Aerobatic Trainer Aircraft Conceptual Design

Aircraft conceptual design is the very beginning of aircraft design process in which one or several aircraft configurations are created to meet the top-level requirements of aircraft. The aim of this paper is to build a more traceable, structured and systematic process of aircraft conceptual design by use of Quality Function Deployment (QFD) method. The conceptual design of a narrow body commercial aircraft is used as an example to illustrate application of the QFD. The top-level requirements of the commercial aircraft are defined. The conceptual design activities such as initial sizing of aircraft, wing configuration design, fuselage configuration design, and empennage configuration design are accomplished using House of Quality (HOQ). During initial sizing of aircraft, the top-level requirements of aircraft are converted into the requirements for aerodynamics, propulsion and weight with the aid of HOQ and constraint analysis. During configuration designs of the wing, fuselage and empennage, their requirements are identified based on the top-level requirements and output of the initial sizing, and consist of requirements of aerodynamics, structure, control, safety, operation, cost and etc. Several HOQs representing the relationships between the requirements and the configuration design parameters are created. After that, the configuration design parameters are selected with the aid of the HOQ and relevant design knowledges. By use of the QFD, aircraft conceptual design activities are represented by a number of the HOQ and the design process of configuration parameters is traceable through the HOQ. It is concluded that the QFD-based aircraft conceptual design is more traceable, structured and systematic than the traditional method.

A holistic review of the current state of research on aircraft design concepts and consideration for advanced air mobility applications

Recent years have seen a renewed interest in rotary-wing aircraft, due to the emergence of urban air mobility. In the conceptual design of such aircraft, fast estimation of rotor performance is important. A challenge is that rotor performance models require a detailed description of the blade geometry, which is not always available at this stage. We have developed a new, robust method for rapidly computing optimal rotor designs, which is derived from blade element momentum theory. In this paper, we derive the rotor design method and prove its convergence for all reasonable model inputs. We verify the design method by comparing the optimal rotor designs it computes to those found by a nonlinear programming algorithm. Thousands of numerical experiments yield relative errors of less than 13 and 5% in terms of the chord and twist profiles, respectively, and less than 0.3% in terms of efficiency, thrust, torque, and power coefficients. In addition, the design method computes optimal designs 170 times faster compared to the nonlinear programming algorithm. Based on these findings, we expect this new rotor design method to be useful in the conceptual design of rotary-wing aircraft.

The fully tailless configuration has lower observability, less structural weight and less drag, and it is considered one of the preferred designs for the next generation of efficient supersonic combat aircraft. In the conceptual design of such novel aircraft, a parametric geometry model is essential for multidisciplinary design analysis and optimization (MDAO). This paper presents a parametric three-dimensional (3D) geometry modeling methodology and tool for MDAO in the conceptual design of a notional supersonic tailless combat aircraft (STCA). The geometries of the STCA components (wing, fuselage and propulsion) are defined specifically by a set of parameters. In particular, the inlet and nozzle geometries are defined with the required details. Based on the geometric relationships among the STCA components, an approach involving master-dependent parameters is proposed. The geometry model generated by the approach has features such as the fuselage being blended smoothly with the wing and the propulsion being well integrated with the fuselage. Moreover, the geometry model can be generated by simply specifying the values of the master parameters, and the number of parameters required to generate the geometry model is reduced substantially. Based on the methodology, a parametric geometry modeling tool for the STCA conceptual design is developed using a Visual Basic (VB) script in the CATIA V5 platform. The applicability of the tool is validated with several case studies.

Blended wing body (BWB) aircraft represent a paradigm shift in jet transport aircraft design that promise benefits over conventional aircraft. A method is presented to enable the conceptual design of BWB aircraft, enabling comparison studies with tube-and-wing aircraft (TAW) based on the same top-level requirements and analysis fidelity. The aim of this work is to make comparative studies between the blended-wing-body aircraft and its conventional tube-and-wing counterpart based upon the same design requirements at conceptual level. By developing a novel geometric parametrisation of the blended wing body, the design possibilities have been increased while maintaining straightforward shaping manipulation and robustness. The mass estimation methods that have been implemented are verified and validated to be within approximately 5% of reference blended wing bodies. Drag estimations perform less accurately with drag being overpredicted by approximately 20%. The cause of this over prediction was largely due to empirical corrections for miscellaneous and unaccounted drag sources as well as induced drag predicted by a vortex-lattice method. Test-case BWB and TAW aircraft were formed in the 150, 250 and 400 passenger classes. The comparisons of the resulting aircraft show that the blended wing body have reduced mass, improved aerodynamic efficiency and higher fuel economy. Trends also show that the improvements over tube-and-wing aircraft increase with aircraft size.

Regulations in the United States stipulate that all airline pilots begin receiving upset prevention and recovery training in 2019. This year, the Federal Aviation Administration (FAA) published guidance to meet those regulations. Two important approvals are necessary before an operator delivers this training. First, the principal FAA inspector for that airline will have to approve the training program. Second, pilots and engineers from the National Simulator Program will have to qualify the simulator used for the training. To add consistency to these approvals, training that covers the main principles of the regulations and guidance is being provided academically and in a simulator to the personnel in both approval paths. To date, the training has been beta-tested with several FAA pilot and engineering groups. This paper covers the training, what has been learned in its early delivery, as well as the key questions arising since the regulations and guidance have been issued. The predominant challenge is to undo the previous technique of trying to use engine thrust as the primary means to recover from a stall.

Nomenclature Aircraft design is a complex multidisciplinary process to determine aircraft configuration variables that satisfy a set of mission requirements. It is very hard for aircraft designers to foresee the consequences of changing certain variables. Furthermore, conventional optimization processes are limited by the type and number of parameters used, resulting in sub-optimal designs. The objective of this research is to test the functionality and implementation of a multidisciplinary aircraft conceptual design optimization method using an adaptive genetic algorithm (GA), as a feasible alternative to the existing sizing and optimization methods. To illustrate the approach the algorithm is used to optimize a medium range commercial aircraft, with takeoff weight as an optimization goal, subjected to constraints in performance and geometric parameters. Adaptive and traditional formulations for the handling of constraints by the GA are tested and compared. Results show the ability of the adaptive GA to unbiased search through the design space of aircraft conceptual designs, leading to more viable aircraft configurations than the traditional GA approach at reduced timeframes, with a lower cost than current aircraft design optimization procedures.

Motion cueing algorithms (MCAs) enable realistic motion cues on limited-workspace platforms for flight training. In upset prevention and recovery training, conventional MCAs often fail to deliver critical sensory cues, which reduces pilot training performance. This study proposes a model predictive control-based MCA optimized using a deep Q-network. Motion cueing weights for specific scenarios are learned under platform constraints. As a result, the optimized MCAs improved overall perceptual tracking by at least 11.6% and increased the subjective ratings by at least 7.8%. These findings demonstrate the effectiveness of the proposed method in enhancing motion realism and training efficiency under high-demand conditions.

Upset Prevention and Recovery Training in Flight Simulators

This paper aims to contribute to developing a potential combined cycle air-breathing engine integrated into an aircraft design, capable of performing flight profiles on a commercial scale. This study specifically focuses on the single expansion ramp nozzle (SERN) and aircraft-engine integration with an emphasis on the combined cycle engine integration into the conceptual aircraft design. A parametric study using computational fluid dynamics (CFD) have been employed to analyze the sensitivity of the SERN’s performance parameters with changing geometry and operating conditions. The SERN adapted to the different operating conditions and was able to retain its performance throughout the altitude simulated. The expansion ramp shape, angle, exit area, and cowl shape influenced the thrust substantially. The internal nozzle expansion and expansion ramp had a significant effect on the lift and moment performance. An optimized SERN was assembled into a scramjet and was subject to various nozzle inflow conditions, to which combustion flow from twin strut injectors produced the best thrust performance. Side fence studies observed longer and diverging side fences to produce extra thrust compared to small and straight fences.

*† All designers are aware of the uncertainty inherent in conceptual aircraft design. Most often it is dealt with by the use of a safety factor or a design margin that is added to constraints or limits to account for variability between a predicted result and the actual result. Both safety factors and design margins can be very effective; however, they may result in an “over designed” aircraft and sometimes an “under designed” one. This may require late design changes, which are typically costly to perform. Engineers understand that a large source of uncertainty in initial aircraft sizing comes from the empty weight equations. They generally use an empirically-based equation or a set of such equations that are a function of major aircraft design parameters, like wing area, sweep, and aspect ratio and have been defined during conceptual design. It is possible to characterize the uncertainty in these equations and to use Monte Carlo simulations to combine all of the component weight uncertainties into a distribution of empty weight. Because empty weight affects the overall size and performance of the aircraft, the uncertainty in the weight equations can be propagated to the vehicle level. The end result is a “probabilistic carpet plot” that is much like a traditional carpet plot, only each design point in the sizing matrix is now represented by distributions of weight and performance values. A “minimally invasive” approach to account for uncertainties in aircraft sizing is provided here. Seemingly insignificant scatter about component weight trend lines leads to large uncertainties in the final design point.

The conceptual aircraft design and its integration with a combined cycle engine for hypersonic cruise at Mach 8 is documented in this paper. The paper describes the process taken to develop a hypersonic aircraft from a conceptual approach. The discussion also includes the design and CFD analysis of the integrated combined cycle engine. A final conceptual hypersonic transport aircraft with an integrated combined cycle engine was achieved through this study. According to the analysis carried out, the aircraft is able to take-off and land at the airports it is intended to be used and will be able to generate enough thrust to sustain hypersonic cruise at an altitude of 30 km.

Maneuver load alleviation, gust load alleviation, and natural laminar flow are integrated into aircraft conceptual design. The concurrent design of the aircraft and its active load alleviation systems can yield significant gains. The simultaneous application of maneuver load alleviation and gust load alleviation to a short-haul aircraft leads to an 11% reduction in fuel burn and 7% reduction in direct operating cost. These savings are diminished if maneuver load alleviation and gust load alleviation are applied independently. The synergy between active load control and natural laminar flow is also explored. It is possible to invest some of the weight savings from load alleviation to achieve extensive laminar flow. The combination of load alleviation and natural laminar flow increases the fuel and cost savings to upwards of 18 and 11%, respectively. It is concluded that load alleviation can shift the transonic transport optimum toward low-sweep, natural laminar flow wings. Sensitivity studies finally demonstrate that contemporary commercial aircraft should be able to meet the control requirements for effective gust load alleviation.

The research presented focuses on investigating the use of Cross-Flow Fan (CFF) as a high-lift device for a Short Take-off and Landing (STOL) aircraft. The wing-embedded fan performance analysis is mostly addressed from an aerodynamic perspective and focuses on using such Active Flow Control (AFC) technology in the conceptual aircraft design process. In particular, the design trade study of an aircraft featuring the fan as a high-lift device applied to a conceptual design of a medium-range multi-purpose aircraft is performed. A sensitivity analysis is employed to investigate the impact of the technology on the aircraft weight, aerodynamics, stability and control, and fight performance. The aircraft design modifications are introduced to maximize the aircraft mission performance given the fan specifications and constraints. Results indicate a reduction of the take-of field length by 18% with the payload penalty of 14%. The aircraft ferry range is also decreased by 7% compared to the baseline aircraft design. The scaling analysis of the aircraft concept is performed to determine the potential market for such technology. The results show that a light General Aviation (GA) airplane or a medium-large size Unmanned Aerial Vehicle (UAV) could benefit more from the wing-embedded CFF compared to more heavy airplanes.

Exploring the performance potential of conventional tube-and-wing aircraft is imperative to bridge the gap for sustainable aviation until the adoption of novel configurations. This paper investigates the strategy of reducing design range and cruise speed for the conceptual design of next-generation single-aisle aircraft. The impacts on fuel consumption are evaluated at both vehicle and fleet levels. Performance analysis shows that simultaneously optimizing aircraft design parameters and cruise Mach number can reduce mission fuel burn by 3% compared to the fixed-Mach optimization. The reduced cruise speed also maximizes the benefit of natural laminar flow technology, yielding an additional 5% fuel saving. Shorter design ranges result in smaller airframes and engines, reducing fuel burn by about 1.4% per 500 nmi decrease relative to the long-range vehicle operating at its off-design missions. Tradeoffs between aircraft range capability and fuel savings are captured by the fleet analysis. The results show that 6% of the flights operated by large single-aisle aircraft become infeasible if the aircraft design range reduces to 1500 nmi. The projected 2050 fleet-level fuel savings vary non-monotonically with reduced design ranges, with the highest reduction of 6.4% achieved by a 2500 nmi design among six optimized candidates.