Seongim Choi

Multifidelity Design Optimization of Low-Boom Supersonic Jets

The practical use of high-fidelity multidisciplinary optimization techniques in low-boom supersonic business-jet designs has been limited because of the high computational cost associated with computational fluid dynamics-based evaluations of both the performance and the loudness of the ground boom of the aircraft. This is particularly true of designs that involve the sonic boom loudness as either a cost function or a constraint because gradient-free optimization techniques may become necessary, leading to even larger numbers of function evaluations. If, in addition, the objective of the design method is to account for the performance of the aircraft throughout its full-flight mission while including important multidisciplinary tradeoffs between the relevant disciplines the situation only complicates. To overcome these limitations, we propose a hierarchical multifidelity design approach where high-fidelity models are only used where and when they are needed to correct the shortcomings of the low-fidelity models. Our design approach consists of two basic components: a multidisciplinary aircraft synthesis tool (PASS) that uses highly tuned low-fidelity models of all of the relevant disciplines and computes the complete mission profile of the aircraft, and a hierarchical, multifidelity environment for the creation of response surfaces for aerodynamic performance and sonic boom loudness (BOOM-UA) that attempts to achieve the accuracy of an Euler-based design strategy. This procedure is used to create three design alternatives for a Mach 1.6, 6-8 passenger supersonic business-jet configuration with a range of 4500 n mile and with a takeoff field length that is shorter than 6000 ft. Optimized results are obtained with much lower computational cost than the direct, high-fidelity design alternative. The validation of these design results using the high-fidelity model shows very good agreement for the aircraft performance and highlights the need for improved response surface fitting techniques for the boom loudness approximations.

Two-Level Multifidelity Design Optimization Studies for Supersonic Jets

The conceptual/preliminary design of supersonic jet configurations requires multidisciplinary analyses tools, which are able to provide a level of flexibility that permits the exploration of large areas of the design space. High-fidelity analysis for each discipline is desired for credible results; however, the corresponding computational cost can be prohibitively expensive, often limiting the ability to make drastic modifications to the aircraft configuration in question. Our work has progressed in this area, and we have introduced a truly hybrid, multifidelity approach in multidisciplinary analyses and demonstrated, in previous work, its application to the design optimization of a low-boom supersonic business jet. In this paper, we extend our multifidelity approach to the design procedure and present a two-level design of a supersonic business-jet configuration, in which we combine a conceptual low-fidelity optimization tool with a hierarchy of flow solvers of increasing fidelity and advanced adjoint-based sequential quadratic programming optimization approaches. In this work, we focus on the aerodynamic performance aspects alone: no attempt is made to reduce the acoustic signature. The results show that this particular combination of modeling and design techniques is quite effective for our design problem and the ones in general and that high-fidelity aerodynamic shape optimization techniques for complex configurations (such as the adjoint method) can be effectively used within the context of a truly multidisciplinary design environment. Detailed configuration results of our optimizations are also presented.

Multi-fidelity Design Optimization of Low-boom Supersonic Business Jets

The practical use of high-fidelity multidisciplinary optimization techniques in low-boom supersonic business jet designs has been limited because of the high computational cost associated with CFD-based evaluations of both the performance and the loudness of the ground boom of the aircraft. This is particularly true of designs that involve the sonic boom loudness as either a cost function or a constraint because gradient-free optimization techniques may become necessary, leading to even larger numbers of function evaluations. If, in addition, the objective of the design method is to account for the performance of the aircraft throughout its mission (T/O and landing, climb, acceleration , etc.) while including important multidisciplinary trade-offs between the relevant disciplines (performance, boom, structures, stability and control, propulsion, etc.) the situation only worsens. In order to overcome these limitations, we propose a hierarchical multi-fidelity design approach where high-fidelity models are only used where and when they are needed to correct the shortcomings of the low-fidelity models. Our design approach consists of two basic components: a multidisciplinary aircraft synthesis tool (PASS) that uses highly-tuned low-fidelity models of all of the relevant disciplines and computes the complete mission profile of the aircraft, and a hierarchical, multi-fidelity environment for the creation of response surfaces for aerodynamic performance and sonic boom loudness (BOOM-UA) that attempts to achieve the accuracy of an Euler-based design strategy. This procedure is used to create three design alternatives for a Mach 1.6, 6-8 passenger supersonic business jet configuration with a range of 4,500 nmi and with a T/O field length that is shorter than 6,000 ft. Optimized results are obtained with much lower computational cost than the direct, high-fidelity design alternative. The validation of these design results using the high-fidelity model show very good agreement for the aircraft performance and highlights the need for improved response surface fitting techniques for the boom loudness approximations.

Supersonic Business Jet Design using a Knowledge-Based Genetic Algorithm with an Adaptive, Unstructured Grid Methodology

In the design of supersonic low-boom aircraft, it is important to balance the aerodynamic performance and sonic boom requirements in a way that represents the best compromise for the overall design. Since the ground sonic boom is typically not a smooth function of the design variables and may actually contain multiple local minima, it is important to select an optimization algorithm that is able to cope with this kind of design space. In this work, we study the use of Kriging approximation models for both boom and performance and use them in conjunction with genetic algorithm techniques to investigate the computational cost and characteristics of such an approach for the design optimization of a low-boom supersonic business jet (SBJ). Direct use of genetic algorithms with high-fidelity CFD analysis tools has been limited by the inherently large computational cost of genetic algorithms (GAs). The use of computationally inexpensive approximation models in lieu of high-fidelity CFD greatly improves the robustness and eciency of the design process for searches in relatively large design spaces. In order to improve the performance of this method, a new hybridization strategy that combines a GA with gradient information is proposed and its improved convergence rate is demonstrated. Regardless, the proposed procedures still require a large number of evaluations of the flow and boom patterns for dierent points in the design space. For this purpose we have built two automated Euler analysis tools that use a CAD-based geometry engine, and both multiblock-structured and unstructured, adaptive meshing techniques (they are named QSP107 and QSP-UA respectively). QSP-UA, has been developed to handle the geometric detail of the complete configuration including wing, fuselage, nacelles, diverters, empennage, etc., and to provide accurate performance and boom data. Results of sample test problems, and a 15-dimensional design case are presented and discussed.

Helicopter Rotor Design Using a Time-Spectral and Adjoint-Based Method

A time-spectral and adjoint-based optimization procedure that is particularly efficient for the analysis and shape design of helicopter rotors is developed. The time-spectral method is a fast and accurate algorithm to simulate periodic, unsteady flows by transforming them to a steady-state analysis using a Fourier spectral derivative temporal operator. An accompanying steady-state adjoint formulation for periodic unsteady problems is then possible and enables an unsteady flow design optimization procedure. The time-spectral CFD analysis is validated against conventional time-accurate CFD and flight test data for a UH-60A Black Hawk helicopter rotor in high speed forward flight. A multidisciplinary structural dynamics and comprehensive analysis coupling is employed for validation study to include blade structural dynamics and to enforce vehicle trim. Application of the adjointbased design method is carried out to optimize blade shape for a UH-60A. An uncoupled aerodynamics only calculation is performed for design application, holding the blade deformations and trim angles fixed to the previously calculated values. Minimization of power is pursued with non-linear constraints on thrust and drag force, resulting in a simplified form of the trim condition. The blade twist distribution, airfoil section shapes, and outboard planform shape comprise over 100 design variables. Starting from the initial configuration, the optimizer found a new design that shows good performance improvement, amounting to a 2% decrease in torque and 7% increase in thrust compared to the baseline UH-60A. The results demonstrate the potential of the time-spectral and adjoint-based design method for helicopter rotors.

Two-Level Multi-Fidelity Design Optimization Studies for Supersonic Jets

The conceptual/preliminary design of superonic jet configurations requires multi-disciplinary analyses (MDAs) tools which are able to provide a level of flexibility that permits the exploration of large areas of the design space. High-fidelity analysis for each discipline is desired for credible results, however, corresponding computational cost can be prohibitively expensive often limiting the the ability to make drastic modifications to the aircraft configuration in question. Our work has progressed in this area, and we have introduced a truly hybrid, multi-fidelity approach in MDAs and demonstrated, in previous work, its application to the design optimization of low-boom supersonic business jet. In this paper we extend our multi-fidelity approach to the design procedure and present two-level design of a supersonic business jet configuration where we combine a conceptual low-fidelity optimization tool with a hierarchy of flow solvers of increasing fidelity and advanced adjoint-based Sequential Quadratic Programming (SQP) optimization approaches. In this work, we focus on the aerodynamic performance aspects alone: no attempt is made to reduce the acoustic signature. The results show that this particular combination of modeling and design techniques is quite effective for our design problem and the ones in general, and that highfidelity aerodynamic shape optimization techniques for complex configurations (such as the adjoint method) can be effectively used within the context of a truly multi-disciplinary design environment. Detailed configuration results of our optimizations are also presented.

Validation Study of Aerodynamic Analysis Tools for Design Optimization of Helicopter Rotors

The key objectives of this paper are to assess the accuracy and the validity of our current aerodynamic analysis tools in predicting the unsteady flow field generated by helicopter rotors and to investigate their applicability to the future design problems. A ReynoldsAveraged Naiver-Stokes (RANS) solver with various turbulence models has been used with necessary modifications for the computation of all test cases. Dynamic stall and massive separation, which are physical phenomena characteristic of helicopter flows, are analyzed first using RANS methodology, and Detached Eddy Simulation (DES) is applied to the simulation of massive separation and compared to corresponding RANS solutions. The periodic nature of the helicopter flowfield strongly motivates the application of TimeSpectral (TS) approach for unsteady RANS computations. In this paper, the TS method is applied to simulate actual flight conditions of the UH-60A helicopter and is verified to be a fast and efficient algorithm which maintains the same level of accuracy as the timeaccurate approach. In addition, the TS method also provides great potential for adjoint based design optimization of helicopter rotors as the governing equations can be reduced to a periodic steady state.

CFD Prediction of Rotor Loads using Time-Spectral Method and Exact Fluid-Structure Interface

The primary objective of this paper is to study time-spectral method for simulating helicopter rotor flows in steady flight. The intent is to compare the accuracy of predicted vibratory loads (both airloads and structural loads) with time-accurate computations and quantify the aliasing error and convergence behavior in a precise manner. The CFD/Comprehensive Analysis coupling method in this paper is different from the stateof-the-art. We implement an exact fluid-structure interface for rotors and formulate a modified delta coupling procedure, that is generic for advanced geometry blades, underlying structural models, and unstructured surface grids. The Counter 8534 flight from the U. S. Army/NASA Airloads Program, the highest vibration flight of this helicopter, is used for validation. The fastest and minimal implementation of the method, with 11 time instances for a 4-bladed rotor, predicts the vibratory normal forces and pitching moments within 5–10% accuracy with respect to time-accurate simulations in about one-fifth the computational time. The largest errors occur in vibratory chord forces (10 to 20%). This level of error generates un-satisfactory levels of error in structural loads. However, the primary source of error is aliasing, which for this flight decreases asymptotically with an increasing number of time instances. We demonstrate an accuracy level of 1% and 0.1% in airloads with 17 and 25 time instances respectively. These correspond to one-third and one-half the computational time of a time-accurate solution. It is concluded that timespectral method in CFD can be used effectively for the prediction of rotor vibratory loads. However, without any anti-aliasing filter, reliable prediction of structural loads requires a number of time instances at least four times the blade number – still at one-third the time, approximately, compared to a time-accurate solution.

Numerical and Mesh Resolution Requirements for Accurate Sonic Boom Prediction

A careful study is conducted to assess the numerical mesh resolution requirements for the accurate computation of sonic boom ground signatures produced by complete aircraft configurations. The details of the ground signature can depend heavily on the accurate prediction of the pressure distribution in the near field of the aircraft. It is, therefore, important to accurately describe the geometric details of complete configurations (including the wing, fuselage, nacelles, diverters, etc.) and to precisely capture the propagation of shock and expansion waves at large distances from the aircraft. Unstructured, adaptive mesh technologies are ideally suited for this purpose as they use mesh points only in the appropriate locations within the flowfield. In this work, we consider a supersonic business jet configuration that was tested at the NASA Langley Research Center. Near-field data were measured at several locations underneath the flight track. The propagation of these near-field signatures from different altitudes can be shown to result in near N-wave ground booms. To examine the effect of both nacelles and empennage, results for three test cases are presented. These test cases represent the complete configuration, the configuration without the nacelles, and the configuration without the nacelles and empennage. Inviscid solution-adaptive unstructured meshes with up to 7.2 million nodes and 42.1 million tetrahedra are used to calculate the pressure distributions at several locations below each configuration where comparisons with experimental data are performed. All near-field pressure distributions are propagated to the ground (from an altitude of 50,000 ft) to predict the ground boom and the perceived noise level of the ground signature. Both the near-field overpressures and ground boom signatures are compared between experimental data and computational fluid dynamics simulation, and the results show good agreement in all cases. The minimum number of mesh nodes and elements and the levels of refinement needed for the accurate computations of near-field pressure distribution and ground boom signature are discussed for each of the three cases.

A Preliminary Study of Open Rotor Design Using a Harmonic Balance Method

An open rotor is one of the next generation aero-engines as it has 30% higher efficiency compared to the conventional turbofan engines. However, a high level of noise loudness has been a major drawback of the open rotor for its commercial use in aviation market. Although there have been a number of efforts to reduce its noise level, an accurate prediction of unsteady and complex flow field of the open rotor makes it difficult for the design methodologies to be applied. This paper introduces one of the state-of-the-art design methodologies to handle the unsteady problem of low-noise open rotor design. A harmonic balance method which is an order of magnitude more efficient than the conventional timeaccurate CFD method is used to analyze open rotor flows. To demonstrate the accuracy of the harmonic balance method, a wind-tunnel experiment of the scaled model of the open rotor is carried out and aerodynamic performances are compared with the harmonic balance predictions. With the steady formulation of the flow governing equations through the harmonic balance method, a design method using a surrogate model is employed to find an optimum configuration that minimizes the noise level and total power at a constant thrust level. A noise prediction is computed using the Farassat formula, derived from the FfowcsWillimas Hawkings equation. Design variables of the blade radii, rotor spacing, and the pitch angle variation of the aft rotor are chosen. A parameter study to investigate the sensitivities of the design parameters to thrust and torque/power levels as well as to the noise loudness is carried out. A genetic algorithm to handle multi-objectives is used in combination with the surrogate model of Kriging response surface. An optimum configuration is obtained from the pareto front of the optimization results and shows the reduction of noise level by 7dB and power level by 4% from the baseline values.