The development process for new aircraft configurations needs to be more efficient in terms of performance, cost and time to market. The potential to influence these factors is highest in early design phases. Thus, high confidence must be established in the product earlier than today. To accomplish this, the concept of virtual product development needs to be established. This implies having a mathematical representation of the product and its associated properties and functions, often obtained through numerical simulations. Building confidence in the product early in the development process through simulations postpones expensive testing and verification to later development stages when the design is more mature. To use this in aerodynamic design will mean introducing more advanced physical modelling of the flow as well as significantly reducing the turn around time for flow solutions. This work describes the benefit of using parallel computers for flow simulations in the aircraft design process. Reduced turn around time for flow simulations is a prerequisite for non-linear flow modelling in early design stages and a condition for introducing high-end turbulence models and unsteady simulations in later stages of the aircraft design process. The outcome also demonstrates the importance of bridging the gap between the research community and industrial applications. The computer platforms are very important to reduce the turn around time for flow simulations. With the recent popularity of Linux–clusters it is now possible to design cost efficient systems for a specific application. Two flow solvers are investigated for parallel performance on various clusters. Hardware and software factors influencing the efficiency are analyzed and recommendations are made for cost efficiency and peak performance. / Report code: LiU-TEK-LIC-2006:27.

Aircraft design has recently experienced a trend away from performance centric design towards a more balanced approach with increased emphasis on engineering an economically successful system. This approach focuses on bringing forward a comprehensive economic and life-cycle cost analysis, which can be addressed by the introduction of a dynamic method allowing the analysis of the future attractiveness of such a concept in the presence of uncertainty. One way of addressing this is through the use of a competitive market model. However, existing market models do not focus on the dynamics of the market, which results in poor predictive capabilities. The method proposed here focuses on a top-down approach that integrates a competitive model based on work in the field of system dynamics into the aircraft design process. The primary contribution is the demonstration of the feasibility of such integration. This integration is achieved through the use of surrogate models, which enabled not only the practical integration of analysis techniques, but also reduced the computational requirements so that interactive exploration as envisioned is actually possible. An example demonstration of this integration is built on the competition in the 250 seat large commercial aircraft market. Two aircraft models were calibrated to existing performance and certification data and then integrated into the system dynamics market model, which was then calibrated with historical market data. This calibration showed a much improved predictive capability as compared to the conventional logit regression models. The resulting market model was then integrated into a prediction profiler environment with a time variant Monte-Carlo analysis resulting in a unique trade-off environment. This environment was shown to allow interactive trade-off between aircraft design decisions and economic considerations while allowing the exploration potential market success in the light of varying external market conditions and scenarios. Another use of the existing outputs of the Monte-Carlo analysis was then realized by visualizing the model variables on a multivariate scatter plot. This enables the designer to define strategic market and return on investment goals for a number of scenarios and then directly see which specific aircraft designs meet these goals.

Internal combustion (IC) engines that consume hydrocarbon fuels have dominated the propulsion systems of air-vehicles for the first century of aviation. In recent years, however, growing concern over rapid climate changes and national energy security has galvanized the aerospace community into delving into new alternatives that could challenge the dominance of the IC engine. Nevertheless, traditional aircraft sizing methods have significant shortcomings for the design of such unconventionally powered aircraft. First, the methods are specialized for aircraft powered by IC engines, and thus are not flexible enough to assess revolutionary propulsion concepts that produce propulsive thrust through a completely different energy conversion process. Another deficiency associated with the traditional methods is that a user of these methods must rely heavily on experts experience and advice for determining appropriate design margins. However, the introduction of revolutionary propulsion systems and energy sources is very likely to entail an unconventional aircraft configuration, which inexorably disqualifies the conjecture of such connoisseurs as a means of risk management. Motivated by such deficiencies, this dissertation aims at advancing two aspects of aircraft sizing: 1) to develop a generalized aircraft sizing formulation applicable to a wide range of unconventionally powered aircraft concepts and 2) to formulate a probabilistic optimization technique that is able to quantify appropriate design margins that are tailored towards the level of risk deemed acceptable to a decision maker. A more generalized aircraft sizing formulation, named the Architecture Independent Aircraft Sizing Method (AIASM), was developed for sizing revolutionary aircraft powered by alternative energy sources by modifying several assumptions of the traditional aircraft sizing method. Along with advances in deterministic aircraft sizing, a non-deterministic sizing technique, named the Probabilistic Aircraft Sizing Method (PASM), was developed. The method allows one to quantify adequate design margins to account for the various sources of uncertainty via the application of the chance-constrained programming (CCP) strategy to AIASM. In this way, PASM can also provide insights into a good compromise between cost and safety.

The primary objective of this thesis was to study, implement, and test low-cost electronic flight control systems (FCS) in remotely piloted subscale research aircraft with relaxed static longitudinal stability. Even though this implementation was carried out in small, simplified test-bed aircraft, it was designed with the aim of being installed later in more complex demonstrator aircraft such as the Generic Future Fighter concept demonstrator project. The recent boom of the unmanned aircraft market has led to the appearance of numerous electronic FCS designed for small-scale vehicles and even hobbyist-type model aircraft. Therefore, the purpose was not to develop a new FCS from scratch, but rather to take advantage of the available technology and to examine the performance of different commercial off-the-shelf (COTS) low-cost systems in statically unstable aircraft models. Two different systems were integrated, calibrated and tested: a simple, gyroscope-based, single-axis controller, and an advanced flight controller with a complete suite of sensors, including a specifically manufactured angle-of-attack transducer. A flight testing methodology and appropriate flight-test data analysis tools were also developed. The satisfactory results are discussed for different flight control laws, and the controller tuning procedure is described. On the other hand, the different test-bed aircraft were analysed from a theoretical point of view by using common aircraft-design methods and conventional preliminary-design tools. The theoretical models were integrated into a flight dynamics simulator, which was compared with flight-test data obtaining a reasonable qualitative correlation. Possible FCS modifications are discussed and some future implementations are proposed, such as the integration of the angle-of-attack in the control laws.

Weight and its estimation have a vital impact in the aircraft design process from the very early phases. When the conceptual design configuration of an aircraft has been created, it should ensure that the estimated weight is sufficiently accurate to meet the performance and cost requirements. The estimation of structural weight in early design stages is mostly performed using historical data; however, this approach does not provide reliable weight estimates for unconventional or unique designs or those that employ advanced materials. One solution for improving the accuracy of conceptual level weight estimation is to improve the fidelity of the methodology, geometry models, and loads through the use of Finite Element Methods (FEM). This thesis is intended to demonstrate an initial application of conceptual-level FEM based weight estimation to aircraft wing structure. / text

A set of techniques is developed that allows the incorporation of flight dynamics metrics as an additional discipline in a high-fidelity aerodynamic optimization. Specifically, techniques for including static stability constraints and handling qualities constraints in a high-fidelity aerodynamic optimization are demonstrated. These constraints are developed from stability derivative information calculated using high-fidelity computational fluid dynamics (CFD). Two techniques are explored for computing the stability derivatives from CFD. One technique uses an automatic differentiation adjoint technique (ADjoint) to efficiently and accurately compute a full set of static and dynamic stability derivatives from a single steady solution. The other technique uses a linear regression method to compute the stability derivatives from a quasi-unsteady time-spectral CFD solution, allowing for the computation of static, dynamic and transient stability derivatives. Based on the characteristics of the two methods, the time-spectral technique is selected for further development, incorporated into an optimization framework, and used to conduct stability-constrained aerodynamic optimization. This stability-constrained optimization framework is then used to conduct an optimization study of a flying wing configuration. This study shows that stability constraints have a significant impact on the optimal design of flying wings and that, while static stability constraints can often be satisfied by modifying the airfoil profiles of the wing, dynamic stability constraints can require a significant change in the planform of the aircraft in order for the constraints to be satisfied.

A set of techniques is developed that allows the incorporation of flight dynamics metrics as an additional discipline in a high-fidelity aerodynamic optimization. Specifically, techniques for including static stability constraints and handling qualities constraints in a high-fidelity aerodynamic optimization are demonstrated. These constraints are developed from stability derivative information calculated using high-fidelity computational fluid dynamics (CFD). Two techniques are explored for computing the stability derivatives from CFD. One technique uses an automatic differentiation adjoint technique (ADjoint) to efficiently and accurately compute a full set of static and dynamic stability derivatives from a single steady solution. The other technique uses a linear regression method to compute the stability derivatives from a quasi-unsteady time-spectral CFD solution, allowing for the computation of static, dynamic and transient stability derivatives. Based on the characteristics of the two methods, the time-spectral technique is selected for further development, incorporated into an optimization framework, and used to conduct stability-constrained aerodynamic optimization. This stability-constrained optimization framework is then used to conduct an optimization study of a flying wing configuration. This study shows that stability constraints have a significant impact on the optimal design of flying wings and that, while static stability constraints can often be satisfied by modifying the airfoil profiles of the wing, dynamic stability constraints can require a significant change in the planform of the aircraft in order for the constraints to be satisfied.

Shock control bumps (SCBs) are considered promising flow control devices for transonic commercial aircraft. By generating a λ-shock structure, 2D SCBs offer large drag savings, but perform poorly when that structure breaks down off-design. Milder-performing 3D devices produce weak vortices, that may offer some boundary layer control, and SCBs also affect buffet via direct impact on shock motions and separation. To date however, design studies have largely ignored complications from the swept, spanwise-varying flows, so this thesis tackles the question of whether SCB arrays can offer useful benefits to the performance of transonic commercial aircraft. Using a numerical infinite-wing model, a simple rotation adaptation is shown to redress deficient on-design drag performance of 3D SCBs under swept flows. With the correct rotation (dependent on height, planform and spacing) bumps follow performance-design trends similar to those in unswept flow. With this knowledge, an array design method is developed to tailor 2D and 3D devices to local flow conditions on an aircraft model, aiming to maximise on-design drag performance. Careful infinite-wing setup means the influence of rotation and array height on performance is replicated on the aircraft. Predicted array designs achieve 74-87% of their estimated local drag savings. However, with wave drag being a smaller percentage of the total, the influence of arrays on lift is more significant and makes the optimal designs shorter than predicted. Strategies for improving off-design drag performance are then evaluated. Stagger, an alternating chordwise translation applied to 3D arrays, broadens operating range and lowers drag penalties by better accommodating off-design shock movements, but offers a less favourable trade-off against on-design drag than simply reducing the array height. However, a 2D array can always outperform a 3D on drag objectives. Lastly, buffet performance is inferred using steady indicators based on trailing edge pressure and shock location. These disagree regarding the impact on buffet onset, unresolvably due to a lack of validation data, but agree that arrays could alleviate flow development post-onset. Optimal array designs depend on prioritised objectives: considering buffet severity and on-design drag, tall 2D (or 3D) arrays; for buffet and minimum off-design drag penalties (similar to the motivation behind vortex generator application), 3D arrays of varying height and stagger. A simple flight fuel consumption model utilising the computed drag data shows that many arrays are neutral or offer small savings (up to 0.3%) across a range of mission profiles. While likely too small to merit application for solely drag purposes, this implies buffet benefits without cost to efficiency. Unsteady tests and proper assessment of buffet onset are needed to confirm this.

The work presented in this thesis was performed with the aim of developing improved prediction methods for aerodynamic and aeroelastic analysis to be used in aircraft design. The first part of the thesis concerns the development of a viscous-inviscid interaction model for steady aerodynamic predictions. Since an inviscid, potential flow, model already is available, the main focus is on the development of a viscous model consisting of a three-dimensional integral boundary layer model. The performance of the viscous-inviscid interaction model is evaluated and it is found that the accuracy of the predictions as well as the computational cost appear to be acceptable for the intended application. The presented work also includes an experimental study aimed at analyzing steady and unsteady aerodynamic characteristics of a laminar flow wing model. An enhanced understanding of these characteristics is presumed to be useful for the development of improved aerodynamic prediction models. A combination of nearly linear as well as clearly nonlinear aerodynamic variations are observed in the steady as well as in the unsteady experimental results and it is discussed how these may relate to boundary layer properties as well as to aeroelastic stability characteristics. Aeroelastic considerations are receiving additional attention in the thesis, as a method for prediction of how flutter characteristics are affected by modeling uncertainties is part of the presented material. The analysis method provides an efficient alternative for obtaining increased information about, as well as enhanced understanding of, aeroelastic stability characteristics. / <p>QC 20170816</p>

This material is based on research sponsored by Air Force Research Laboratory under agreement number FA8650-09-2-3938. The U.S. Government is authorized to reproduce and distribute reprints for Governmental purposes notwithstanding any copyright notation thereon. The views and conclusions contained herein are those of the authors and should not be interpreted as necessarily representing the official policies or endorsements, either expressed or implied, of Air Force Research Laboratory or the U.S. Government. / This work seeks to develop multidisciplinary design optimization (MDO) methods to find the optimal design of a particular aircraft called an Efficient Supersonic Air Vehicle (ESAV). This is a long-range military bomber type of aircraft that is to be designed for high speed (supersonic) flight and survivability. The design metric used to differentiate designs is minimization of the take-off gross weight. The usefulness of MDO tools, rather than compartmentalized design practices, in the early stages of the design process is shown. These tools must be able to adequately analyze all pertinent physics, simultaneously and collectively, that are important to the aircraft of interest. Low-fidelity and higher-fidelity ESAV MDO frameworks have been constructed. The analysis codes in the higher-fidelity framework were validated by comparison with the legacy B-58 supersonic bomber aircraft. The low-fidelity framework used a computationally expensive process that utilized a large design of computer experiments study to explore its design space. This resulted in identifying an optimal ESAV with an arrow wing planform. Specific challenges to designing an ESAV not addressed with the low-fidelity framework were addressed with the higher-fidelity framework. Specifically, models to characterize the effects of the low-observable ESAV characteristics were required. For example, the embedded engines necessitated a higher-fidelity propulsion model and engine exhaust-washed structures discipline. Low-observability requirements necessitated adding a radar cross section discipline. A relatively less costly computational process utilizing successive NSGA-II optimization runs was used for the higher-fidelity MDO. This resulted in an optimal ESAV with a trapezoidal wing planform. The NSGA-II optimizer considered arrow wing planforms in early generations during the process, but these were later discarded in favor of the trapezoidal planform. Sensitivities around this optimal design were computed using the well-known ANOVA method to characterize the surrounding design space. The lower and higher fidelity frameworks could not be combined in a mixed-fidelity optimization process because the low-fidelity was not faithful enough to the higher-fidelity analysis results. The low-fidelity optimum was found to be infeasible according to the higher-fidelity framework and vice versa. Therefore, the low-fidelity framework was not capable of guiding the higher-fidelity framework to the eventual trapezoidal planform optimum. / Air Force Research Laboratory / Ph. D.