Large Passenger Aircraft Research Articles

Numerical Prediction of Fatigue Life for Landing Gear Considering the Shock Absorber Travel

Due to the complexity of the landing gear’s (LG) structural integrity and its loads under various static or dynamic working conditions, the fatigue life assessment for LG is a highly challenging task. On the basis of the whole geometric model of a large passenger aircraft’s main landing gear (MLG), the quasi-static finite element model (FEM) of the whole MLG is established, and the high-cycle fatigue issue of the Main Fitting (MF) is studied by considering the variation in shock absorber travel (SAT). Firstly, the ground loads under actual fatigue conditions are equivalently converted into the forces acting on the center of the left and right axles of the MLG, and based on these spatial force decompositions, the magnitude and direction of the load for 12 different basic unit load cases (ULC) are obtained. That is, the stress of the MLG under actual fatigue conditions can be obtained by superimposing these ULCs. Then, considering that the SAT of the MLG varies under different fatigue conditions, and to reduce the number of finite element (FE) simulations, this article simplifies all the SAT experienced by the MLG into seven specific values, so as to establish seven quasi-static FEMs of the MLG with the specified stroke of the shock absorber. In this way, the fatigue stress of the MLG with any actual SAT can be obtained by interpolating the stress components of the seven FEMs. Only 84 FE simulations are needed to efficiently obtain the fatigue stress spectra from the ground load spectra. Finally, according to the material S-N curve and Miner’s damage accumulation criterion, evaluate the fatigue life of the Main Fitting. The results of the stress component interpolation and superposition method show that at least five different SATs of the whole MLG’s FEM are needed to effectively convert the fatigue loads into a stress spectrum. The fatigue life prediction results indicate that the minimum lifespan of the MF is 53164 landings, which means that the fatigue life meets the requirement design.

The impact and application of aircraft cooling systems on aircraft

Aircraft cooling systems were developed along with steam aeroplanes. In the state of high-speed operation, any temperature imbalance will cause an irreversible situation. In the evolution of generations of aircraft, aircraft cooling systems should also be upgraded and gradually returned to two categories, air cycle cooling system and evaporation cycle cooling system. The advantages and disadvantages of the two systems are obvious, and they complement each other, so in the modern background, there are also examples of some large passenger aircraft such as Airbus 350 combining the two cooling systems, which will also become the future development trend.

Analysis and Experimental Study on Vibration Isolation Performance of Full-scale Turbofan Engine Mounting

At present, most of the turbofan engine use the articulated link mount to connect the engine to the wing, and the analysis and experimental research of articulated link mount are important means in the development of engine mount. In this paper, the vibration isolation performance analysis of a certain turbofan engine mounting was carried out, and a full-scale turbofan engine mounting vibration isolation test platform was established based on the four-terminal parameter method, and the engine simulation device - eccentric rotation excitation tester was developed, which simulates the real engine vibration through rotor imbalance, and the vibration isolation efficiency of a certain type of engine mount at the high and low pressure rotor frequency was obtained. The results show that the simulation and test results are very close, the error is within 15% at the low-frequency, and the error is within 5% at the high-frequency. The research of this paper can provide test evaluation means for my country to independently carry out the design and improvement for large passenger aircraft engine installation devices, and provide test basis for its design and finalization.

Analysis and verification of vibration isolation performance of engine mount based on low-frequency approximation method

At present, most of the turbofan engine use the articulated link mount to connect the engine to the wing. The mechanical modeling, analysis and experimental research of articulated link mount are important means in the development of engine mount. In this paper, multibody dynamics method is used to establish the dynamic theory equation for a certain type of turbofan engine mount, and the low-frequency approximation method of the MNF file is proposed, which greatly improves the calculation speed on the basis of ensuring the calculation accuracy. In addition, a full-scale engine mounting system vibration isolation performance test platform was developed, and the vibration isolation efficiency of a certain type of engine mount at the high and low pressure rotor frequency was obtained. The results show that the simulation and test results agree well, the error is within 15% at the low-frequency, and 5% at the high-frequency. This work provides theoretical method and test evaluation means for the design and improvement of large passenger aircraft engine installation devices.

An Event-Driven Link-Level Simulator for the Validation of AFDX and Ethernet Avionics Networks

Aircraft are composed of many electronic systems: sensors, displays, navigation equipment, and communication elements. These elements require a reliable interconnection, which is a major challenge for communication networks since high reliability and predictability requirements must be verified for safe operation. In addition, their verification via hardware deployments is limited because these are costly and it is difficult to try different architectures and configurations, thus delaying design and development in this area. Therefore, verification at early stages in the design process is of great importance and must be supported with simulation. In this context, this work presents an event-driven link-level framework and simulator for the validation of avionics networks. The tool presented supports communication protocols commonly used in avionics, such as Avionics Full-Duplex Switched Ethernet (AFDX), as well as Ethernet, which is used with static routing. Also, the simulator provides accurate results by employing realistic models for various devices. The proposed platform was evaluated in the Clean Sky’s Disruptive Cockpit for Large Passenger Aircraft architecture scenario, showing the capabilities of the simulator. Verification speed is a key factor in its application, so the computational cost was analyzed, proving that the execution time is linearly dependent on the number of messages sent and that the increase in the number of nodes has few quadratic components.

An event-driven link-level simulator for validation of AFDX and Ethernet avionics networks

Aircraft are composed of many electronic systems: sensors, displays, navigation equipment and communication elements. These elements require a reliable interconnection, which is a major challenge for communication networks as high reliability and predictability requirements must be verified for safe operation. In addition, their verification via hardware deployments is limited because these are costly and make difficult to try different architectures and configurations, thus delaying the design and development in this area. Therefore, verification at early stages of the design process is of great importance that has to be supported via simulation. In this context, the present work presents an event-driven link level framework and simulator for the validation of avionics networks. The presented tool supports communication protocols such as Avionics Full-Duplex Switched Ethernet (AFDX), which is a common protocol in avionics, as well as Ethernet, which is used with static routing in such scenarios. The simulator also uses realistic element models to provide accurate results. The proposed platform is evaluated in Clean Sky’s Disruptive Cockpit for Large Passenger Aircraft architecture scenario. The speed of the verification is a key factor, so the computational cost is analyzed, proving that the execution time is linearly dependent on the number of messages sent.

A Follow-up Loading System during the Retraction and Extension Process of a Large Civil Aircraft

This article introduces a design scheme for a distributed load follow-up loading system for large passenger aircraft flaps. The follow-up loading system includes tape lever components, follow-up pulley and drive components, force control actuators, fixed pulley and cable components, and support structures. At the same time, a loading control system was also designed. The dynamic loading system can simulate the deformation of flaps under distributed aerodynamic loads during the entire retraction and extension process, greatly improving the fidelity of the test environment and comprehensively assessing the interference between flaps and surrounding structures under all operating conditions.

Analysis of the Application and Benefits of Aircraft Electric Wheel Systems during Taxi and Take-Off

An electric wheel hybrid power system is designed for driving a large single-aisle passenger aircraft during the take-off and ground taxi phases, which consists of an APU, an energy storage system, and a motor. In the taxi phase, the electric wheel hybrid power system works alone, and the turbofan engine does not work, reducing fuel consumption and pollution emissions. During the take-off rolling phase, the electric wheel hybrid power system and turbofan engine work together to reduce the thrust requirement of the turbofan engine. This article establishes an aircraft kinematic model, hybrid power system model, and a mechanical wheel model. The feasibility of the collaborative work of the electric wheels and the turbofan engines is verified by simulations. By utilizing the established hybrid system of electric motor wheels, the fuel consumption can be reduced, and the emissions of CO, HC, and NOX can also be diminished to varying degrees. The input of motor power leads to lower turbine inlet temperature, thereby enhancing the turbofan engine’s service life by approximately 4.3% and saving operational costs.

Adjoint-based robust optimization design of laminar flow airfoil under flight condition uncertainties

Due to the high sensitivity of laminar wings to uncertain factors, although there have been a lot of engineering application attempts, so far laminar flow technology has not been effectively applied to the wings of large passenger aircraft. To address this issue, we develop an adjoint-based robust optimization design framework. The framework mainly includes a Reynolds-Averaged Navier-Stokes solver coupled with the simplified eN method, coupled adjoint equations considering the transition, a gradient-enhanced polynomial chaos expansion, and a statistical moment gradient solver. The robust optimization designs of laminar airfoils subject to uncertainties in flight conditions for subsonic and transonic conditions are performed and compared with the results with deterministic optimization. The results show that the robust design can improve the ability of the laminar airfoil to resist the uncertain disturbance of flight conditions and reduce the mean and standard deviation by reasonably balancing the deterministic and uncertain performance. The mean and standard deviation of the drag coefficient can be reduced by 28%-36% and 20%-71%. Meanwhile, a longer and more stable laminar flow region can be obtained. In contrast, the deterministic design may lead to a large decrease in the robustness of performance. The successful results demonstrate the effectiveness of the established robust design method for laminar airfoils, and could be applied and extended for future laminar flow wings design. Although the present framework is verified by 2D optimization examples, it has the potential to be applied to quasi-3D and 3D laminar wings, as well as actual large passenger aircraft design.

Impact at aircraft level of elastic efficiency of a forward-swept tailplane

This paper evaluates the impact at aircraft level of the elastic efficiency of an advanced rear-end concept for a large passenger aircraft, exploiting a low-fidelity yet reliable aeroelastic approach. The innovative concept leverages a forward-swept horizontal tailplane to unlock a tail-fuselage connection such that a structural opening in the aircraft's rear-end is avoided. This installation allows for weight reduction in the structure, resulting in a positive impact on aircraft fuel burn. Moreover, a forward-swept tail has a different aerostructural behaviour that can be exploited to reduce its size with further weight and aerodynamic drag savings. In this respect, elastic efficiency is a crucial parameter for measuring the impact of this configuration at the aircraft design level. It takes into account both aerodynamic and structural characteristics, making it a comprehensive measure of effectiveness. Two different tail arrangements are being considered for an A320 neo-like aircraft: an innovative forward-swept design and a conventional layout that is equivalent. The results indicate that the forward-swept horizontal stabilizer has a higher elastic efficiency compared to the conventional tail arrangement. This could potentially lead to a reduction of the tailplane surface by approximately 2%, while still maintaining the same stability and control characteristics. This reduction in tail size unlocks the potential for fuel savings of approximately 0.5% on a mission profile of 3,400 nautical miles. Elastic efficiency is just one of the advantageous features of this innovative concept. By incorporating all the innovations proposed by the advanced rear-end concept, a weight reduction of up to 20% at the component level is expected. This could potentially result in fuel savings of approximately 2% for an aircraft similar to the A320neo, which has a mission range of 3,400 nautical miles.

Three-Dimensional Point Cloud Data Pre-Processing for the Multi-Source Information Fusion in Aircraft Assembly

Wing-body assembly is a key part of aircraft manufacturing, and during the process of wing assembly, the 3D point cloud data of the components are an important basis for attitude adjustment. The large amount of measured point cloud data and the obvious noise affect the quality and efficiency of the final assembly. To address this problem, research on the pre-processing method of the component point cloud data is carried out. Firstly, a feature-enhanced point cloud resampling method is proposed to preserve key features such as part contours in the resampling process. Then, a multi-scale point cloud data noise filtering method is proposed, which can effectively filter out the outliers. The experimental results show that the proposed method improves the speed and accuracy of the subsequent point cloud analysis effectively and is successfully applied to the assembly process of a large passenger aircraft, laying the foundation for high-quality assembly.

Research on Integration of Design and Manufacturing of High Thickness Honeycomb Sandwich Structure with Lager Curvature

Based on the characteristics of honeycomb sandwich structure with high thickness and large curvature, the idea of integration of design and manufacturing was proposed to solve the problem of mismatch among design, process and material. Facing the particularity of this kind of structure, the honeycomb deformation simulation model was established. And the composite building block verification method was simplified aiming at solving the problems of sliding and collapse of high thickness honeycomb. A variety of honeycomb stabilization methods were proposed. And the verification contents at the material level, element level, sub component level and component level were specified. Taking the landing gear door parts of large passenger aircraft as a case, the idea of design and manufacturing integration and the building block process verification method of honeycomb sandwich structure was practiced leading to the forward design of gear door parts. The honeycomb deformation simulation model was preliminarily established, which provided a reference for the forward design of similar structural parts.

Adjoint-based robust optimization design of laminar flow wing under flight condition uncertainties

It is an inherent uncertainty problem that the application of laminar flow technology to the wing of large passenger aircraft is affected by flight conditions. In order to seek a more robust natural laminar flow control effect, it is necessary to develop an effective optimization design method. Meanwhile, attention must be given to the impact of crossflow (CF) instability brought on by the sweep angle. This paper constructs a robust optimization design framework based on discrete adjoint methods and non-intrusive polynomial chaos. Transition prediction is implemented by coupled Reynolds-Averaged Navier-Stokes (RANS) and simplified eN method, which can consider both Tollmien-Schlichting (TS) wave and crossflow vortex instability. We have performed gradient enhancement processing on the general Polynomial Chaos Expansion (PCE), which is advantageous to reduce the computational cost of single uncertainty propagation. This processing takes advantage of the gradient information obtained by solving the coupled adjoint equations considering transition. The statistical moment gradient solution used for the robust optimization design also uses the derivatives of coupled adjoint equations. The framework is applied to the robust design of a 25° swept wing with infinite span in transonic flow. The uncertainty quantification and sensitivity analysis on the baseline wing shows that the uncertainty quantification method in this paper has high accuracy, and qualitatively reveals the factors that dominate in different flow field regions. By the robust optimization design, the mean and standard deviation of the drag coefficient can be reduced by 29% and 45%, respectively, and compared with the deterministic optimization design results, there is less possibility of forming shock waves under flight condition uncertainties. Robust optimization results illustrate the trade-off between the transition delay and the wave drag reduction.

Influence of Single External Store on Geometrical Nonlinear Flutter of High-aspect-ratio Wings

High-altitude and long-endurance UAVs, distributed electric propulsion aircraft and large passenger aircraft often use high-aspect-ratio wings, which have obvious geometric nonlinear large deformation effect, and are usually arranged with a certain number of external stores, which have an impact on their geometric nonlinear aeroelastic characteristics. Based on the “quasi-linear” hypothesis, the updated Lagrange incremental finite element method and doublet-lattice method are used to perform the geometrically nonlinear static aeroelastic analysis, and the mass, stiffness and aerodynamic matrices of large deformation are updated, and the geometrically nonlinear flutter characteristics of high-aspect-ratio wings are calculated. The influence of the position and mass of the external store on geometrical nonlinear flutter characteristics of high-aspect-ratio wing structure is investigated. Results show that the change of the position and mass of the external store causes the relative position between the elastic axis and the center of gravity of the wing, and the wing’s large deformation equilibrium state to change, resulting in differences in the stiffness characteristics of the wing and the natural vibration frequency of the equilibrium state, thus affecting the wing flutter characteristics. When a single external store is installed near the wing root, changing the mass and direction of the store has no significant effect on the geometric nonlinear flutter characteristics of the wing. The improvement of wing flutter speed is most obvious when the hanging point is moved to the leading edge of the wing near 0.65 times the wingspan.

Subjective evaluation of the acoustic annoyance in a large passenger aircraft cabin

People demand better acoustic comfort for higher travelling quality and security in aircraft. It is necessary to evaluate and predict the subjective annoyance caused by the noise in aircraft cabins. This study investigates the noise-induced annoyance in a large passenger aircraft. We recorded the noise at 21 positions in the aircraft cabin without passengers and selected 21 stimuli during the cruising. Each stimulus has a duration of 5 seconds and a sound pressure level in the range of 72-81dB(A). The psychoacoustic parameters such as loudness, sharpness, roughness and articulation index were also calculated. Twenty-four subjects evaluated the subjective annoyance of the stimuli by the absolute magnitude estimation method. Results showed the noise annoyance at the middle section of the cabin is significantly higher than that at the front or rear section. The principal component analysis and correlation analysis found that annoyance is mainly affected by loudness, sharpness and roughness and dominated by loudness. We proposed a multiple linear regression model between the subjective magnitudes of annoyance and the psychoacoustic parameters to evaluate and predict the noise annoyance in the aircraft cabin well.

Influence of geometric parameters of leading edge horn-ice on stall characteristics of airfoil

The irregular characteristics and geometric randomness of horn-type ice are significant while the behaviors of separation are complex. The mechanism and factor leading to the essential change of stall performance are still not clear. According to actual requirements of airworthiness certification of a large passenger aircraft, a family of horn-type ice shapes with different parameters are constructed based on typical supercritical airfoil and icing environment. By combining with wind tunnel test and numerical simulation methods, the parameter sensitivity of airfoil stall characteristics to the change of ice height and angle is systematically analyzed and the essential influence mechanism of the ice shape parameters leading to the change of separation bubble development under stall process is summarized, which provides a theoretical basis for ice airworthiness certification of large passenger aircraft.

Conceptual design, AC loss calculation, and optimization of an airborne fully high temperature superconducting generator

With the trend of multi-electrification and hybrid-electrification of aircraft, the airborne high temperature superconducting (HTS) generator is considered as the key technology of the new generation of aviation power generation system and has received wide attention. The high critical current and low loss of superconductors can significantly improve the output power, power density, efficiency and other performance of generators. In this work, conceptual design of a fully HTS generator for aircraft is carried out to address key issues in terms of AC loss distribution, which serves as the ground of cryogenic design in the future. A 10 MW fully superconducting generator is designed to be used in a large passenger aircraft (with more than 150 seats and a maximum takeoff weight of about 100 tons). Based on the finite element method (FEM) with the so-called T-A formulation, the electromagnetic characteristics of the generator are analyzed, and AC losses of the stator and rotor are calculated. To calculate the rotor loss in realistic dynamic-field environment, a technique named “rotation-domain exchange” is developed and implemented. By using such technique, accurate calculation of AC loss of the rotor in realistic dynamic electromagnetic environment is achieved with no need of simplified boundary condition as previously reported work. Based on the model, the stator coil structure is optimized to reduce the AC loss and the rotor coil structure is optimized to reduce the HTS tape usage. Results show that the optimized stator AC loss is 15.645 kW, and the rotor AC loss is 6.268 kW, indicating that the rotor AC loss of a fully superconducting generator with narrow air gap and large number of coil turns is not neglectable as previous work assumed. Thus, precise calculation and dedicated optimization of the rotor loss is critical since cooling of the rotor is much more challenging than the stator as the former remains in motion.

Recent developments of common numerical methods and common experimental means within the framework of the large passenger aircraft program

For the exploration and validation of the integration of efficient propulsion concepts and new technologies for future large passenger aircraft, it is necessary to establish suitable design, evaluation and measurement tools. Therefore, as one major activity of the Large Passenger Aircraft (LPA) Platform 1 of the Clean Sky 2 initiative, the so-called Cross-Capability Demonstrator (XDC), has been set up to develop and demonstrate powerful numerical and experimental methods for aerodynamic, aeroacoustics, and aeroelastic simulation and measurement tasks. The paper will give an overview of the activities to be performed within the XDC and present some of the latest achievements related to this demonstrator.

Pull control of material supply for low-volume assembly lines: a reorder point method for aerospace manufacturing

In many medium- and high-volume assembly environments, e.g., automobile manufacturing, pull control of material supply facilitates resilient and lean manufacturing processes through consumption-based replenishment in aggregated quantities of one or more complete material requirements of an assembly cycle. However, in material-intensive low-volume environments such as aerospace manufacturing, this aggregation at the assembly level would lead to excessive supply; thus, material must be replenished gradually over an assembly cycle. The relevant research is limited, and grouping-based principles proposed for such environments can lead to suboptimal task aggregation and undesirable preemption. This paper presents a reorder-point-based pull control method that enables highly precise replenishment. It was developed and evaluated through a simulation study of a wing assembly line for large passenger aircraft. The simulation results suggest that the proposed method is highly effective for the complex assembly environment investigated and can reduce work in process even further than existing methods.

Research and Verification on the Solution of Bleed Air Temperature Deviation of Civil Aircraft Pneumatic System

The pneumatic system of civil aircraft is equipped with a precooler (PCE) to cool the high-temperature bleed air from the engine to meet the design requirement of a downstream user system. The PCE is usually a fork-flow heat exchanger. This heat exchange structure will make the outlet temperature difference especially large, which shall lead to the deviation of the inlet air temperature of the downstream wing anti-icing system and air conditioning system and cause the temperature of bleed air supplied to the wing anti-ice system unstable. During a flight test of a certain large passenger aircraft, it was found that the bleed temperature at the PCE outlet was lower than the piccolo inlet temperature of the wing anti-ice system when the wing anti-ice system was operated with two engine bleed systems on. This paper creatively proposes design ideas for the three kinds of mixers and determines the solution of adding a mixer downstream of the PCE to solve the temperature deviation problem. The mixer is installed at the hot side outlet of the PCE and used to mix the temperature field at the PCE outlet so that the air supplied to the downstream customer systems meets the design temperature requirements. According to the simulation calculation and laboratory test verification results, the design scheme of a certain large passenger aircraft is selected, which has successfully solved the problem of temperature stratification at the outlet of the PCE and supported the follow-up flight test.