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2023 | Book

The Proceedings of the 2021 Asia-Pacific International Symposium on Aerospace Technology (APISAT 2021), Volume 1

Editors: Sangchul Lee, Cheolheui Han, Jeong-Yeol Choi, Seungkeun Kim, Jeong Ho Kim

Publisher: Springer Nature Singapore

Book Series: Lecture Notes in Electrical Engineering

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About this book

This proceeding comprises peer-reviewed papers of the 2021 Asia-Pacific International Symposium on Aerospace Technology (APISAT 2021), held from 15-17 November 2021 in Jeju, South Korea. This book deals with various themes on computational fluid dynamics, wind tunnel testing, flow visualization, UAV design, flight simulation, satellite attitude control, aeroelasticity and control, combustion analysis, fuel injection, cooling systems, spacecraft propulsion and so forth. So, this book can be very helpful not only for the researchers of universities and academic institutes, but also for the industry engineers who are interested in the current and future advanced topics in aerospace technology.

Table of Contents

Frontmatter

Aerodynamics and Design

Frontmatter
Asia–Pacific Region’s Aircraft Design and Development Strategy for Single-Aisle Aircraft in the Oligopoly Market of Boeing and Airbus

There have been severe supply and demand imbalances in the global aircraft operation situation, but various factors make it difficult to address this market imbalance. Considering the number of commercial jet flights in December 2019, Asia–Pacific (AP) countries operate 34%, but supply is 0%. North America operates 25%, while Boeing supplies 55%. Europe operates 21%, while Airbus supplies 44%. Why this imbalance happened, how to address it, and how to develop aircraft will be suggested.In the 1960s, European countries had tried to correct the severe mismatch between aircraft supply and demand. France and Germany agreed to design and develop aircraft for the development of the European aerospace industry by international joint development and established Airbus in 1970. The A320 was designed and developed, and was finally launched in 1988. Airbus outperformed Boeing by taking a 52% market share with 305 units in 2003, which contributed significantly to the European economy. Europe established EADS in 2000 and integrated the European aeronautical defence industry, which contributed greatly to European peace. In 2020, Airbus delivered 566 units, occupying a 78% market share (Boeing 162), remaining the world’s No. 1 aircraft supplier and maintaining Europe’s economy and peace. Regional Comprehensive Economic Partnership (RCEP) countries should benchmark the Airbus A320 success stories to correct AP aircraft demand–supply distortions and lay the groundwork for Asia–Pacific aerospace industry development. To achieve long-term goals, it is effective to divide international joint development into three stages. In the first stage, individual companies (company-company) promote international co-development, and it is similar to the current environment of AP countries. The second stage is the case of Japan-promoted international joint development (JADC-Boeing) (companies consortium-company), and the third is international joint development between countries, as in the case of Airbus. This study is based on a quantitative approach to provide clear findings. AP aircraft manufacturers should secure the basis for the RSP international joint development project.

Myung-kwan Ahn, Jong-soo Lee, Heung-jae Kim, Hyeong-yu Jang
Study on Wind Resistance Characteristics of Multi-rotor UAV

Based on the wind field characteristics of high-rise buildings, this paper studies the wind resistance stability of multi-rotor UAV flight platform, conducts research on the design technology of high-wind resistance flight platform, improves the safety and reliability of fire fighting and extinguishing UAV, and provides technical support for the design and optimization of fire fighting and extinguishing UAV. Based on various aerodynamic data under no-load and on-load conditions, the wind resistance of the UAV is analyzed, and a set of wind resistance evaluation methods for the multi-rotor fire fighting and extinguishing UAV is proposed. The wind static model of the multi-rotor fire fighting UAV is established to evaluate the wind resistance of the UAV. The results show that: the designed fire-fighting UAV can theoretically resist the wind of level 7 (16 m/s), and the UAV can achieve the balance of force and torque by adjusting its attitude under the horizontal action of level 7 wind.

Zhan-ke Li, Shi-gao Su, Jin-shuo Cao, Si-jie Luo
Inverse Design Method of Transonic Airfoil Based on Deep Neural Network

This research introduces a deep neural network-based transonic airfoil inverse design approach. A deep neural network is established using the predicted pressure coefficient distribution, angle of attack, and Mach number around the airfoil as inputs and parameterized airfoil parameters as outputs. The network solves the transonic flow field of an airfoil using the Euler equation. The neural network is trained to grasp the impacts of geometric changes on the pressure coefficient and shock wave at various places on the airfoil surface. Using the finite volume approach and multigrid acceleration method, the Euler equation is employed to batch compute the airfoil database and build the training set. As the training set's output vector, the CST parameterization technique is utilized to parameterize each synchronously generated airfoil into 10 parameters. The results demonstrate that the deep learning approach developed in this research is capable of achieving the reverse design of a transonic airfoil, and that the trained neural network can generate the airfoil with the requisite aerodynamic properties rapidly and directly. The generalization error of the anticipated airfoil geometry information is less than 1.5 percent when compared to the accurate results. The results reveal a high level of precision. The impact of the deep neural network's shape and the activation function of neurons on the subjects under investigation is also examined. The results reveal that the activation function used is critical to the outcome, and the design of the neural network also has an impact on the outcome.

Zhiliang Bai, Wei Zhang, Ruyue Wei
High-Rise Building Wind Field Simulation

High-rise building fire extinguishing UAV is a new equipment used in high-rise building fire extinguishing and rescue. However, the turbulent environmental airflow in the high-rise building area and the complex wind field pose a threat to the stability of fire fighting and fire fighting drones, which puts forward higher requirements for the stability and wind resistance of multi-rotor UAV flight platforms. Therefore, aiming at multi-rotor drones that perform fire fighting tasks in high-rise buildings, the characteristics of wind fields between buildings in high-rise buildings are studied to provide technical references for the design and optimization of fire-fighting drones, which not only has important theoretical significance, but also has high practical value. With multi-rotor firefighting drones as the background, based on the computational fluid dynamics (CFD) theory and Fluent software, the wind field characteristics around high-rise buildings are studied, carrying out numerical simulation analysis of the wind field around single buildings and multiple buildings to provide a reference for the actual flight of fire extinguishing drones between buildings. The analysis results show that: Compared to the single building, the wind field around the building has changed a lot when the two buildings are arranged, and the aerodynamic interference between the two buildings is significantly weakened as the distance between the buildings increases; Layout and the surrounding buildings play a shielding role. The wind and shadow areas are superimposed on each other, and the air flow is complicated. Drones should avoid flying in these complex wind fields as for as possible.

Zhan-ke Li, Si-jie Luo, Shi-gao Su, Jin-shuo Cao, Xiao-min Zhang
Study on Aerodynamic Characteristics of “Propeller/Wing” System with Overset Mesh Method

In this paper, the propeller unsteady aerodynamic and “propeller/wing” aerodynamic interaction are investigated using the Computational Fluid Dynamics (CFD) method. Based on the true three-dimensional (3D) propeller simulation, a steady RANS solution for ANSYS FLUENT was presented using an actuator-disk approach, with a user-defined function that was proposed for the commercial solver. The full 3D unsteady RANS numerical simulation work was performed using four models: (1) a clean wing with span length of 1 m; (2) single propeller model composed of four blades; (3) the “propeller/wing” model containing propeller and 1-m wing; (4) a clean wing with span length of 3 m. On the other hand, the actuator-disk model contains three models: one disk model, one disk/wing model, and two disk/wing models. The unsteady aerodynamic results revealed that propeller slipstreams affect the wing characteristics by changing the lift and drag. Meanwhile, when the freestream angle of attack is not equal to 0°, the distribution of the pressure coefficient of the blade is very different at various phase positions. Actuator-disk method can greatly reduce the calculation cost and provide satisfied prediction of the trends of the distributed propeller aerodynamic performance.

Zhang Zhitao, Xie Changchuan, Yang Chao
An Ontology Based Single Source of Truth (SSOT) Construction Approach for Aircraft Modeling and Simulation

Single source of truth (SSOT) is a key element in the realization of digital engineering or model-based system engineering. During today’s aircraft development, huge number of models are built and simulated for different design purposes at different development stages. This brings a crucial challenge for achieving and maintaining model consistency. The single source of truth is a solution to meet model consistency challenge. It emphasizes on building an authoritative data source and all other models shall refer back to it, so that no inconsistent duplicate exists and model consistency is ensured.In this paper, an ontology based single source of truth construction approach is proposed. Ontology is brought out to provide a generic and language-independent description framework for aircraft and SSOT can be built based on this ontology. Then, with the model transformation mechanism introduced, which defines the matching relations between ontology and models, the reference ontology can be transformed into different kinds of aircraft design models and also support the transformation between different models. To confirm the usability of this approach, a practice of this approach on a fix wing unmanned aircraft system (UAS) development is introduced According to the practice, the ontology based single source of truth construction approach initially confirms its feasibility and usability.

Yuchen Zhang, Chuangye Chang, Weijia Wang, Gang Xiao
Experimental Investigations on Aerodynamic and Psychoacoustic Characteristics of Loop-Type Propeller

This study proposes a low noise loop-type propeller that provides lift by loops rather than blades, aiming to reduce the noise sound pressure level (SPL) and improve sound quality. Noise data were acquired via a hover stand test in an anechoic chamber. We then made a comparison of the SPL and psychoacoustic metrics (loudness, sharpness, roughness, and fluctuation strength) of the DJI Phantom 3 propeller and the loop-type propeller using acoustic analysis. Furthermore, we discussed the tonal noise and broadband noise components of both propellers. The aerodynamic force and torque results show that the two propeller types provide very similar figure of merit values, while the DJI Phantom 3 propeller shows higher lift and torque coefficients. In terms of aeroacoustics, the loop-type propeller can reduce the noise SPL at different positions under 5400 rpm. With Fastl and Zwicker's psychoacoustic annoyance model, we predict and demonstrate that the loop-type propeller can improve a human's psychological response to the sound via higher-level broadband noise and lower-level tonal noise components.

Jianwei Sun, Koichi Yonezawa, Eiji Shima, Hao Liu
Effect of Dynamic Micro Vortex Generator on Corner Shock Wave Boundary Layer Interactions Based on DES

Based on the OpenFOAM numerical simulation platform, the dynamic mesh technology and Detached-eddy simulation (DES) are used to study the flow field characteristics of corner shock wave boundary layer interactions (SWBLI) when the micro vortex generator (MVG) moves downstream at a certain speed. The inlet Mach number is 2.5, and the speed of the MVG moving downstream is 0 m / s, 10 m / s and 20 m / s. The results show that when MVG moves downstream, the high pressure area in SWBLI region will gradually weaken or even disappear. The boundary layer thickness of SWBLI will gradually decrease; this shows that dynamic MVG has a good control effect on SWBLI. As the movement speed increases, the control effect becomes more obvious. MVG controls SWBLI through the wave-system structure generated by MVG and the streamwise vortex pair of wakes.

Yong-sheng Zhao, Jun-fei Wu, Jian Zhou
Aerodynamic Characteristics of A Compound Deflected LEF/TEF Rotor

In order to analyze the comprehensive influence of a compound leading edge flap (LEF) and trailing edge flap (TEF) on the rotor aerodynamic characteristics of a 2-D airfoil and a 3-D rotor, a grid generation method including the LEF/TEF is developed. Based on the Navier–Stokes equation, a high-precision analysis method of aerodynamic characteristics under compound deflected LEF/TEF is established. Through a numerical analysis, variations of the aerodynamic characteristics under different LEF/TEF deflection states are studied and the effects of compound flap deflections on the leading-edge vortex and shock wave in 2-D and 3-D cases are investigated thoroughly. The results indicate that the LEF deflects downward and the range of the leading-edge vortex decreases; the pressure distribution on the leading edge of airfoil with differently deflected TEF is nearly identical at a large Mach number. The effect of the compound deflected LEF/TEF on the aerodynamic characteristics is approximately equivalent to the superposition of the separate deflection effects of the LEF/TEF.

Hualong Wang, Xiayang Zhang, Guoqing Zhao, Qijun Zhao
Analysis of Aerodynamic Characteristics and Influence of Parameters of the Quad Tiltrotor Aircraft

In this work, computational fluid dynamics (CFD) and momentum source methods are combined to perform aerodynamic characteristic analysis and to determine the influence of aerodynamic layout parameters on the design of a quad tiltrotor aircraft. First, a numerical method is established to simulate the whole-aircraft interference flow field of a quad tiltrotor aircraft. Then, the aerodynamic characteristics of the rear rotor under various interferences are analyzed based on the established CFD method, and the interference mechanism of the quad tiltrotor aircraft explored. Finally, an influence analysis of parameters under different rotor rotation direction combinations, wing installation angle, rotor transverse distance, and rotor longitudinal distance are carried out. The numerical results indicate that the aerodynamic characteristics of the rear rotors are greatly affected by whole aircraft interference in forward flight, and that the thrust of the rear rotor is reduced by 36.7% relative to an isolated rotor. The left rotors and right rotors rotate in opposite directions, which is conducive to greater wing lift. Increasing the longitudinal and transverse distance between the front and rear rotors can reduce interference of the front wings with the rear wings and improve the aerodynamic characteristics of the rear wings.

Jinshuai Shi, Muyang Lin, Xiayang Zhang, Qijun Zhao, Guoqing Zhao
A Two-Step Optimization Method Using POD-Based Geometric Parameterization for Aerodynamic Shape Optimization

A two-step optimization method is proposed for aerodynamic shape optimization (ASO), and the efficiency of the 2nd-step optimization is improved by proper orthogonal decomposition-based (POD-based) geometric parameterization. The two-step optimization method is constituted by particle swarm optimization (PSO) combined with Hicks-Henne shape functions and steepest descent algorithm (SDA) bonded with POD-based parameterization method. After the 1st-step optimization, superior data (SD) given with better values of objective functions which can satisfy all the constraints are filtered to extract POD bases. The POD bases only cover design space near the 1st-step optimized solution and can parameterize the geometric shape with fewer design variables (DVs). Fewer DVs and smaller design space can improve the efficiency of the 2nd-step optimization,. The two-step optimization method is validated by a case of NASA Rotor 37 aiming to increase peak adiabatic efficiency. The efficiency of the 2nd-step optimization is improved by using POD-based geometric parameterization, which is proved by comparing with SDA using the conventional parameterization method.

Chenliang Zhang, Yanhui Duan, Guangxue Wang, Hongbo Chen
Anti-icing Performance of Engine Inlet Cone by Hot Air Film

Numerical simulation and icing wind tunnel test were carried out to understand the anti-icing performance of engine inlet cone by hot air film. The cones with different slots and holes were investigated under the typical icing conditions. For slot film structures, as the inclination angle decreases from 90° to 30°, the heating efficiency increases about 17% and the amount of ice decreases about 26% under the same hot air mass flow rate. Compared with inclination angle, the effect of slot width on heating and accretion is much weaker. For circular hole, as the diameter of the holes decrease from 6 to 3 mm, the averaged wall temperature increases by about 2–5 °C, and the amount of ice decreases more than 50%. The results of icing wind tunnel tests agree well with numerical results. Increasing the temperature and flow rate of hot air can obviously reduce the amount of ice on the cone. Overall, reducing the slot inclination angle and diameter of circular hole can both improve anti-icing performance.

Xinwei Jiang, Guochao Liu, Yundan Li, Qi Jia, Jianjun Zhou
Aerodynamic Characteristics of Close UAV Formation

The aim of this work is to study the aerodynamic characteristics of solar unmanned aerial vehicle (UAV) wake surfing. First, the application of the following methods to this kind of aircraft was compared: rapid method (i.e., horseshoe vortex model, lifting line theory and nonlinear lifting line theory) and computational fluid dynamics (CFD). The results showed that nonlinear lifting line theory (NLLT) is more accurate in the calculation of the aerodynamic force. The flow field of wake surfing was evaluated with CFD and the local angle of attack was calculated with a local lift coefficient from the CFD results. It was found that the linear method that added induced angle of attack directly to the rear wing overestimated the influence of wing-tip vortices on the rear aircraft. The classical vortex model was introduced to explain this error in a way of vortex decay, which means the rear wing can influence trailing vortices of the lead wing. The error is quantitatively studied by introducing vortex core radius into the calculation, and it was found that the closer trailing vortices act on the rear plane, the stronger the decay shows. After analysis of the characteristics, the vortex core radius was finally introduced into the fast method to modify errors caused by vortex decay. The modified method can reduce the amount of computation in engineering evaluation and give more accurate predictions.

Ziyu Li, Zhengping Wang, Zhou Zhou
Conceptual Design and System Level Analysis of Tilt-Duct eVTOL Aircraft

To take the potential aerodynamic advantages of ducted fan, a novel aircraft configuration utilizing tilt-duct eVTOL concept is developed in this study. The current paper focuses on the overall performance estimation of the tilt-duct aircraft via conceptual design and system level analysis for a typical eVTOL design mission with maximal take-off weight of 2500 kg, 4 passengers, 241 km/h cruise speed, 120 km flight range including reserves. For state-of-the-art battery and electric motor technologies, technically feasible solutions have been found to satisfy the design requirements. For design passenger payload and flight range mission, the maximal take-off weight is 2443 kg. The work in this paper lays the foundation of more detailed disciplinary studies and optimization for tilt-duct eVTOL aircraft, which could further help the real-world implementation of environmentally friendly, cost-effective and safe urban air mobility.

Jiechao Zhang, Yaolong Liu, Tianhong Jiang, Yao Zheng
Research on Wind Tunnel Test of the Total Pressure Probe Layout for Civil Aircraft

The total pressure probe of civil aircraft plays an important role in the atmospheric data system. It is used to measure the total pressure of the incoming flow during flights. The total pressure probe should be installed at a correct position on the nose which local angle can be covered by local angle requirement of the total pressure probe in all flight envelope. According to the above wind tunnel tests of the total pressure probe with part model and full model were designed. The characteristics of total pressure loss of total pressure probe were obtained in the full model wind tunnel test. In the full model wind tunnel test a simplified fuselage with two types of miniature probe was designed to carry out the aerodynamic layout test of total pressure probe. Two types of miniature probes include a similar shaped probe and a seven-hole probe which measure the local total pressure and the local angle of the probe respectively. The variation characteristics of the local angle and the total pressure loss coefficient with the altering fuselage angle of attack and sideslip angle were obtained in the full model test. From the two aspects above, the satisfaction of the local angle and the total pressure loss coefficient for the probe installation position on the aircraft nose can be determined. The design of the miniature probe fills the technical blank of the total pressure probe wind tunnel test model. This type of wind tunnel test can accurately examine the reasonability of the layout of the total pressure probe for civil aircraft.

Xing Zhou, Yuan Zhong, Jiaqi Song
Study on the Dynamic Aerodynamic Performance of Airfoil with Direct Force Measurement

An in-depth study on the dynamic aerodynamic performance of the pitching oscillating airfoil is carried out in the two-dimensional test section of the NF-3 low-speed wind tunnel of Northwestern Polytechnical University. The experimental model is a span-wise three-section force measuring model, and the force measure is only performed in the middle section of the model to reduce the influence of the sidewall interference of the wind tunnel. In the experiment, the transient angle of attack of the model is collected, the inertial force and pitching moment on the middle section of the model are calculated, and the data collected from the balance is subtracted to correct the influence of the model's inertia on the results. The results show that the angle of attack exceeding the positive or negative static stall angles of attack is a necessary condition for the lift and pitch moment coefficients to produce a large hysteresis loops. As the oscillation reduced frequency increases, the dynamic stall is delayed, the lift coefficient hysteresis loop increase, the drag coefficient increase, and the pitch moment coefficient near the maximum angle of attack decrease. When the angle of attack is less than the static angle of attack of stall or exceeds a small range, with the increase of the reduced frequency of the airfoil oscillation, the pitch moment coefficient of the airfoil decreases when it goes up and increases when it goes down. With the increase of the oscillation amplitude, the hysteresis loops of both dynamic lift coefficient and pitching moment coefficient of the oscillating airfoil increase. As the average angle of attack increases, the angle of attack of airfoil enters the positive stall zone more, the lift coefficient hysteresis loop increases, and the minimum pitch moment coefficient decreases. The Reynolds number has no obvious effect on the hysteresis loop of lift, drag and pitch moment coefficients; however, in the downward process, as the Reynolds number increases, the lift recovery advances, and the hysteresis loop decreases.

Yuqin Jiao, Chunsheng Xiao, Dengke Wu
The Research of Correlationship and Critical Ice Shape Acquisition in CIRA-IWT Icing Wind Tunnel

Icing wind tunnel is a mature way to obtain ice shape, and the SAE has conducted ice shape correlation work in the icing wind tunnels as IRT, BRAIT, COX, Goodrich, etc. On the basis of this, the NACA0012 standard model was adopted to implement series ice shapes test in CIRA-IWT, meanwhile, the ice shape was also calculated by CFD for comparison, then the icing parameters sensitivity was analysed. All this work confirmed the correlationship of CIRA-IWT with other icing wind tunnel around the world and was the basis of the following critical ice shape research. According to the FAR25 App.C, the 45 min holding ice shape acquired under continuous cloud icing condition can result in the most severe aerodynamic degradation of aircraft, which called the critical ice shape. A method based on the total water collection (TWC) was proposed to determine the critical icing parameters such as velocity, altitude, temperature, LWC and MVD. Furthermore, the method was verified in CIRA-IWT by observing the ice shedding at the end of 45 min duration and by the comparison of ice shape calculated by CFD. Finally, the critical ice shape prediction method based on the combination of CFD and icing wind tunnel test was the future research direction of this paper.

Bai Feng, Yan Wei, Li Haixing
Parametric Studies of the Fuselage Geometry of a Double-Bubble Aircraft Configuration

The double-bubble fuselage aircraft configuration is a suitable aircraft concept for achieving higher environmental performance. This study aims to improve the aerodynamic performance of the fuselage geometry design of a double-bubble aircraft configuration (JAXA's technology reference aircraft, TRA2035A). In this regard, to understand the aerodynamics of the fuselage geometry, CFD analyses were conducted considering several geometric parameters, such as the height and length of the nose and tail parts. The parametric studies regarding only the fuselage geometry showed that higher nose and lower tail positions improved the lift-to-drag ratio and reduced the trim drag by nose-up pitching moment. Therefore, the fuselage geometry of a wing-body configuration was re-designed. The re-designed geometry showed a higher aerodynamic performance than a baseline geometry. Moreover, the aerodynamics of the wing-body configuration exhibited the same trend as that of the fuselage-only configuration.

Kenya Takahashi, Ryutaro Furuya, Toshiyuki Nomura, Dongyoun Kwak, Katsuyoshi Fukiba
Wind Tunnel Free Flight Model Attitude Determination Based on Non-contact IR Optical Measurement

Model position and attitude measurement is one of the key techniques for wind tunnel free flight test. The state-of-the-art model attitude is determined by a wind vane sensor mounted near the nose of the model, angular sensors or attitude and heading reference system located in the model. However, the large wind vane sensor will influence the flow distribution of the model surface, which increases the difficulty of system modelling and identification. The inner sensors occupy the limited space in the model, making it more difficult to design model and inertia similarity match. This paper proposes a solution using OptiTrack system mainly applied in motion capture for model position and attitude based on wind tunnel free flight reality. Static test and Φ3.2 m wind tunnel dynamic test results show that the static position measurement accuracy is better than 0.07 mm and angular accuracy is better than 0.05°. Dynamic tests indicate the good following performance of the system, meeting the no less than 100 Hz requirement of free flight test data update.

Litao Fan, Bowen Nie, Yong Jiang
Design and Performance Analysis of a Fuel Cell Powered Heavy-Lift Multirotor Drone

The growing use of unmanned aerial vehicles (UAVs) in various sectors has created market demands for drones which can carry heavier payloads and have longer flight endurance. As the endurance of a UAV is highly dependent on the weight of the payload that it carries, drone makers are exploring new and more efficient power sources to produce UAVs that can fly longer while carrying a heavy payload. Previous studies have mainly presented the potentials of fuel cell (FC) as a power source solution for mini to small sized fixed-wing UAVs. The design considerations and performance when implementing the FC system in larger sized UAVs for the purpose of carrying heavier payload and have extended flight endurance has yet been evaluated. This paper describes the design and construction of a 39 kg fuel cell powered heavy-lift medium sized hexacopter and presents the results of its performance, using an analysis of real data from flight tests. It also discusses the possibility of extending flight endurance by increasing the amount of hydrogen fuel using a larger hydrogen tank. In order to provide enough power for the 39 kg medium-sized hexacopter, a 5.2 kw hydrogen fuel cell system was produced by linking two 2.6 kW class power modules. Experiments were conducted to check the performance of the 5.2 kW hydrogen fuel cell system before mounting the system into the hexacopter. Results showed that hydrogen discharge pressure of the hydrogen tanks distributed in Korea cannot cope with the required pressure of 1 bar when loaded with the 5.2 kW power module, so a regulator was used to control discharge pressure. The hexacopter was specially designed to be loaded with the linked FC systems and optimized in terms of weight and structural strength. Structural strength analysis was performed using NASTRAN and PASTRAN to check the design safety margin of the composite skin and the internal structures during structural optimization. Special design considerations were also given to the airflow intake to ensure that the FC system was provided with enough airflow to reach high performance. The flight control system communication and power platform were also modified to adapt to the linked power modules. Flight tests were performed to check the performance and endurance of the UAV. The flight test results demonstrated the feasibility of the designed fuel cell powered heavy-lift hexacopter to perform stable flights. The consumption rate of hydrogen mass was evaluated from the flight test results and the flight time was estimated for the case of a larger hydrogen fuel amount. The consumption rate of hydrogen mass was calculated to be 6 g/min and it was estimated that the flight time could be increased by 6 min by using a larger hydrogen tank to store an additional 36 g of hydrogen fuel.

Allie Foong Yi Chia, Kyoung Moo Min
Multi-objective System Optimization of Suborbital Spaceplane by Multi-fidelity Aerodynamic Analysis

An effective approach has not been established for the reusable launch vehicle, particularly for spaceplanes, whereas the conceptual design method has been established to some extent for aircraft. Spaceplanes that have wing like airplanes flight various environment, and the airframe and trajectory design problem are closely linked. Therefore, the multidisciplinary optimization method is required to optimize the airframe and the flight trajectory design at the same time in the spaceplane conceptual design. Due to computational cost constraints, a low-accuracy Computational Fluid Dynamics (CFD) method was used in a previous study to evaluate the aerodynamic characteristics of the airframe. This has been a source of concern regarding the accuracy of optimization calculations. In this study, a multi-fidelity approach is applied where the low fidelity and high-fidelity CFD are used in a complementary way. This surrogate model was connected into the manned spaceplane’s Multidisciplinary Design Optimization (MDO) framework. As the load limit was reduced, the wing area grew larger and the initial mass increased.

Shintaro Tejika, Takahiro Fujikawa, Koichi Yonemoto
Experimental Validation of Nonlinear Coupled Constitutive Relations in Continuum Flows

The nonlinear coupled constitutive relations (NCCR) model has recently gained great success in stable numerical predictions of high-speed flows over three-dimensional complex flight configurations. However, there is an urgent demand for performing essential experimental measurements to validate the new theoretical model before further promotion to engineering applications. In this paper, the numerical simulations of the NCCR model with Jiang’s undecomposed algorithm and Navier–Stokes-Fourier (NSF) equations by in-house code are in comparison with experimental measurements to validate the applicability and accuracy of the NCCR theory in hypersonic flows. The wind tunnel experiments of HB-2 model were conducted at Ma 5 with five different angles of attack ( $$\mathrm{\alpha }=\pm 4^\circ , \pm 2^\circ , 0$$ α = ± 4 ∘ , ± 2 ∘ , 0 ) in the Φ120 mm hypersonic wind tunnel. It was found that the flow field simulation results by NSF equations and the NCCR model, including aerodynamic force and shock wave structure position, are consistent with the experimental results. Most importantly, the NCCR model could completely recover the flow field solutions of NSF equations in hypersonic continuum flows. Thus, the accuracy and physical consistency of the NCCR model is critically examined and validated for hypersonic flow in continuum region.

Shuhua Zeng, Yunlong Qiu, Zhongzheng Jiang, Weifang Chen
A Propeller Evaluation and Selection Tool for Multicopter and VTOL Design

Along with the emergence of electric vertical takeoff and landing (eVTOL) aircraft arose the need to predict propeller performance, namely thrust and power, at unconventional free stream angles. For cases where this angle is either $$0^\circ$$ 0 ∘ (hover) or $$90^\circ$$ 90 ∘ (forward flight), it is not difficult to estimate the performance. However, in the case of long-range eVTOLs, where the free stream angle is usually in between these angles, the propeller behaves differently. Blade element theory (BET) models can account for this but are impractical due to their large number of geometric parameters. As a substitute, we present an empirical model built on wind tunnel data that uses only propeller diameter and pitch to define the geometry, while capturing the effects of free stream angle on thrust and power. Herein, we discuss the experiment used to gather data for our model, its mathematical derivation, validation against an in-house BET model, and limitations. We end with an example demonstrating how to apply the model to solve the non-trivial problem of selecting an appropriate propeller for a tiltrotor aircraft with maximum range.

Franco Maurice Staub, Yuji Shimizu, Dai Tsukada, Shosuke Inoue, Emery Premeaux, Chris Raabe, Takeshi Tsuchiya
Design and Development Status of Experimental Winged Rocket WIRES#015

Research on winged reusable space transportation systems and flight demonstration by experimental winged rockets called WInged REusable Sounding rocket (WIRES) have been conducted by researchers and laboratory graduates since 2005. WIRES#015 has been developed by Tokyo University of Science, the university start-up SPACE WALKER Inc., and JAXA in collaboration with the German Aerospace Center (DLR) and Swedish Space Corporation (SSC). It has a total length of 4.6 m, launch mass of 1000 kg, and maximum flight altitude of 5.5 km, which is a comprehensive technology demonstrator for technical issues such as the liquid oxygen/liquefied natural gas engine, reaction control system, airframe structure, and propellant tanks made of carbon fiber reinforced plastics and autonomous guidance, navigation, and control systems. Currently, WIRES#015 is in the sustaining design phase. The airframe structure will be manufactured by March 2022 in cooperation with Toray Carbon Magic Co., Ltd. and Mooncraft Co., Ltd. and subjected to structural tests. In 2023, the captive firing test (CFT) will be conducted. The first flight demonstration will be performed in early 2024. Three additional flight demonstrations are planned from 2025 to 2026. In this paper, we describe the development status of WIRES#015 and its development schedule.

Akira Watanabe, Koichi Yonemoto, Takahiro Fujikawa, Takahiro Matsukami, Tsuyoshi Otsuki, Yuto Kitazono, Yasuhiro Koshida, Masaaki Murakami, Tomataka Watanabe, Hina Atarashi, Hidetoshi Takeyama, Shintaro Tejika, Yuito Fujii, Raizo Matsuda, Yusuke Mine, Ayaka Yamazaki, Sho Yoshida, Toshiki Morito
Multiobjective Optimization of a Highly Maneuverable Supersonic Airfoil Using Multifidelity EGO

In supersonic flight regimes, maneuverability reduces owing to the retreat of the aerodynamic center. This paper describes an optimization that objective is the shift forward of the aerodynamic center to solve the maneuverability problem. For this optimization, multifidelity efficient global optimization (EGO) was used, which is an extension of EGO that uses a surrogate model for an efficient search and allows using two different evaluation functions with different fidelities. B-spline curves, which allow for a flexible airfoil shape, were used to generate the airfoil shape. We could determine the type of the airfoil shape that would move the aerodynamic center forward from the pressure distribution and control points of the B-spline curve based on the optimization results. Therefore, we succeeded in finding the geometric features that determine the tradeoff between the aerodynamic center position and lift–drag ratio. The contribution of the geometry of the airfoil to the aerodynamic center is discussed herein.

Tomotaka Watanabe, Koichi Yonemoto, Takahiro Fujikawa, Ayaka Yamazaki
A Computational Study on Unsteady Aerodynamic Forces Around a Pitching Airfoil with Shock and Shock-Induced Separation

The present study computationally investigates the characteristics of unsteady aerodynamic forces around an oscillating airfoil in the transonic flow regime. Particular attention is paid to the role of shock wave and shock-induced boundary layer separation in unsteady aerodynamics. The Reynolds-averaged compressible Navier–Stokes equations are solved with SA and SST turbulence models. A well-known forced-pitching NACA64A010 airfoil experiment (Sanford and Gerald in AIAA J 11:1306–1312, 1980, [6]) is simulated, and the freestream Mach number, Reynolds number, and reduced frequency are set to 0.8, 1.2 × 107, and 0.2, respectively. The pitching airfoil with mean angles of attack of 0°, 2°, 4°, and 6° having the amplitude of 1° is parametrically simulated. It is observed that at the mean angle of attack of 0°, there is a phase delay of the lift coefficient against the angle of attack due to the delay of a shock wave movement over the airfoil surface. In contrast, a phase-advanced feature of unsteady aerodynamics appears in increasing the mean angle of attack (e.g., 4° and 6°). There is a phase transition of unsteady aerodynamics from the delay to advance, significantly caused by the appearance of shock-induced boundary-layer separation. The mean angle of attack around 3° may correspond to a transition condition between the phase-delayed and phase-advanced features. The present study demonstrated that the trend of the unsteady aerodynamic characteristics around the transonic oscillating airfoils largely changes with the mean angle of attack. The shock wave, the shock-induced separation, and their interaction play a crucial role in determining the unsteady aerodynamics such as the phase-delayed or the phase-advanced features.

Noah D. Oyeniran, Hiroshi Terashima
A Multi-fidelity Method-Based Aerodynamic Design Strategy for Preliminary Prop-Rotors

A design strategy to provide suitable baseline prop-rotors for sophisticated applications is presented. Different from those aiming at improving aerodynamic performance of a matured baseline blade fully based on high-fidelity methods, the new strategy is capable of efficiently working out proper layouts with no baseline, which is more beneficial for customizing prop-rotors in industrial applications. The proposed design strategy is introduced through the design process for the initial blades for a self-developed, unmanned tilt-rotor aircraft. The influences of thrust-weighted solidity and thrust requirements in cruise on the aerodynamic performance of the resulting blades are also discussed. Under pre-set design requirements for hover and cruise flight modes, a representative prop-rotor can reach the figure-of-merit of 0.7 in the design hover state and reach cruise efficiency of 0.62 in cruise state due to a relatively lower thrust demand. Using this strategy, part of the design results can be quite close to the requirements, vastly reducing the design space and supporting more advanced optimizations.

Hang Zhang, Qijun Zhao, Guoqing Zhao
Aerodynamics Characteristics and Flight Dynamics Analysis of Multi-body Aircraft

As a new aircraft concept composed of independent aircrafts, multi-body aircraft (MBA) is connected together by hinge in wing tip to wing tip configuration. The dynamic characteristics of this kind of aircrafts are different from traditional aircrafts. If the independent aircraft is stable, MBA will be unstable in this configuration. In this paper, a dynamic modeling method of MBA was proposed. The state-space vortex lattice method (VLM) was used for the calculation of aerodynamics coefficients. Newton–Euler equation was used to obtain the flight dynamic model with absolute coordinates. In addition, trim and stability analysis were presented based on the dynamic model and the particular characteristics of this aircrafts were illustrated.

Chao An, Linpu Wang, Changchuan Xie, Chao Yang
Hydrophilic Coatings for Natural Laminar Flow

The presence of surface contaminants on the wing leading edge of aircraft can expedite boundary-layer transition from laminar to turbulent. For aircraft wings to maintain natural laminar flow and reduce skin-friction drag, it is important to ensure their surfaces remain free from contaminants and/or insects. Japan Aerospace Exploration Agency has recently developed a method for maintaining laminar flow over wings. It involves the application of hydrophilic coatings on wing surfaces, which can be subsequently washed clean with water. These coatings prevent insects from adhering to the wing surface. The results of the coupon and ground tests performed in this study reveal that compared to conventional aircraft coatings, the proposed hydrophilic coatings demonstrate washabilities improved by up to 20%.

Hidetoshi Iijima, Yoshimi Iijima, Takeo Suga, Hidehiko Minami, Kazumi Takao
Modeling of Crossflow-Induced Boundary Layer Transition

Analytical models for the primary and secondary instabilities of three-dimensional boundary layer are developed by assuming locally parallel shear flow and scale similarity at each position. More specifically, both linear and nonlinear growth rates of the stationary crossflow modes are estimated as functions of a normalized wavenumber and a Reynolds number, which are locally defined by the characteristic velocity and length scales of the inflectional profile of shear flow. By choosing one of them as the primary mode, the maximum linear growth rate of the secondary mode is also estimated by assuming scale similarity. The quantitative expressions of these scaling laws are determined so that they agree with several results of direct numerical simulation and numerical stability analysis. Transition to turbulence is triggered when a secondary mode grows nonlinearly. In other words, if a nonlinear primary mode is less unstable to secondary instability, it may not cause transition but laminarize the boundary layer. For various three-dimensional boundary layers, our models are expected to be useful for quickly predicting whether a certain crossflow mode leads to transition or laminarization under a certain disturbance level.

Makoto Hirota, Yuki Ide, Yuji Hattori
On the Characteristics of Anti-contamination Devices on Attachment-Line of Subsonic Transport Aircraft

The natural laminar flow (NLF) wing design approach effectively reduces the fuel consumption and environmental impact by reducing the friction drag for a subsonic civil transport aircraft. This study conducts a numerical analysis based on the Reynolds-averaged Navier–Stokes equations applied to a technology reference aircraft, TRA2022 (120 pax) developed by Japan Aerospace Exploration Agency (JAXA), to analyze the characteristics of a passive flow control device known as the anti-contamination device (ACD), on an attachment line. It is introduced to prevent the propagation of contaminated flow from the fuselage to the wing. Three different types of ACDs are considered: split ACD, chevron ACD, and streamwise groove. The split ACD can provide a wider area to apply the NLF design because it is placed at the wing-body junction. However, it significantly affects the aerodynamic characteristics. Conversely, the chevron ACD and streamwise groove do not significantly affect the characteristics. The split and chevron ACD do not cause flow separation or additional disturbance to isolate the contaminated flow. The streamwise groove includes a turbulent vortex along the groove along with the mechanism used to prevent the contamination flow, which corresponds to this vortex. The effect of the angle of attack between the end and beginning of the cruise flight is analyzed. The change in the boundary layer thickness immediately upstream of the apex of the chevron ACD is approximately 14% of the ACD height, and it can be increased locally based on the ACD shape. The span of the split ACD and the height of the chevron ACD must be considered this effect to ensure their decontamination characteristics.

Keisuke Ohira, Naoko Tokugawa
Preliminary Investigation on Flyover Noise of Propeller-Powered Aircraft Using a Phased Microphone Array

With people’s heightened awareness of environmental protection, the noise level of large civil aircraft during take-off and landing are monitored with strict restrictions by the ICAO (International Civil Aviation Organization). This has triggered people’s interest in learning about the noise characteristics of aircraft components. Hence, researchers are developing phased microphone array test technology to analyze the characteristics of external noise sources of civil aircraft and verify the effect of noise reduction design. At present, most aviation research institutions around the world, have developed their own phased microphone array test technology. They have also carried out the corresponding flight noise test verification, widely used in aircraft low noise design and verification. To better understand the functioning of the external noise of civil aircraft and development the flight noise test capability of civil aircraft, COMAC Beijing Aircraft Technology Research Institute (BATRI) has also carried out the research on the flyover noise test technology. This research, based on a phased microphone array, has designed microphone array and built a set of phased microphone array test systems. For the research, a multi-arm spiral array with a diameter of 15 m and 84 channels were arranged, at the end of the runway of Zhanghe airport in Hubei Province. The domestic single-engine propeller-powered aircraft is used for multiple flights, such as different flight altitudes, different flight speeds, landing gear put-down, and retraction configurations. The test results provide effective test data. Based on this test data, Doppler effect correction is carried out on the flight test data, the test data of different working conditions are compared and analyzed by using the conventional Beamforming method and the noise source identification results are analyzed. The test data processing results show that the main noise source of single-engine propeller aircraft is the blade passing frequency noise of the propeller, and the pure tone noise has obvious characteristics, which are conducive to data correction. The results also indicate that the conventional beamforming noise source inversion algorithm can accurately identify the noise location. However, the sound source location accuracy has a certain deviation due to the influence of array design and aircraft flight condition control. This test has verified the microphone array test system, test process, data post-processing process, and accumulated field flight noise test experience. It also lays a foundation for later large-scale array and large-scale civil aircraft flight external noise source positioning tests.

Zhang Yingzhe, Zhang Weiguang, Lin Dakai, Liu Peiqing
Natural Laminar Flow Design for Highly Swept Wings

Natural laminar flow design has been applied to the next generation of subsonic aircraft, which are expected entry into service in the 2020s, with the aim of reducing greenhouse gas emissions, which is one of the Sustainable Development Goals. The natural laminar flow wing design was applied to highly swept wings, such as the main wing and the vertical tail plane of TRA2022, which is defined as a 120-pax technical reference aircraft in subsonic aircraft research of the Japan Aerospace Exploration Agency. An automatic inverse design system with inverse problem design tools and flow solvers outside of the system was developed. A natural laminar flow design was carried out more efficiently with the help of various innovations. A natural laminar flow design was successfully achieved for not only the main wing but also the vertical tail plane. In both cases, the laminar flow region was extended by more than 20% in the area. This result shows that the natural laminar flow wing design technology of the Japan Aerospace Exploration Agency is effective for highly swept wings. The objective of the natural laminar flow wing design of the Japan Aerospace Exploration Agency has been achieved.

Naoko Tokugawa, Kosuke Toyoda, Fumitake Kuroda, Yoshine Ueda, Takahiro Ishida
Research on Aerodynamic Characteristics of General Long Endurance Vertical Take-Off and Landing Fixed Wing UAV

This paper proposes a new type of pure electric vertical take-off and landing fixed wing general long endurance UAV, and carries out a series of design, calculation and analysis on it. Firstly, use CATIA to perform three-dimensional modeling of the UAV, determine the aerodynamic shape of the UAV, and analyze the calculation method. Secondly, use CFD to divide the mesh, set the boundary conditions, and verify the independence of the mesh. Finally, Fluent is used to calculate the aerodynamic characteristics, the lift-drag characteristic curve, the pitching torque curve, the corresponding pressure cloud graph and the aerodynamic characteristics of the whole UAV under different lateral wind speeds at different attack angles are obtained. The calculation results show that the aerodynamic characteristics of the UAV are good and meet the design index requirements.

Zhan-ke Li, Liang-yang Zhang, Si-jia Zhang, Hai-bo Wei, Hai-yang Han
Flux Limiting Schemes for Fine Resolution Detonation Wave Cell Structure

The dynamic feature of the detonation wave is characterized by the cell structure formed by the traces of triple points in the detonation wave-front and transverse shock wave behind the detonation wave. In this paper, detonation wave is numerically traced by the local maximum pressure and the fully resolved results for the dynamic structure of the detonation wave is studied. A comparative study is carried out for the selection of different numerical schemes which can provide fine resolution without sacrificing the overall accuracy and robustness. For this study, a weak condition is purposefully chosen in such a way that making a single detonation cell to propagate in a two-dimensional (2D) channel of unit width. Grid resolution study has also been carried out for a wide range of radial grid spacing (Δy) By considering the finest resolution results in 256 million grid points in 2D simulation, a moving computational window technique is adopted to calculate only the vicinity of detonation wave front where important physics happens. For low resolution cases, numerical fluxes are evaluated by Roe, RoeM, AUSMDV and AUSMPW+ schemes and high resolution was achieved by 3rd-oder MUSCL, 3rd-order WENO, 5th-order WENO and 5th-order oMLP schemes. This paper shows the full details about the comparative performance of each scheme to capture dynamics structure of detonation cell structure including the dynamic maximum value on Von Neumann spike.

Jaehoon Ryu, Mohammed Niyasdeen, Jeong-Yeol Choi
Experimental Investigation of Performance for Korean Electric Unmanned Helicopter on Ground

Korean Electric Unmanned Helicopter(K-EUH) has been developing to achieve the major performance target of 60 km flight distance with 20 kg payload. The first prototype weighing 126 kgf was manufactured and was tested. Objectives of this ground test is to investigate whether the helicopter has sufficient performance for take-off, and to acquire know-how on the operation of the electric power system. The ground test equipment has loadcells for measuring thrust of the main rotor, and the tail rotor. Rotating power of rotors is supplied by the battery system of nominal DC 355 V. As a result of the test, it was confirmed that the maximum thrust of the main rotor was 1.27 times of MTOW, and the tail rotor also confirmed to have the thrust capable of anti-torque of the main rotor. Experiments have also shown that it is advantageous to configure the electric power system at high voltage. If the required current is low by feeding the high voltage, heat generation and wire thickness can be reduced. The other significant fact is that the voltage drop of battery becomes more severe when the required power is high such as landing approach. Since such a sudden voltage drop reduces the rotating speed of the rotor, a dangerous situation may occur.

Sanghyun Chae, Hong-Tae Yeo, Kyuhoon Lee, Byung Il Yoon
Aerothermodynamic Shape Optimization of a Hypersonic Lifting Body

Hypersonic lifting vehicles experience severe aerodynamic heating while on hypersonic flight regime during atmospheric reentry or unpowered glide. Complex flow interaction is observed in a certain range of flight conditions that results in high heat flux banded region on the windward surface, which should be taken into consideration for thermal protection designing. To obtain desirable surface heat flux distribution and aerodynamic efficiency of a lifting body, shape optimization to reduce the interaction intensity and increase the spanwise distance of the banded heating region on the windward surface was performed in this study. Both single- and multi-objective designs were developed with multiple constraints, including volume, lift-to-drag ratio, the heating rates in stagnation region, and the base section profile of the body. The free-form deformation method was employed for evaluating surface deformation followed by the transfinite interpolation method to update the computational mesh. The range of Pareto front obtained was consistent with the results of single-objective optimization. Additionally, a contradictory and non-linear correlation between the two objectives was elucidated. The optimal solutions exhibited superior comprehensive aerodynamic performance, and the aerothermodynamic shape optimization method was demonstrated to be efficient and practicable.

Haoge Li, Chengrui Li, Weifang Chen, Wenwen Zhao
Experimental Studies on Behavior of Laminar Separation Bubble Formed on the Hofsass Espada Airfoil Near Stall

For micro air vehicles development, it is necessary to investigate the aerodynamic characteristics at low Reynolds numbers. Hofsass Espada airfoil has been developed, and its aerodynamic characteristics at low Reynolds numbers have been studied. This study uses wind tunnel experiments to understand the flow field around the Hofsass Espada airfoil near the stall with a laminar separation bubble. In this study, simultaneous flow visualization and time-resolved surface pressure measurements were conducted. Based on the chord length, the Reynolds number is 6.5 × 104. At the angle of attack near stall, flow around Hofsass Espada airfoil switches between the formation of a laminar separation bubble and a large separated region over the airfoil after the bubble burst. This behavior is similar to that of a relatively thick airfoil, such as the NACA0012. Owing to the smaller leading-edge radius than the NACA0012 airfoil, quasi-periodic switching behavior of flow does not exist.

Masashi Kawakami, Kenichi Rinoie
Influence of Turbulence Intensity on Flow Field Around NACA0012 Wing at Reynolds Number of 20,000

The flow field around an NACA0012 wing is visualized in a towing tank using particle image velocimetry at a low Reynolds number of 2.0 × 104. A semi-span model of the NACA0012 wing with an aspect ratio of 3.0 is used for the experiment. The visualization is performed at the streamwise cross section along the centerline of the wing model. In the experimental investigation of the turbulence effects, a turbulence grid is installed upstream of the wing model. The turbulence is generated by moving the turbulence grid and the wing model at the same velocity. The effect of turbulence on the flow field is investigated by comparing the results obtained without using the grid. The flow field visualization around the wing model is achieved in a turbulent flow. It is found that turbulence leads to the reattachment of the separated boundary layer on the wing surface.

Masataka kase, Makoto Mizoguchi, Hajime Itoh
Analysis of Rotor Noise During Ramp Collective Pitch Increase by CFD Method

Based on FW-H equations and the CFD method, aeroacoustic characteristics of rotor during collective pitch aperiodic variation are calculated and analyzed. At first, a set of analysis method for aperiodic rotor aeroacoustic characteristics is developed. The aerodynamic and aeroacoustic characteristics of BO-105 rotor in hover and NACA rotor in a ramp increase of collective pitch are calculated, and the employed numerical analysis method is validated through comparisons with experimental data. Then, the aeroacoustic characteristics of the NACA rotor during an oscillating collective pitch is analyzed, and the sound pressure peak at different collective pitch is discussed in detail, and some conclusions are obtained.

Xi Chen, Kai Zhang, Qijun Zhao, Weiqi Wang, Siyu Chen
Analyses on Aerodynamic Interactions of Quad-Tiltrotor Aircraft with Variable RPM and Diameter

A CFD method for whole aircraft aerodynamic interactions of the quad-tiltrotor is established, where the momentum source method is adopted for the simulation of the rotor flowfield, and the body-fitted grid method is employed for the simulation of the fuselage. Based upon the established CFD method, the aerodynamic interaction characteristics between the front rotor/rear rotor/wing of the quad-tiltrotor under forward and transition flight state are studied, and the effects of rotor diameter, RPM, and forward flight speed on aerodynamic interaction characteristics of quad-tiltrotor are analyzed. The numerical results show that the use of rotors with large diameter may not be able to meet the performance requirements in high-speed cruise, and the aerodynamic efficiency of the rear rotor is slightly higher. As the diameter and RPM decrease, the differences of aerodynamic efficiency for the front and rear rotors decrease. When the forward speed and the tilt angle of the rotors are small, the efficiency of the rear rotor is lower than that of the front rotor, and the interferences between rotors weaken with increasing rotor diameter and RPM.

Hongliang Wang, Qijun Zhao, Guoqing Zhao, Xiayang Zhang, Jinshuai Shi
Separation Characteristics of a Two-Stage-To-Orbit Winged Rocket by Aerodynamic Interaction Analysis

The booster and orbiter stages of two-stage-to-orbit (TSTO) launch vehicles are separated at supersonic or hypersonic speed regimes under non-negligible dynamic pressure. An aerodynamic interaction exists between the vehicles during the aforementioned separation phase owing to the presence of complicated flow fields with shock wave (SW) interactions. Therefore, the separation operation and attitude control considering such interactions are crucial. In this study, the construction of a database comprising aerodynamic characteristics based on the results obtained using computational fluid dynamics (CFD) analysis is discussed. Based on the results obtained through CFD analysis, the pitch-up and pitch-down moments act on the orbiter and booster stages, respectively when two vehicles are located close. This makes the two vehicles move away from each other. By contrast, the pitch-down and pitch-up moments act on the orbiter and booster stages, respectively, when the two vehicles are located far. The aerodynamic interaction affects the pitching moment may reverse during the separation.

Tsuyoshi Otsuki, Takahiro Fujikawa, Koichi Yonemoto
Aerodynamic Performance of a Low-Speed Blended-Wing-Body Aircraft with Distributed Ducted Fans

The aerodynamic performance of a low-speed, blended-wing-body aircraft with distributed ducted fans was studied. A second-order finite volume flow solver was used to conduct flow simulations around the aircraft. The distributed ducted fans were modeled by applying different boundary conditions at the inlet and outlet of the ducts. Three different geometric models, namely, clean airframe configuration (C1), airframe with a boundary layer ingestion (BLI) fan system (C2), and airframe with a non-BLI fan system (C3), were created to analyze the impact of the BLI fan system on the aerodynamic characteristics of the blended-wing-body aircraft. Then, the aerodynamic performance of these configurations was examined at three design points, namely, climb, cruise, and glide. A comparison of C1 and C3 showed that the installation of the distributed ducted fan system increased the total airframe dissipation in the glide condition. A comparison of the total airframe dissipation of C2 and C3 in the cruise condition revealed that considerable aerodynamic performance improvement was generated by the BLI effects. The aerodynamic performance of the low-speed, blended-wing-body aircraft was sensitive to the installation position of the ducted fan system. The findings of this study provide useful insights into the design of distributed ducted propulsion systems.

Wenyuan Zhao, Jianghao Wu, Yanlai Zhang
Numerical Study of the Aerodynamic Performance of Two Coaxial Flapping Rotary Wings Under Wake Interaction

Flapping rotary wing (FRW) is a composite flapping wing layout proposed in the last decade for micro air vehicle (MAV) design. FRW flaps actively in the vertical direction coupled with passive horizontal rotation, ensuring that the high-lift mechanisms from wing flapping at low Reynolds numbers and the high aerodynamic efficiency of the rotary wing are applied in this wing layout. In actual MAV designs, multiple FRWs have various arrangements, one of which is the arrangement where FRWs locate coaxially and rotate in the same direction. In this type of arrangement, a complex flow interaction exists between wakes from the upper and lower wings, but it has never been focused on in previous research. Thus, this study investigates the effect of wake interaction on the aerodynamics of two coaxial FRWs by using the lattice Boltzmann method and explores the influence of rotation speed, flapping phase difference ( $$\Delta \theta$$ Δ θ ), and wing vertical distance between two wings on the lift and rotating moment, which determine the rotation speed of actual FRW. With inverse flapping (i.e., $$\Delta \theta =\pi$$ Δ θ = π ), the wings interact strongly, and the MAV can reach the maximum rotating moment. An increase in vertical distance between the two wings and rotation speed can weaken the interaction. Our design allows for high rotating speed and low wing loads, thereby reducing the torque requirements on the motor. This study can enhance our understanding of the complex wake interaction produced by multiple FRWs at low Reynolds numbers and provide theoretical guidance for the design of MAVs, especially FRW MAVs.

Songtao Chu, Chao Zhou, Jianghao Wu
Twin Support Vector Regression and Its Application on Aerodynamic Design

Surrogate modeling is playing an increasingly important role in multidisciplinary design optimization (MDO) related to different areas of aerospace science and engineering, Support vector regression (SVR), due to its good behavior related to numerical noise filtering and highly nonlinear function modeling, is promising as an alternative modeling method. However, SVR is time-consuming for high dimension large-scale samples problem. Since Twin support vector regression (TSVR) method shows faster modeling speed, this work aims to evaluate the modeling abilities by numerical examples, and introduce the TSVR method into aerodynamic designs to explore its potential in the aerospace science. Through series of numerical examples, including low-dimensional numerical examples, high-dimensional numerical examples and numerical examples with noises, it is shown that TSVR takes much less time for modeling while keeping high modeling accuracy. Then, taking RAE2822 airfoil as an example, TSVR and SVR is compared, in which TSVR still shows higher modeling efficiency and accuracy. It’s preliminarily proven of the TSVR has great potential in aerodynamic design.

Pei-Xia Lu, Ke-Shi Zhang, Peng-Hui Wang
H-Force of Rigid Rotor in Forward Flight of Multi-rotor

Multi-rotors have certain outstanding features, such as VTOL (Vertical Take-Off and Landing) ability and a simple mechanical system. However, they tend to lack endurance performance. Several types of fixed-wing VTOL have therefore been studied and proposed to minimize this drawback. Our proposal of a novel fixed-wing VTOL is one of these. It has a simple mechanism and has low drag because it is not equipped with control surfaces and stabilizers. Like a multi-rotor, this VTOL is controlled only by the speed of the rotors. In a previous study of our novel VTOL, we conducted a fundamental wind tunnel test of the multi-rotor with and without wings, and acquired data on the drag of the multi-rotor under condition of forward flight. In this paper we describe the results of theoretically examining the drag caused by the rotor. The paper is set out as follows. Section 1 introduces the novel VTOL and the purpose of this research, Sect. 2 describes the wind tunnel tests and the drag measured during forward flight, Sect. 3 describes the theoretical examination of the H-force caused by rotor spin and Sect. 4 verifies the experimental results based on the results of our theoretical study. We draw our conclusions in Sect. 5.

Yasushi Morikawa, Takeshi Tsuchiya
Numerical Investigation of Transonic Flutter Characteristics of a Supercritical Airfoil

This study numerically investigates the transonic flutter characteristics of a supercritical airfoil and discusses the similarities and differences compared to a conventional symmetric airfoil. A wing-section model with two-degree-of-freedom was adapted. An SC2-0610 supercritical airfoil was used for the investigation, and the results were compared with those of the NACA64A010 airfoil. Overall, similar flutter characteristics, such as transonic dip appearance, were observed between the two airfoils. The effect of the shape of the supercritical airfoil appears as the angle of attack effect; for example, the supercritical airfoil with an angle of attack of −2°(0°) shows qualitatively similar flutter characteristics to the NACA64A010 airfoil with an angle of attack of 0°(2°). Furthermore, in the present study, the supercritical airfoil exhibits unique flutter characteristics under high Mach number conditions because of the development of flow separation near the trailing edge on the lower surface. In particular, when the Mach number was 0.875, the flow reattachment near the trailing edge on the lower surface generated a high-pressure region, which induced a unique flutter oscillation under the second torsion mode.

Toma Miyake, Hiroshi Terashima
Control of Wing-Engine-Slat Cut Out Flow Separation Using Nacelle Nozzle Modification

A numerical study on active flow control (AFC) to a swept wing in a take-off configuration was conducted. The focus of this study is the feasibility of the innovative active flow control (AFC) method by using the engine jet, and the control of the jet was achieved through the modification of the nacelle nozzle geometry. The modification included a partial cut or extension at the trailing edge of the nacelle. This AFC method is designed mainly for the flow control of wing-engine-slat cut-out position and it is an alternative design to the outboard nacelle strake. The target configuration integrated with this AFC design was a typical twin-engine jetliner with close-coupled High Bypass Ratio (HBR) engine nacelles. The basic take-off configuration was integrated with regular powered nacelles. Two modified configurations with different engine nozzle designs were also investigated and compared. The surface flow visualization and the spatial particle traces illustrated how the different nozzles affected the engine exhaust flow and the further impact on the aerodynamic performance of the wing. The simulation results showed this innovative AFC device can mitigate the aerodynamic performance degradation caused by the installation of large-diameter nacelles, and suppress the flow separation near the wing-engine-slat cut-out position. Meanwhile, the AFC method did not require additional mechanical mechanisms or any special maintenance, nor had any weight penalty. Therefore, this type of AFC device shows a high application potential in commercial transport aircraft.

Zheng Cui, Jiangang Wang, Xi Du, Dakai Lin, Yihua Cao
Developments of Free-Flight Testing Facility for Aerodynamic Assessment of Martian Entry Capsule

The evaluation of the aerodynamic characteristics of a Martian entry capsule is crucial for future Mars missions. However, the uncertainty in the dissociation rate of CO2 makes it difficult to predict them accurately with a CFD analysis. To improve the prediction accuracy of aerodynamic characteristics of hypersonic vehicles, the development of an experimental database is inevitable. In this study, the free-flight tests with a Martian capsule model were performed with using a light gas gun at Japan Aerospace Exploration Agency. The aerodynamic coefficient was determined by reconstructing the trajectory and attitude from high-speed visualization data of the free-flight capsule. The target velocity to reproduce the flight environment was 4.2 km/s for an ambient pressure of 11 kPa. Currently, with a cylindrical projectile, the muzzle velocity of 4.2 km/s was achieved at the pressure of 11 kPa. The shadowgraph image of free-flight capsule model was successfully obtained at the velocity of 3.4 km/s. The template matching method was applied to a capsule shadow, and the pitch angle was determined as 16.0°. With respect to the yaw angle, an improvement in the analysis procedure is necessary.

Satoshi Nomura, Kyosuke Itabashi, Masahito Mizuno, Kazuhisa Fujita
Multi-fidelity Modeling via Regression-Based Hierarchical Kriging

Advances in computational technology have enabled engineers to obtain large amounts of data efficiently. However, generating datasets from a single source is still burdensome and inefficient, so there have been many studies on data fusion techniques using multiple sources of data: e.g., low-fidelity data from computational fluid dynamics simulations and high-fidelity data from wind tunnel experiments. The hierarchical Kriging surrogate model is one of these data fusion methods which is known to be simple, robust, and accurate. It is based on the Kriging surrogate model, which predicts quantities of interest using an accumulated database based on the interpolation method. However, serious problems arise when it is built based on the noisy datasets for the following reasons: 1) noise in the dataset induces ill-conditioned correlation matrix; and 2) errors at lower-fidelity levels due to the noise accumulate as the fidelity level increases, adversely affecting the accuracy of the final model. These issues can be mitigated by a method called regression-based Kriging, which incorporates the adaptive regression factor (nugget) to the correlation matrix. This paper extends the corresponding regression technique to the hierarchical Kriging surrogate model, which will be called regression-based hierarchical Kriging. Considering that data fusion methods synthesize data with different fidelity levels to build a surrogate model efficiently, the regression-based hierarchical Kriging model can consider the different degrees of noise at each fidelity level. Herein, this model is demonstrated with several numerical examples with artificial noise and is finally validated with real engineering problem. These examples show that the proposed method has better performance than interpolation-based hierarchical Kriging when trained with the noisy data.

Sunwoong Yang, Yu-Eop Kang, Kwanjung Yee
Evaluation of the Use Cases of eVTOLs with High Potential in the Philippines and Thailand

This study aims to select and evaluate the promising use cases of multirotor electric vertical takeoff and landing (eVTOL) aircraft in Philippine and Thai markets. With the high potential of eVTOL aircraft for practical use in Southeast Asia, identifying which countries and use cases are promising is necessary. We conducted interviews with doctors of two institutions with helicopter emergency medical services (HEMS) in Thailand to elucidate current challenges in HEMS and airframe requirements for “emergency medical services (EMS)” use case. Next, we conducted interviews with 12 local stakeholders in the Philippines and Thailand to gain information on the challenges surrounding the implementation of eVTOLs for the “Sightseeing/Leisure” and “Urban Air Mobility (UAM)” use cases within the specified locations. Then, we quantitatively evaluated the stakeholder perceptions using the Quantitative Strategic Planning Matrix to select promising use cases. Finally, we performed Fermi estimate to calculate the market size with constraints. The results showed that HEMS in Thailand needs two-seater eVTOLs to dispatch only a doctor to the field as in Japan; however, the feasibility of this use case is lower than other use cases considering the specifications required by doctors. The Sightseeing/Leisure use case in Cebu is the best use case for eVTOL implementation with stakeholder perceptions in mind, whereas the UAM use case in Bangkok is better for sales.

Aki Nakamoto, Patrisha Armie W. Bas, Masaru Nakano
Numerical Investigation of Droplet Impact on the Surface by Multiphase Lattice Boltzmann Flux Solver

The dynamic behaviors of the micro-sized water droplet collision onto the wings of the aircraft are essential to the flight safety. The details on the small droplet in the airflow in contact with the aircraft wing surface play a quite important role in the ice accretion process. In this paper, multiphase lattice Boltzmann flux solver coupled with phase field method is applied to simulate the water droplet impact onto the solid hydrophilic/hydrophobic surface to further understand the interactions between droplet and surface at mesoscopic level. The reliability and accuracy of the numerical method is validated by the comparison with experimental data and computational results in other literatures, which shows that the solver is capable of predicting the droplet dynamic behaviors. Then, the effects of different physical parameters such as impact velocity, droplet diameter, surface contact angle and impact inclination angle, are systematically studied. The computational results reveal that when the collision is normal to the surface, the water droplet may experience spreading phase, recoiling phase as well as rebounding phase and finally shows the adhesion state or detachment from the surface. The higher velocity and larger diameter contribute to spread the droplet wider and jump higher during the droplet impact process. And a shorter physical time is taken to reach the spreading factor maximum for higher velocity while it is opposite for the droplet with a lager diameter. Moreover, the whole evolutionary process of smaller-sized droplet is accelerated and smaller diameter as well as higher contact angle of the surface advances the droplet detachment from the hydrophobic surface. It is also found that the surface with higher contact angle impedes the droplet spreading and removes the temporal lag of its performance in lifting up the upper end of droplet during recoiling phase and rebounding phase, which is distinct to the results of higher velocity and larger diameter. Besides this, droplet impact with an inclination angle causes reduction on the spreading factor maximum and jump height after detachment from the surface due to the decrease on the normal velocity of the droplet. And the increase of the tangential velocity accounts for the longer contact time with the surface for the droplet, and causes the difference of the spreading factors in spreading directions, which forms an oval contact area on the surface until the droplet detaches. The analysis and quantitative comparison of the temporal morphology evolutions of the micro-sized droplet in this paper help to reveal the interaction mechanism between the different-sized droplets and surfaces with different properties, which can be considered specially in the numerical prediction of the aircraft icing.

Qingyong Bian, Chang Shu, Ning Zhao, Chengxiang Zhu, Chunling Zhu
Computation of Hypersonic Transitional Flows Over Cones Using Gas-Kinetic Scheme Coupled with the Turbulence Model

Accurate prediction of the laminar-turbulent transition with the aerodynamic force and heat in hypersonic boundary flows is essential for the design of a hypersonic vehicle. However, subject to the complex boundary layer mechanism of the angle of attack (AOA) effects, the numerical prediction with high accuracy of hypersonic transitional flows over a cone at a small AOA is still a challenging job in current computational fluid dynamic (CFD) modeling. In this paper, our objective is to improve the prediction accuracy of the traditional turbulent model in the boundary layer flow problems. We employ a gas-kinetic scheme (GKS) strongly coupled with the turbulent kinetic energy equation in the shear stress transport (SST) turbulence model. Langtry-Menter's two-equation transition model is also implemented under the GKS framework. The turbulent relaxation time obtained from the turbulence model is also implemented into the kinetic equation of the BGK model as an enhanced particle collision time related to turbulent fluctuation. We further validate this method for several classic cases of the hypersonic transitional flows over cones with various free-stream conditions at zero AOA. We also compare the numerical solutions by the current GKS coupled with the turbulent kinetic energy equation and traditional Reynolds-Averaged Navier–Stokes (RANS) method. The numerical results confirm the consistency of the transition onset locations predicted by the proposed method with the experiments. The new coupled method also significantly improves numerical accuracy. The heat flux is also in good agreement with the experimental data. The hypersonic flows past a sharp cone are also computed at various small angles of attack ranging from 2° to 4°, where the numerical results agree with the existing experimental data. Using our results we also analyze the impact of AOA on a sharp cone.

Chengrui Li, Hualin Liu, Zhongzheng Jiang, Weifang Chen
Effect of Solidity on Efficiency for High-Advance-Ratio Propeller

The use of a propeller is an option for powered flight on Mars. Powered flight in the Martian atmosphere is quite peculiar in that the propeller operates with low speed under low Reynolds number conditions. The improvement of the performance of the propeller for Martian flight is investigated, focusing on the blade solidity. First using the blade element momentum theory, the effect of the increase in solidity is investigated. Then, in order to verify this solidity effect experimentally, wind tunnel tests are carried out. Two propellers with different solidity are specially designed and tested. Through the measurement of thrust and torque, the effect of the solidity is discussed in terms of efficiency.

Koju Hiraki, Hiroki Kai
Aerodynamic Study of Cone-Derived Waverider as Supersonic Transport

In this research, a new type of supersonic transport (SST) configuration is investigated that can contribute to improve aerodynamic as well as noise (sonic boom) performances, by generating supplemental lift force from generated shock waves. Inspired by the previous study of a cone-derived waverider conceptual design that achieved a significant increase in of lift-to-drag ratio ( $$L/D$$ L / D ), a concave study is examined on the waverider (supersonic) configuration. The main purpose of this study is to increase $$L/D$$ L / D value as well as to reduce the sonic boom strength by designing the lower side of the waverider configuration. The effect of the concave shape added on the bottom side of the body is examined in this study. The numerical simulation approaches are used to evaluate the objective performance functions of $$L/D$$ L / D and sonic boom strength. As a result, the effect of the concave shape added on the lower surface of the body influences the sonic boom strength. Various types of concave models are examined and important design knowledge for the waverider configuration is extracted. Strong shock waves on the lower side of the body cause high-pressure regions to generate higher lift force, and its efficiency depends on the concave position and/or size.

Nomin Buyanbaatar, Yuhei Ishikawa, Wataru Yamazaki
Boundary-Layer Transition Control by Plasma Crossflow Reduction in Swept Wing

The boundary-layer transition caused by the crossflow instability in a swept wing is controlled by a plasma actuator (PA). In this study, the control concept is to suppress the crossflow generated at leading edge of the wing by using PAs. The PAs are installed at the leading edge so that the body force can be applied in the direction that cancels the crossflow. An attempt was made to delay the boundary-layer transition by driving a PA. In a wind tunnel experiment, we investigate the transition position and velocity in the boundary layer at U = 20 m/s. The experimental results showed that the transition position delayed when PA was driven. From the spectral analysis of the velocity in the boundary layer, we found that the boundary-layer transition process changed from crossflow instability mode when PA was driven.

Yuto Miwa, Ikuya Yoshimi, Takashi Matsuno, Dongyoun Kwak
Numerical Investigations of Ground Effect of a Quadcopter

The hovering in ground effect (IGE) of a quadcopter is investigated numerically in terms of flow patterns and aerodynamic characteristics with two different rotor configurations with two different rotor-to-rotor distances. A model of variable-pitch-controlled rotors revealed complicated outwash flow patterns around the quadcopter, showing a pronounced directivity, particularly when the rotor-to-rotor distance was shortened. A narrow and intense outwash (jet) region formed between the rotors. A similar outwash pattern was also found in the diagonal plane. Our results further show that the rotor configuration effect on the aerodynamic performance of the quadcopter can vary significantly with changes in altitude from the ground. Moreover, the rotor configuration with a short rotor-to-rotor distance displayed a relationship between aerodynamic characteristics and altitude that was similar to that of a single rotor.

Koichi Yonezawa, Kazuki Akiba, Hao Liu, Hideaki Sugawara, Yasutada Tanabe, Hiroshi Tokutake, Shigeru Sunada
A Review of Recognition in Military Airworthiness Regulatory Frameworks

Military aviation does not have a globally standardised airworthiness framework like that maintained by ICAO for civil aviation. As a result, the distinct and unique nature of each State’s military airworthiness regulatory framework makes safety assurance difficult during joint operations, collaborations in aircraft development or during procurement of foreign aviation assets. The process of recognition of military airworthiness frameworks can, to some extent, alleviate many of these difficulties and help to build trust, allow resource sharing, enable mutual learning, provide economic benefits, and enhance safety assurance between collaborating Military Airworthiness Authorities. A review of the various military aviation forums, alliances, and collaborations is presented here in which recognition activities are described and evaluated. Attention is then focused on the most successful recognition process currently in use worldwide, namely the European Military Airworthiness Document–Recognition (EMAD-R). The main reason for its success is that it incorporates much of the globally accepted ICAO oversight and airworthiness requirements in its Military Authorities Recognition Question-set (MARQ), thus providing the necessary confidence for recognition between collaborating Military Airworthiness Authorities. This work concludes by examining the EMAD-R process and showing it can effectively be used for recognition of disparate airworthiness frameworks even outside the European paradigm.

Zawar A. Nawaz Bhatti, Nicholas S. Bardell

Structures and Materials

Frontmatter
Material Property Measurement of 3D Printed Carbon Fiber Composite Using a Digital Image Correlation Method

Three-dimensional (3D) printing of carbon fiber composite is a promising manufacturing process for aerospace and other industry fields because it is fast and easy to implement and requires less labor. Knowledge of the material properties of the 3D printed composite is important for applications to aerospace and other industrial structures. In this study, a digital image correlation (DIC) technique is employed to measure the material properties of 3D printed composites. This is a non-contact measurement method, and able to measure the full-field displacement and strain of the structure. This study provides a suitable testing method for determining the material properties of 3D printed continuous fiber composite materials according to ASTM standards using the DIC method.

Feng Quan, Rui Hai Xin, Nam Seo Goo
Research on Helicopter with Emergency Floating System Impact Load Characteristics During Ditching Based on ALE Method

This paper based on the ALE algorithm, the helicopter with emergency floating system load characteristics during ditching are numerically simulated and analyzed. Provide support for the design of emergency floating system, gasbag installation point and personnel evacuation channel. First of all the ditching model of the helicopter and the floats is established, and the accuracy of the simulation model and the calculation method is verified by comparing the results of the scale model and the physical testing, verified the accuracy and effectiveness of the simulation model, modeling method and calculation method. Then by using this method establish the full-size model. Based on this method, the influence of different roll angles, horizontal velocities and vertical velocities on the helicopter impact load during ditching is studied. This paper provides support for the design of emergency floating system, gasbag installation point and personnel evacuation channel. It not only overcomes the difficulty that the full-size prototype cannot carry out physical experiments, but also can effectively shorten the development cycle and save development funds.

Zhijie Feng, Aokun An, Huajin Lei, Chuanyi Ke
Study on shapes of Double Cylindrical Structure for Wing Twist Morphing

We previously developed a twisted morphing wing system based on a double-cylinder structure that consists of outer and inner parts. Here, trade-off studies on the shape of the outer part of the double-cylinder structure are carried out. The open section of the outer part has four beams. The equivalent bending stiffness, equivalent torsional stiffness, and buckling force are calculated for various shapes of the open section using previously derived formulas, and their relationships are obtained. The effects of the shape parameters, namely beam height, width, aspect ratio, and length, on the equivalent bending stiffness, torsional stiffness, and buckling force are investigated, and the optimal cross-sectional shape of the outer part for wing twist morphing is determined. The open section should have low torsional stiffness to achieve wing twist morphing, high bending stiffness, and a large buckling force to bear the aerodynamic load. The results indicate that the smallest equivalent torsional stiffness is achieved by using a cross-sectional shape with a large aspect ratio. Moderately small torsional stiffness and moderately large bending stiffness are achieved with a small aspect ratio. Therefore, the shape of the open section can be designed by considering the trade-off between bending and torsional stiffness. Furthermore, the open section length should be set to avoid buckling of the beam and realize a large twist angle.

Hiroaki Tanaka, Yusuke Arai
Multi-objective Optimization for Resilient Operation of Adaptive Morphing Flap Divided into Span Directions

This study proposes a multi-objective optimization method to enable the recovery of lifting force using the remaining flaps after failure of one flap, such that the flap angle returns to zero. Recoverability is regarded as a resilient operation in this study. Therefore, a multi-objective optimization problem is formulated to minimize the bending moment at the wing root from the structural point of view, to minimize the sum of the flap angle differences for remaining flaps corresponding to the minimum effort of the operation and the difference of the lifting force from the viewpoint of flight mechanics in terms of the flap angles of the remaining flaps, where the lifting force distribution is obtained by vortex lattice method. The validity of the proposed resilient operation is demonstrated through numerical examples, and the efficiency of the adaptive wing divided flaps is evaluated. Additionally, the lifting force distribution and the flap angle distribution along the span direction are shown for different flap failures.

Nozomu Kogiso, Kento Konishi
On the Evaluation and Design Methods of Structural Recoverability

Usually, maintainability, which is often expressed by means of MTTR (Mean Time to Repair), is used to characterize whether it’s easy to repair when the equipment structures have some common damages during the normal service period. But, it’s hard for equipment structures to avoid accidental and unexpected damages in service, such as accident damages, battle damages, etc. And then recoverability can be used to express whether it’s easy to repair and recover when equipment structures suffer some accidental and unexpected damages. After expounding the basic concept of structural recoverability, the significance of structural recoverability to maintaining structural operational integrity, which is categorized into structural static operational integrity and structural dynamic operational integrity, is discussed. Then, parameters Rc(t) (Recovery Degree), MTTRC (Mean Time to Recovery), MRCC (Mean Recovery Cost), and MWTTRC (Mean Work Time to Recovery) are proposed to characterize and measure structural recoverability. Furthermore, the normal evaluation, assessment, and design methods of structural recoverability are put forward primarily. Finally, a brief example of recoverability evaluation for accidentally damaged vehicle front bumper structures of three brands of the same model class is shown through analyzing MTTRC and MRCC, and followed by the recoverability design discussion on an aircraft monolithic backswept wing structure constructed by ribs, stringers and skins. It is also stated here that the concept, measurement, evaluation, assessment, and design methods of structural recoverability are suitable for other structures, such as building structures, bridge structures, and other infrastructure structures, etc.

Yuting He
Simulation and Analysis of Automatic Passenger Door Actuation System

Aiming at the requirement of automatic operation of passenger door of a civil aircraft, a set of mechanism and electric actuation system suitable for all-electric automatic intelligent operation of passenger door is designed. The system can complete the automatic operation of unlocking, opening pressurization prevention door, unlatching, lifting, opening and the electric operation of girt bar mechanism, and has the function of manual operation all the mechanism. In the system design and refines the top-down design method adopted in analysis of systems engineering, and at the same time in the mechanism design of organization performance analysis Real-time adjustment and optimization, to ensure that automation involved institutions include lock mechanism, latch mechanism, pressurization prevention door mechanism, lifting mechanism, open mechanism, gust lock, flight lock mechanism and girt bar mechanism to meet the design performance requirements, the performance of latch mechanism, lifting mechanism, opening mechanism and girt bar mechanism is analysed emphatically. It can be used as the basis to choose the parameters of the downstream transmission train and the power source. It can better design the downstream drive line and choose the suitable power source. According to the two-dimensional diagram of the mechanism of CATIA, a two-dimensional linear frame mechanism model was established. On the basis of considering the size of the entity, the dynamic simulation model was established by using the dynamic simulation analysis software LMS Virtual Lab Motion, and the simulation analysis was carried out by adding constraint joints and mechanical model. According to the Hertz contact force model, the contact model is added by using solid contour, and the empirical data is used to make the model closer to the actual contact, so that the simulation analysis results have a certain reliability. The design and simulation analysis and optimization of the parallel method makes it possible to scientifically analyse and verify the performance of the mechanism before the physical manufacturing and molding, realizes the coordination of design and analysis, exposes the development risk in advance, and reduces the cost and time of the physical verification. After the model design, the wire frame simulation analysis method can be simply and quickly changed to carry out a new round of simulation calculation. There is no need to rebuild the model and change the entity to complete the simulation. Fast multi-round iterative calculation greatly shortens the development cycle. In this paper, it is found that the difference between the two results is small by using MATLA kinematics simulation analysis and local processing calculation, which can be used as the basis for the load design of the passenger door actuation system.

Weijuan Zheng, Yi Wu, Wenjing Zhi, Dongping Liu
Research on the Aerodynamic Failure Load for Civil Aircraft

This paper introduces the failure list selecting methods based on the system function safety evaluation report, and introduces the flight loads calculation methods of the time at the failure occurrence and continuous flight after the failure based on the selected failure list. Taking a large civil airplane as an example, it proposes the principle of failure list selecting, and creates the concept of control surfaces jam biased Normally Encountered Position rather then the extreme deflection. Finally it quantitatively analyzes failure load such as elevator jam, aileron jam, rudder jam, stabilizer jam and flap actuator disengagement. During the procedure it compares with the conventional flight load. The analysis of elevator jam show that it would create an asymmetrical moment at the tail due to asymmetric deflection of the left and right elevators. But the moment is not covered by normal load, which constitutes the critical case of the rear fuselage. And also the flap actuator disengagement would be the serious situation in load design. So both of them should be carried out. This paper comes up with an effective, reliable failure load calculation and analysis method.

Linlin Tan
The Vibration Transfer Path Analysis Based on One Large Passenger Aircraft

The transfer path analysis (TPA) method, which combines the testing and calculating, is a very useful and accurate tool for the structural vibration analysis, and has a wide application in the automotive field. However, the aircrafts have a very complicated loading condition, so the TPA method according to the aircrafts should be studied further. Nowadays, a lot of structural transfer path analysis cases are based on the commercial software, such as LMS and Head. Not mastering the core analysis technology of TPA technology is adverse to the further research in the aviation field, so it is necessary to carry out the foundational TPA technology. This article based on one large passenger aircraft, the vibration transfer paths from the engine and the landing gear to the four seat rail points in the cabin were identified. The vibration contribution and the sensitivity analyses of each path was also analysed. All the above were achieved through basic programming.

Li Yixuan, Li Kaixiang, Liu Jijun
Manufacturability and Deformation Performances of CFRP Twist Morphing Wing Structure with Applying the Electrodeposition Resin Molding Method

To achieve efficient flight by minimizing energy consumption, a morphing wing that allows large and smooth deformation similar to a wing of a migratory bird is required. Previously, the authors have developed a migratory-sized unmanned aerial vehicle (UAV) with a twist morphing wing. The twist morphing wing was manufactured using a carbon fiber reinforced plastics (CFRP) with applying the electrodeposition resin molding (ERM) method, which was developed by authors. In the ERM method, resin impregnation proceeds in a liquid by electrophoresis. Thus, manufacturing process is very efficient because neither pressurization nor evacuation are necessary. On the other hand, although the molding of the twist morphing wing is efficient, an optimal design is important for the twist morphing wing to lower the torsional rigidity to reduce the driving force of the morphing and increase the bending rigidity to support aerodynamic loads. Therefore, in this study, the spar of the twist morphing wing was designed to be arranged in a triangle. Furthermore, the twist morphing wing structures were fabricated by the ERM method, and by a 3D printer for benchmarking. So, torsion and bending deformation properties of the twist morphing wing were confirmed under the aerodynamic loading condition, and validated by finite element analysis. As a result, it was confirmed that the manufacturing efficiency of the ERM method was superior to that of the 3D printer. Additionally, the torsional rigidity of the twisted morphing wing was almost the same, regardless of the manufacturing method. Furthermore, the bending rigidity of the specimen manufactured by the ERM method was higher than that of the 3D printer.

Kazuaki Katagiri, Daekwi Kim, Choong Sik Park, Sonomi Kawakita, Masato Tamayama, Koki Kayano, Shinya Honda, Katsuhiko Sasaki, Makoto Yamazaki
Doubly Coupled Antagonistic Shape Memory Alloy Actuator for Morphing Wing

Shape memory alloy (SMA) wires have received considerable attention for use in actuators in morphing wings. This research presents an extension of our previous study on an antagonistic SMA wire actuator to a doubly coupled antagonistic SMA wire actuator. The actuator is composed of one main antagonistic system and two sub-systems. The experimental results demonstrated the potential of the proposed system to drastically reduce the wire length compared to a single antagonistic system.

Shingo Suzuki, Atsuhiko Senba, Tadashige Ikeda, Masato Tamayama, Hitoshi Arizono
Dynamical Analysis of Rotor Loads in Helicopter Circling Flight Condition

In order to study the dynamic loads of a helicopter rotor during circling flight, the aerodynamic and structural dynamic models of the rotor are developed. In terms of aerodynamics, the free wake method based on relaxation iteration is adopted, wherein the wake model is distorted for circulating flight. With respect to structural dynamics, the Rayleigh–Ritz method is employed in mode superposition manner. The dynamic loads and properties of the rotor blade are calculated and analyzed. The accuracies of the modeling methods are verified by a numerical analysis and experimental comparison. Thereafter the influences of circling radii are emphasized. The distortion of wake changes the distribution of induced velocity and the aerodynamic load increases with a decrease of the circling radius. However, the dynamic loads share the same variation trend as the circulating radius.

Xu Zhou, Xiayang Zhang, Bo Wang, Qijun Zhao
Study on Laser Shock Peening Technology for Hole Position of Aviation Parts

In engineering applications, the opening position is a typical stress concentration structure on aircraft components, which is prone to fatigue cracks under alternating loads. In order to study the effect of different power density and laser shock peening trajectory on the residual stress at the opening position of AL7050-T7451 alloy and improve the fatigue life of aviation parts. The results show that when multi-point impact peening is adopted, the peening effect is better when the power density is 3 GW/cm2, and the Mises average stress value is 178 MPa. When single-point impact peening is adopted, the Mises stress values under different power densities have little difference, and the power density has little effect on the impact peening effect. The single-point impact peening with small power density is the best, about 3 GW/cm2. By comparing multi-point and single point, when the laser power density is reduced, the peening effect can be improved by changing the process. The research results have certain reference significance for the opening position of laser shock peened aviation parts, and lay the foundation for improving the fatigue life of laser shock peened hole structure.

Zhi Yang, Xiangyu Ding, Sijie Ma
Calculation Method and Experimental Verification of Optical Fiber Limit Value of Aeroengine Rotor Blade

In this paper, the implementation and validation of the proposed methods for optical fiber limit value calculation were investigated. The intersection distance method (IDM) and finite element drawing method (FEDM) were proposed and compared with the common similar triangle method (STM). The advantages and disadvantages of the three methods were analyzed from the aspects of accuracy, limitation and computing resources. The simulation results were verified by the fan rotor blade dynamic test. By comparing the calculated amplitude obtained by dynamic stress inversion with the measured amplitude, it is found that the calculated results of the three methods are in good agreement with the measured amplitude and could be used to estimate the amplitude under low stress condition. In torsional mode, the errors between the amplitude calculated by the STM, IDM, FEDM and the measured amplitude are 1.8%, 1.3% and 3.1% respectively. In high order mode, the amplitude difference obtained by each calculation method is up to 12%, and the errors between the amplitude by the FEDM and the measured amplitude is the lowest, reaching 0.3%. Furthermore, the errors of different calculation methods under stress higher than 200 MPa were analyzed. In conclusion, the results indicate that FEDM is the most accurate method for any vibration modes, and STM and IDM are also applicable in simplified calculation of low-order vibration modes of blades.

Qing Du, Jian Zhang, Yixiong Liu, Fayong Wu
Analysis of Laser Positioning Milling Process for Aircraft Cabin Door Opening

One of a main characteristics of aircraft manufacturing is to solve a lot of interchange and coordination problems in process preparation and production. Traditionally, analog quantity transfer method is adopted, which can not meet the manufacturing requirements of “low cost, short circle and high quality”, based on the three dimensional design model and the advanced measurement device of aircraft, it is an effective way to establish and adopt the interchange and coordination model of digital quantity transfer to meet the needs of modern aircraft development, with the rapid development of computer and numerical control technology, digital technology is gradually applied to the field of aviation industry, the level of aircraft design and manufacturing technology has been greatly improved. In terms of assembly technology, digital technology is used in some parts, but analog quantity transmission is still the main coordination method of many aircraft manufacturing enterprises, and special tooling assembly is still used in assembly. Hoped that all aircraft manufacturing enterprises can achieve new breakthroughs in digital assembly technology. The application of digital technology will bring great changes to the manufacturing and inspection process. Under traditional conditions, after assembly, the products can be measured and inspected by analog equipment such as tooling and standard measuring tools. Under the conditions of digital manufacturing, due to the application of direct installation technology such as numerical control technology and laser tracker, the necessary tools such as frame and measuring tools are reduced. In this paper, the problem of interchange and coordination in milling the opening of rear cargo door frame of a cargo aircraft studied. Referring to the traditional aircraft interchange and coordination theory, and combined with the characteristics of aircraft digital manufacturing technology. The key technology of aircraft manufacturing interchange and coordination based on digital quantity transfer is studied. The laser aided positioning and milling technology for the opening of rear cargo door frame of a cargo aircraft is based on rear cargo door frame product digital model and Drawings, and with the help laser tracker, the positioning, marking and milling of frame door opening of the aircraft fuselage door frame are carried out. The advantages of this technology way are: it can save a set of tooling for milling the door frame opening, and no matter how to modify the drawing/digital model size of the door frame opening, this technology way can complete the milling of the door frame with the fastest speed, and adapt to and meet the latest size requirements of the drawing/digital model (because the use of laser tracker does not need to spend the time and cost of designing, modifying and manufacturing tooling).

Shengxiong Li
Numerical Simulation of TC17 Titanium Alloy Thin Blade Strengthened by Laser Shock Processing

Laser shock processing (LSP) is a new surface strengthening technology, it can produces relatively large residual compressive stress on the metal surface and a certain depth, and refines the grain on the metal surface, so as to strengthen the metal. In this paper, the stress distribution and deformation of thin blade edge of TC17 titanium alloy under multi-point impact were studied by numerical simulation. By controlling the action distance of plastic wave, the process parameters of single-sided laser shock strengthening of thin blade edge were designed. The results show that the solution time of dynamic shock is 10000 ns according to the energy curve. By changing the boundary constraints of thin blade, such as bottom constraint, cantilever constraint and both ends constraint, to simulate the actual processing situation, the distribution of residual stress in bottom constraint is the most uniform, the increase of residual stress in the processing area is more obvious, and the material deformation is the least. Based on the commonly used processing parameters of 3 mm laser spot and 20 ns pulse width, 30% overlap ratio is adopted to avoid uneven impact or excessive local stress. The residual stress is about 104, 247 and 450 MPa when 2, 3 and 5 J energy are used respectively; the blade edge deformation is controlled at about 4 μm. In this paper, the actual multi-point machining is simulated, which provides a theoretical reference for the selection of the actual machining parameters of thin blade.

Sijie Ma, Xiangyu Ding, Zhi Yang, Jiayu Hao, Wei Xiao, Longgang Fan
Aerodynamic Study for Wing Shapes Based on a Stuffed Bird

A fundamental aerodynamic study for the effect of wing shape, based on a stuffed bird, is presented. Birds fly efficiently and optimally under various conditions by changing the shape of the wing quickly during flight. Aerodynamic forces on bird models with several flow conditions generated by changing the shape of the wing are examined with wind tunnel experiments and numerical analyses to understand how the aerodynamic characteristics change with the wing shape of the stuffed bird using similar 3D printer models. The goal is to explore the optimum wing shapes, and flow control devices according to flight conditions and flight methods of delaying flow separation on the wing surface at the maximum lift and controlling stall. The results show that the aerodynamic characteristics of the bird models were confirmed through the experiment and the analysis. The lift coefficients of the bird models are relatively close in both data obtained in the wind tunnel test and the numerical analysis, except for the stall angles. It is considered that the stall angles caused errors due to angle of attack adjustment, surface roughness of the similar 3D printer models to the stuffed bird, and the method of data reduction of the experiments and the analyses. The stall angles of the bird models occur at angles of attack of 20 to 25 degrees, which is higher than those obtained with conventional wings. From the flow visualization patterns of the experiments and the analyses, the angle of attack where the stall occurred due to the separation of the wing clarified that the shape of the leading edge of the wing effectively prevented the separation on the bird wing, where a simple shaped slat delayed separation and controlled the stall moderately.

Y. Tanahashi, T. Kato, M. Amano, M. Sagata, N. Kishimoto, T. Ikeda
Investigation of Temperature Generalization for PCM-Based Cooling Plate Used for Remotely Located Electronics

In the present study, investigation of temperature generalization is performed to characterize the complex heat transfer process within a PCM -based cooling plate used for electronics remotely located in aircraft. Firstly, a simplified physical model is abstracted from the PCM-based cooling plate serving in actual applications where the cooling plate undergoes a complex heat transfer process with thermal conduction, thermal storage, thermal convection, and thermal radiation coupling with each other. Dimensionless numbers are defined to generalize the physical quantities associated with transient heat transfer with phase change. The comprehensive heat transfer process is then numerically simulated with the entropy-porous model to predict the melting process of PCM. Transient dimensionless temperature curves are presented in terms of dimensionless numbers including Rayleigh (Ra), Fourier (Fo), and Stefan (Ste) numbers. Transient heat fluxes of natural convection and thermal radiation during the whole heat transfer process are characterized by regression analysis. Finally, a modified product of dimensionless numbers is derived based on the results of regression analysis. It is found that the temperature could be generalized well with the modified product of dimensionless numbers. The results could provide guidelines for generalizing the performance of PCM-based cooling plates utilized in remotely located electronics and could be served as a quick design tool.

Bo Wu, Bin Wang, Liang Zhao, Junjie Han
Structural Design and Analysis of Large-Load Fire-Fighting Multi-rotor UAV

This paper designs a large-load multi-rotor fire-fighting UAV. Use CATIA three-dimensional drawing software for auxiliary design to control the size and weight of the drone within a reasonable range. After obtaining the three-dimensional model of the fire-fighting drone, HyperWorks finite element analysis software was used to perform finite element analysis on the airframe structure composed of carbon fiber composite materials and aluminum alloy. The results showed that the strength and rigidity of the airframe structure meet the design and use requirements.

Zhan-ke Li, Hai-yang Han, Yi Wu, Hai-bo Wei, Liang-yang Zhang
Research on Bearing Fault Recognition Based on PSO-MCKD and 1D-CNN

This paper proposes a bearing fault feature extraction and recognition method based on particle swarm optimization optimization Maximum Correlated Kurtosis Deconvolution (MCKD) and one-dimensional convolutional neural network to solve the non-stationarity of rolling bearing fault signals, Non-linear and complex characteristics, as well as the problems of noise interference and unclear fault characteristics in the process of fault identification. First, the multi-channel signals of the rolling bearing is analyzed, in order to select the signal containing the impact component as the fault feature. Next, the signal containing the fault feature is filtered through MCKD, where the best parameters of MCKD are obtained by improving the particle swarm algorithm to achieve the feature enhancement of the main signal. Finally, a one-dimensional convolutional neural network (One Dimensional Convolutional Neural Network, 1D-CNN) is used to model the characteristic signals under different damage conditions in order to obtain the fault recognition model of the rolling bearing.The experimental results show that the method can effectively extract the main characteristic signals of the faulty bearing, and realize the accurate identification of the bearing fault in the noisy environment.

Yinling Wang, Xianming Yin
Low-Speed Wind Tunnel Testing of a Passive Camber Morphing Airfoil Using a 3D-Printed Compliant Mechanism

Experimental investigations into the aerodynamics of the passive trailing-edge morphing airfoil model are presented. While most morphing airfoils are controlled using actuators, the passive morphing airfoil deforms due to the balance between the airfoil surface’s static pressure and the model flexibility. In our previous study, the designed morphing airfoil model had an interesting characteristic. As lift force increased, the camber of the airfoil increased and a higher lift was achieved. Meanwhile, the internal structure of the model was complex; thus, 24% thickness airfoil (NACA0024) had to be used, and maintaining a model-manufacturing accuracy was difficult. Considering adaptation to aircraft, it is necessary to study thinner airfoils in terms of aerodynamic performance and to manufacture models more accurately. Therefore, a new passive morphing airfoil model was manufactured based on NACA0012. The internal structure was simplified to realize thickness reduction and high model-manufacturing accuracy. The deformation and aerodynamic characteristics were measured through low-speed wind tunnel experiments. Two deformation characteristics, (1) decreasing of the effective angle-of-attack (AOA) and (2) increasing of the camber, were observed. In addition, the deformation gradually increased as the AOA increased. The maximum lift coefficient increased by 13% compared with the rigid airfoil. Further, the lift coefficient of the passive airfoil tailed off slower than the rigid airfoil at a stall angle.

Shoko Kai, Shogo Takazawa, Shuji Ochi, Taro Imamura, Tomohiro Yokozeki, Kenichi Rinoie
Surrogate Model-Based Parametric Structural Design of a Composite Tiltrotor Blade

Tiltrotor aircraft combine the vertical take-off and landing capability of a helicopter with the high cruising speed of a fixed-wing airplane. During flight, the aircraft switches between helicopter and airplane modes. Meanwhile, structural vibration and deformation are affected by changes in the aerodynamic loads, collective pitch, and rotation speed of blades. Therefore, it is crucial to consider dynamic characteristics in different modes. This paper presents a surrogate-based structural design framework of a tiltrotor blade to speed up calculations while maintaining accuracy. The shape generation method, rapid finite element method (FEM) model generation, and simplified boundary conditions are implemented to compute the blade modal frequencies, mass and inertia. A B-spline is used to construct the shape of the blade with the distribution of airfoils, twist angles, chord length, swept angles, and dihedral angles. These geometric features are realized by approximation, rotation, translation, and scaling of the control points. The spanwise distribution of the C-beam area and the width and length of the skins are selected to parameterize and establish the FEM model. The blind kriging model is applied to develop a surrogate model. The precision of this surrogate model is evaluated and compared to the kriging model and backpropagation neural network (BPNN) model based on the blade of an eight-ton-weight aircraft. A blade with a linear spanwise distribution of the C-beam area is designed using these three surrogate models trained by 13 samples. The results reveal that the error of the blind kriging model concerning the first four modal frequencies is less than 0.5% and lower than that of the other two models. A gimbal hub is modeled using a combination of a simply supported and a clamped boundary condition at the root. The two boundary conditions and two working conditions are unified into one working condition since a linear relationship is found. Moreover, the effects of the segmented linearly distributed cross-sectional area of the C-beam and the geometric parameters of the skins are investigated. The tip and root of the C-beam have the most significant influence on the low-order modal frequency. The parameters of the skins have opposite effects on the 1st and 2nd flap frequencies and the 1st torsion frequency. This framework provides a complete design process on the basis of geometric features and design parameters in the structure design, reducing the design variables. The use of the surrogate model and simplified working conditions reduce the consumption of time in the design. Conclusions about the tiltrotor blade have reference significance in the design stage.

Fan Sun, Chen Jiang, Yang Shen, Haowen Wang
Metadata
Title
The Proceedings of the 2021 Asia-Pacific International Symposium on Aerospace Technology (APISAT 2021), Volume 1
Editors
Sangchul Lee
Cheolheui Han
Jeong-Yeol Choi
Seungkeun Kim
Jeong Ho Kim
Copyright Year
2023
Publisher
Springer Nature Singapore
Electronic ISBN
978-981-19-2689-1
Print ISBN
978-981-19-2688-4
DOI
https://doi.org/10.1007/978-981-19-2689-1

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