animation of parametric surfaces using elliptic Partial Differential Equations (PDE) which are produced via the PDE method. Compared with traditional surface generation techniques, the PDE method is an effective technique that can represent complex three-dimensional (3D) geometries in terms of a relatively small set of parameters. A PDE-based surface can be produced from a set of pre-configured curves that are used as the boundary conditions to solve a number of PDE. An important advantage of using this method is that most of the information required to define a surface is contained at its boundary. Thus, complex surfaces can be computed using only a small set of design parameters. In order to exploit the advantages of this methodology various applications were developed that vary from the interactive design of aircraft configurations to the animation of facial expressions in a computer-human interaction system that utilizes an artificial intelligence (AI) bot for real time conversation. Additional applications of generating cyclic motions for PDE based human character integrated in a Computer-Aided Design (CAD) package as well as developing techniques to describe a given mesh geometry by a set of boundary conditions, required to evaluate the PDE method, are presented. Each methodology presents a novel approach for interacting with parametric surfaces obtained by the PDE method. This is due to the several advantages this surface generation technique has to offer. Additionally, each application developed in this thesis focuses on a specific target that delivers efficiently various operations in the design, modelling and animation of such surfaces.The project files will not be available online

Computer-Aided Geometry Modeling

Techniques in computer-aided geometry modeling and their application are addressed. Mathematical modeling, solid geometry models, management of geometric data, development of geometry standards, and interactive and graphic procedures are discussed. The applications include aeronautical and aerospace structures design, fluid flow modeling, and gas turbine design

Aerodynamic Shape Optimization of Common Research Model Wing–Body–Tail Configuration

Peer Reviewedhttps://deepblue.lib.umich.edu/bitstream/2027.42/140640/1/1.C033328.pd

Conceptual design methodology for blended wing body aircraft

The desire to create an environmentally friendly aircraft that is aerodynamically efficient and capable of conveying large number of passengers over long ranges at reduced direct operating cost led aircraft designers to develop the Blended Wing Body(BWB) aircraft concept. The BWB aircraft represents a paradigm shift in the design of aircraft. The design offers immense aerodynamics and environmental benefits and is suitable for the integration of advanced systems and concepts like laminar flow technology, jet flaps and distributed propulsion. However, despite these benefits, the BWB is yet to be developed for commercial air transport. This is due to several challenges resulting from the highly integrated nature of the configuration and the attendant disciplinary couplings. This study describes the development of a physics based, deterministic, multivariate design synthesis optimisation for the conceptual design and exploration of the design space of a BWB aircraft. The tool integrates a physics based Athena Vortex Lattice aerodynamic analysis tool with deterministic geometry sizing and mass breakdown models to permit a realistic conceptual design synthesis and enables the exploration of the design space of this novel class of aircraft. The developed tool was eventually applied to the conceptual design synthesis and sensitivity analysis of BWB aircraft to demonstrate its capability, flexibility and potential applications. The results obtained conforms to the pattern established from a Cranfield University study on the BlendedWing Body Aircraft and could thus be applied in conceptual design with a reasonable level of confidence in its accuracy

Elements of computational flight dynamics for complete aircraft

This work summarizes the effort to introduce high–fidelity methods in aircraft design process. Flight dynamic characteristics are crucial in aircraft design. Simulation tools evaluate the aircraft statical and dynamical stability and manoeuvrability usually are based on a previously computed tabular aerodynamic model.During the conceptual design phase, the aerodynamic database is usually computed with semi–empirical methods. These tools rely on existing configurations databases (statistical methods) or linear aerodynamic hypothesis (e.g. vortex lattice method), and so are not suitable for innovative designs. The exploitation of such methods may lead to evaluation errors in the design process, which can be found only in the following steps and so may be very expensive to rectify via additional work, wind tunnel and flight testing, enlarging the time–to–market and increasing the whole life cycle product cost.The adoption of high–fidelity, physical based aerodynamic models starting from the very first steps of the aircraft design would reduce the uncertainty of current design procedures and prevent costly aircraft retrofitting. Computational fluid dynamics may be utilized to achieve the required high–fidelity, but, because of the substantial computational cost, it is currently used only during ensuing design steps. In this thesis the steps towards an autonomous high–fidelity flight dynamics analysis are presented.A tool for generating the aerodynamic tables with the semi–empirical United States air force stability and control digital compendium with the common parametric aircraft configuration schema is developed. The function for the flow solver Edge is updated and both scripts are implemented and validated inside the computerised environment for aircraft synthesis and integrated optimisation methods.Reduced order models to overcome computational fluid dynamics limitations for automated generation of aerodynamic tables are then presented. Two methods are developed in order to obtain a more efficient approach for samples positioning inside the flight envelope domain. Emphasis is given on the ability to capture nonlinearities appearance in the flow field with only a few computations over the whole flight envelope. The methods rely on Kriging interpolation, and are validated for semi–analytical functions and for real test cases. This may permit to reduce the number of required computational fluid dynamics solutions to use the flight simulator of a factor of some tens, without compromising the main aircraft statical and dynamical behaviour results.A test case is then presented, showing the statical and dynamical aircraft stability comparison between different geometry configurations, by use of reduced order models and with a low computational budget. The limitations of a derivatives based aerodynamic model are then presented for a test case, highlighting the differences with a computational fluid dynamics and flight dynamics full–coupled model. A blocks software architecture is used to obtain a tool open and customizable.A computational fluid dynamics based optimization loop is then used to analyse the longitudinal trim conditions of a test case, presenting the derivatives aerodynamic model limitations. The geometry optimization feasibility, considering the aircraft stability as objective, is assessed. A model based on aerodynamic derivatives is assumed for the representation of the aerodynamic loads, because traditionally used by flight dynamics tools. Advances in this direction are discussed

Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2018-2020 Period Phase 2

Lockheed Martin Aeronautics Company (LM), working in conjunction with General Electric Global Research (GE GR) and Stanford University, executed a 19 month program responsive to the NASA sponsored "N+2 Supersonic Validation: Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2018-2020 Period" contract. The key technical objective of this effort was to validate integrated airframe and propulsion technologies and design methodologies necessary to realize a supersonic vehicle capable of meeting the N+2 environmental and performance goals. The N+2 program is aligned with NASA's Supersonic Project and is focused on providing system level solutions capable of overcoming the efficiency, environmental, and performance barriers to practical supersonic flight. The N+2 environmental and performance goals are outlined in the technical paper, AIAA-2014-2138 (Ref. 1) along with the validated N+2 Phase 2 results. Our Phase 2 efforts built upon our Phase 1 studies (Ref. 2) and successfully demonstrated the ability to design and test realistic configurations capable of shaped sonic booms over the width of the sonic boom carpet. Developing a shaped boom configuration capable of meeting the N+2 shaped boom targets is a key goal for the N+2 program. During the LM Phase 1 effort, LM successfully designed and tested a shaped boom trijet configuration (1021) capable of achieving 85 PLdB under track (forward and aft shock) and up to 28 deg off-track at Mach 1.6. In Phase 2 we developed a refined configuration (1044-2) that extended the under 85 PLdB sonic boom level over the entire carpet of 52 deg off-track at a cruise Mach number of 1.7. Further, the loudness level of the configuration throughout operational conditions calculates to an average of 79 PLdB. These calculations rely on propagation employing Burger's (sBOOM) rounding methodology, and there are indications that the configuration average loudness would actually be 75 PLdB. We also added significant fidelity to the design of the configuration in this phase by performing a low speed wind tunnel test at our LTWT facility in Palmdale, by more complete modelling of propulsion effects in our sonic boom analysis, and by refining our configuration packaging and performance assessments. Working with General Electric, LM performed an assessment of the impact of inlet and nozzle effects on the sonic boom signature of the LM N+2 configurations. Our results indicate that inlet/exhaust streamtube boundary conditions are adequate for conceptual design studies, but realistic propulsion modeling at similar stream-tube conditions does have a small but measurable impact on the sonic boom signature. Previous supersonic transport studies have identified aeroelastic effects as one of the major challenges associated with the long, slender vehicles particularly common with shaped boom aircraft (Ref. 3). Under the Phase 2 effort, we have developed a detailed structural analysis model to evaluate the impact of flexibility and structural considerations on the feasibility of future quiet supersonic transports. We looked in particular at dynamic structural modes and flutter as a failure that must be avoided. We found that for our N+2 design in particular, adequate flutter margin existed. Our flutter margin is large enough to cover uncertainties like large increases in engine weight and the margin is relatively easy to increase with additional stiffening mass. The lack of major aeroelastic problems probably derives somewhat from an early design bias. While shaped boom aircraft require long length, they are not required to be thin. We intentionally developed our structural depths to avoid major flexibility problems. So at the end of Phase 2, we have validated that aeroelastic problems are not necessarily endemic to shaped boom designs. Experimental validation of sonic boom design and analysis techniques was the primary objective of the N+2 Supersonic Validations contract; and in this Phase, LM participated in four high speed wind tunnel tests. The first so-called Parametric Test in the Ames 9x7 tunnel did an exhaustive look at variation effects of the parameters: humidity, total pressure, sample time, spatial averaging distance and number of measurement locations, and more. From the results we learned to obtain data faster and more accurately, and made test condition tolerances easy to meet (eliminating earlier 60 percent wasted time when condition tolerances could not be held). The next two tests used different tunnels. The Ames 11 ft tunnel was used to test lower Mach numbers of 1.2 and 1.4. There were several difficulties using this tunnel for the first time for sonic boom including having to shift the measurement Mach numbers to 1.15 and 1.3 to avoid flow problems. It is believed that the 11 ft could be used successfully to measure sonic boom but there are likely to be a number of test condition restrictions. The Glenn 8x6 ft tunnel was used next and the tunnel has a number of desirable features for sonic boom measurement. While the Ames 9x7 can only test Mach 1.55 to 2.55 and the 11 ft can only test Mach 1.3 and lower, the Glenn 8x6 can test continuously from Mach 0.3 to 2.0. Unfortunately test measurement accuracy was compromised by a reference pressure drift. Post-test analysis revealed that the drift occurred when Mach number drifted slightly. Test measurements indicated that if Mach number drift is eliminated, results from the 8x6 would be more accurate, especially at longer distances, than results from the 9x7. The fourth test in the 9x7, called LM4, used everything we learned to comprehensively and accurately measure our new 1044-02 configuration with a full-carpet shaped signature design. Productivity was 8 times greater than our Phase 1 LM3 test. Measurement accuracy and repeatability was excellent out to 42 in. However, measurements at greater distances require the rail in the aft position and become substantially less accurate. Further signature processing or measurement improvements are needed for beyond near-field signature validation

Structural-Electromagnetic Simulation Coupling and Conformal Antenna Design Tool

Airborne and spaceborne radar has long been an effective tool for remote sensing, surveillance, and reconnaissance. Most airborne systems utilize antenna arrays that are installed inside the moldline of the aircraft or in radomes that protect the array from in-flight loads. While externally-mounted arrays can offer the advantage of larger apertures, sensor-vehicle interactions often result in performance degradation of both systems. The presence of an externally-mounted array will increase the vehicle’s drag and potentially affect the dynamics and control of the vehicle. In addition, in-flight structural loads will deform the array, thus resulting in relative phase errors. While there exist a multitude of physics-based simulation tools to determine the effects of the array on the aircraft performance, existing tools are not sufficient for generating deformed arrays necessary for determining in-flight array performance. In response to this need, a computer tool for analyzing antennas undergoing structural loads is developed. The Antenna Deformation Tool (ADT) has two primary uses: generating deformed geometry from the output of a structural Finite Element Model (FEM) for use in an Electromagnetic (EM) simulation, and designing conformal antenna arrays. The two commercial software packages ADT is optimized for are MSC NASTRAN and ANSYS HFSS. Specifically, ADT is designed to generate a deformed 3D Computer Aided Design (CAD) model from a NASTRAN structural mesh. The resulting CAD model is compatible with HFSS electromagnetic simulation software for the assessment of the effects of loads on performance. The main purpose for the development of ADT is to facilitate studies of how structural deformations affect airborne antenna arrays performance and to provide the capability to perform studies easily and quickly using different antennas on the same structural model. ADT capabilities are demonstrated using several representative airborne antenna array structures. ADT is also demonstrated in the design of conformal antenna arrays. ADT can import CAD geometry and deform it according to a prescribed deformation field. The deformation field can either be determined from structural simulations or be provided by the user. This functionality allows the user to take an existing planar antenna design and conform it to a desired shape. Within the scope of airborne antenna arrays, this would allow an engineer to conform the antenna to the moldline of the aircraft or other support structure. Currently, ADT can interpret only quad and triangular 2D elements from NASTRAN. In addition, its ability to interpret a surface from a point cloud is limited to surface meshes in which there are exactly four explicit vertices, or surfaces which have a fairly even boundary with no major discontinuities and can be divided into four even segments. ADT is tested on NASTRAN structural models of small to medium complexity, and the geometry generated from simple models is used in HFSS simulations with success (with occasional post processing required). The antenna deformation submodule shows favorable performance with sheet and solid CAD geometry, though post-processing is required in the case of the latter. Results of some deformed antennas simulated with HFSS in the 200 MHz range are presented. The surface error of the geometry produced by ADT varies with the type of input mesh, with curved meshes and surfaces having higher errors. In terms of average element edge length, the maximum surface error is up to 1% for surfaces with no to small curvatures, and up to 3.6% for highly curved surfaces. This translates to about 0.17% of the mesh diagonal. ADT contains a set of classes and functions which provide ample capabilities for surface generation from meshes, and the process implemented is mostly automatic, requiring minimal user intervention. Due to ADT defining deformed geometry purely on separate meshes, adjacent surfaces are not associative and continuity between them is not guaranteed, which inherently can result in small intersections. These intersections can cause meshing problems with HFSS; however, these issues can be mitigated by adding a small offset. While demonstrated applications are still limited, ADT promises to substantially contribute to the design of aircraft-integrated antennas and multifunctional structures. With very limited capabilities for generating and assessing deformed antenna geometry currently existing, ADT represents a unique tool. ADT could be used not only in developing the next-generation of airborne remote sensing technologies, but to characterize in-flight performance of existing systems as well