Flight Loads Research Articles

Experimental and Numerical Modal Analysis of a Composite Rocket Structure

Finite Element Analysis (FEA) is a powerful tool that can aid in the engineering design process to reduce cost and time. However, it is best used in conjunction with experimental data, through which its numerical results can be verified. This paper presents the experimental and numerical modal analyses of an experimental rocket aerostructure to verify the accuracy of the numerical models. This aerostructure has been through flight loads and a recovery. The first numerical results for the rocket showed a 96% difference with the experimental ones. Subsequently, three mass refinements were made to create calibrated FEM models whose results differed from the experimental ones by 19% to 8%. Additionally, as expected, the FEM results tended to overestimate the stiffness of structures. The numerical simulations for all components were performed through ANSYS software, and the experiments were conducted using the hammer tap test with laser vibrometers as sensors.

Overview of criteria to estimate aerodynamic limits of the flight envelope of a transonic aircraft based on RANS simulations

The aerodynamic limits of the flight envelope of an aircraft play an important role for flight characteristics and loads, so they must be estimated already in the early stages of aircraft design. However, without wind tunnel tests or costly scale-resolving simulations, it is difficult to make an accurate prediction. From a computational perspective, this problem can be split into two parts: (1) generating a valid flow solution for edge-of-the-envelope conditions and (2) extracting the exact lift limit from these computations. Not only the first point can be very challenging, but also the second point. At low Mach numbers, the lift limit may be detected relatively easily, but already at moderate Mach numbers, maximum lift and the lift limit do not necessarily coincide any more, as the actual achievable lift may be limited by the onset of transonic buffet before maximum lift is achieved. Since buffet itself can only be predicted accurately with high effort using time-resolving CFD methods, indirect criteria are required for estimating the lift limit based on steady RANS simulations. This is the focus of the present paper. For this purpose, various criteria proposed in the literature are presented, applied to a test case and comparatively evaluated. For the test case, steady RANS simulations of the generic wing-fuselage configuration LEISA at different Mach numbers and angles of attack are carried out. Finally, the advantages and disadvantages of the different criteria are identified and discussed.

Statistical Algorithms with Phase-Sensitive Detection for the Number of Hovers and S Turns in Aircraft Flights

Compared to maximum state parameters, such as maximum altitude and Mach, the number of hovers and S turns can be used as process parameters representing the complexity of military aircraft maneuvers when classifying big flight mission data to compile flight load spectra for structures. This study developed intelligent statistical algorithms based on yaw angle data from flight parameters such as the number of hovers and S turns. Using the median-crossing de-redundant function of Phase-Sensitive Detection (PSD) and analyzing the characteristics of 360° hovering flight parameters, a statistical algorithm for the number of hovers during a flight profile is presented. Using the split-half function of PSD, a triangle layering algorithm based on the yaw angle signal was developed to count the number of S turns during a flight profile, where the signal of each sublayer is segmented into median-crossing intervals to eliminate the redundant median-crossing marks from the previous layer. Compared with artificial means, the statistical results of the flight example showed that the developed intelligent algorithms are effective.

A Comparative Regression Analysis between Principal Component and Partial Least Squares Methods for Flight Load Calculation

This study investigates and compares various multivariate regression methods, including principal component regression (PCR) and partial least squares regression (PLSR), for flight load analysis and demonstrates their high learning efficiency and strong generalization capabilities, making them highly suitable for this purpose. The flight load data of a civil aircraft use altitude, Mach number and load factors as input parameters, which are used as sample data to establish regression models for predicting wing loads under different flight conditions. The accuracy of all regressions are confirmed through evaluation, with PLSR being the most efficient. In the comparison of computational times, it was found that the computational efficiency of regression methods was significantly superior to traditional panel methods. The flight load calculation shows that PCR and PLSR can significantly improve analysis efficiency and provide new insights into efficient flight load analysis.

Finite element analysis of the underslung carrier rocket effect on stress and strain distribution in a structure of the MiG-29 aircraft

The paper presents selected aspect of the analytical-conceptual research project entitled: Airborne-Rocket Launch System for Delivering Satellite Payloads into Low Earth Orbit – Feasibility Study. The aim of the project is to conduct series of aerodynamic, strength, and aeroelastic simulations of Polish aging combat aircraft (Mig-29 and Su-22) to explore applicability of both fighters as an airborne platform for carrying out carrier rockets with a detachable satellite payload. This work presents exemplary analyses in the area of airframe loads and structural strength and deformability of MiG-29 for predicted operational variants with carrier rockets put to the hardpoints under fuselage. The numerical simulations were conducted for a structural discrete model of the aircraft prepared for finite element analysis in MSC Software. Model development involved such aspects as precise discretization of geometric model, declaration of material constants, identification of structural properties, introduction of suitable merging connections for included airframe assemblies, and final validation of model mass and stiffness. The model was analyzed in MSC Nastran software with application of linear “Statics” solution. External flight loads introduced into the model were calculated for specific points of the flight envelope—the highest values of load factor were taken into consideration ( n=9). The counterpart aerodynamic force distribution in a form of a set of equivalent lumped forces was calculated. Then, the aerodynamic and weight loads were added to the model as a set of forces applied to specific structural nodes. Additionally, weights of applicable carrier rockets were taken into account as concentrated forces applied to under-fuselage suspension hardpoints. Calculations were performed for several rockets of mass values between 250 and 1200 kg. Parametric dependencies were investigated as an effect of missile mass and size on stress and strain distribution over the whole structure. The areas of stress cumulations were identified. On the basis of static structural deformation, maximum wing tip displacements were assessed.

Load calibration experiment analysis for scaled model of large aspect ratio wing

Load calibration experiment is the key work for flight load verification and flight test of civil aircraft. The efficiency of the flight load verification is meaningfully determined by the reliability of load calibration experiment. This paper aims to theoretically and experimentally study the load calibration of civil aircraft wing with large aspect ratio and elastic beam based on the scaled model, which includes the design of fixed device for wing loading, the data processing, the basic hypothesis test, the selection of regression variables and optimisation of regression equation, etc. Firstly, the systematic error is eliminated based on the sample difference method. Then, on the basis of the multiple linear regression method, the multi-collinearity among variables is weakened by using the multi-collinearity verification technology. Subsequently, the abnormal sample points are eliminated by the means of cyclic residual analysis. Finally, the optimized variables are screened with the method of the stepwise linear regression, and the load equations for the civil aircraft wing with large aspect ratio can be established. It is demonstrated that the calculated results verifies well with the engineering requirements by comparing with experimental results, and the error is smaller than 5%.

Flight Load Calculation Using Neural Network Residual Kriging

Flight load calculation, an important step in aircraft design and optimization, typically involves millions of computations and requires significant computing resources and time. Improving the efficiency of flight load calculations while maintaining accuracy is therefore of great significance for shortening research and development cycles. This study investigated and compared multiple algorithms, including the neural network model, the Kriging surrogate model, and the neural network residual Kriging (NNRK) model, for flight load analysis. The accuracies of all models were confirmed through evaluation, with NNRK being the most efficient, making it highly suitable for flight load analysis. The flight load data of a civil aircraft, including the total weight, the center of gravity, the pitch moment of inertia, the altitude, the Mach number, the airspeed, the velocity pressure, the pitch rate, the load factor, and the angle of attack as input parameters, were used as sample data to establish models, for predicting wing loads under different flight conditions. The accuracies of all regressions were confirmed through evaluation, with NNRK being the most efficient. The flight load calculation shows that NNRK can significantly improve analysis efficiency and provide new insights into efficient and comprehensive flight load analysis.

Fatigue failure analysis on the 2nd bulkhead beam of the Portuguese Air Force Epsilon TB-30 aircraft

Military aircraft perform demanding maneuvers with a wide range of load factor levels, allowing the growth of fatigue cracks that shall be duly monitored to prevent unexpected failures. In order to monitor the operation of the Portuguese Air Force Epsilon TB-30 fleet that performs basic and elementary piloting training, two systems were installed in these aircraft, allowing the recording of load factor data in the aircraft center of gravity and strain data in two critical locations.This study aims to evaluate and compare the growth of fatigue cracks on the 2nd bulkhead beam, one of the aircraft critical locations, through experimental tests and computational simulations. The methodology used in this work comprised the application of real flight load sequences to a 2024-T351 aluminum specimen that is representative of this critical location.Main results from these fatigue crack growth tests and simulations showed that NASGRO model is the one that comes closest to experimental data with maximum deviations of approximately 8%, presenting a deviation of 18% between the representative and real structures at the final rupture of the specimen. Additionally, experimental results provided knowledge about the component behavior regarding fatigue of materials, namely the evolution of the local strain as a function of the crack length.This study can be used by the Portuguese Air Force engineers to adjust the aircraft maintenance program according to the actual operation regime and consider a possible life extension of this fleet.

Automated sizing chain of aircraft structures: A Vertical Tail Plane use case

Current developments in preliminary design emphasise the potential benefit of addressing detailed sizing cases in the early design stages. This implies a sound understanding of the loads that the structure must withstand and for which it must demonstrate compliance. In aircraft structural practices, stringent safety requirements demand robust designs. It is understood that the foremost limitation and very challenging aspect, in the aviation sector, consists in complying with those requirements while targeting a contained mass. In this respect, the presented work aims to propose an automated implementation of the dimensioning chain by foregrounding a specific subsystem application, id est the Vertical Tail Plane. The conditions corresponding to a set of critical flight loads are identified and used for the sizing process. The use case intends to standardise the process, narrowing down the large number of unknowns on the overall loosely defined design method.

Analysis of the Functional Capabilities of Technical tools for Measuring Aircraft Mechanical Parameters

The purpose of research is to analyze the functionality of classical and advanced sensors for controlling the mechanical parameters of aircraft using the example of strain and vibration sensors to identify their current areas of development.Methods. Research methods are based on the concepts of the theory of sensory systems, the theory of diagnostics and forecasting of the technical condition of aircraft. Methods of multi-criteria analysis, parametric and structural synthesis are used. The principles of operation, as well as the functionality of the main classical sensors used in aircraft to control the parameters of deformation and vibration, are analyzed. A critical assessment of the possibilities of using the analyzed sensors for implementation in various tasks of aviation diagnostics of mechanical parameters has been made.Results. It has been established that the impact of flight loads on the airframe and critical components of aircraft is accompanied by the appearance of hidden deformations in the form of mechanical stresses, which are divided into two components: normal and tangential. Analytical dependencies are obtained for calculating the above quantities using fiber-optic sensors with distributed Bragg cells that convert the change in their own linear dimensions into a change in the reflected wavelength. A necessary condition for obtaining correct measurement results is the temperature compensation of the cells, which makes it possible to localize the places of deformations with an accuracy up to the location of a particular cell. The practical results of using alternative sensors for detecting hidden deformations (cracks) based on radio frequency identification methods in various frequency ranges are presented.Conclusion. The development of the method for diagnosing stress-strain states of aircraft complex units is the use of frequency-Doppler fiber-optic sensors with a high signal-to-noise ratio and a spherical radiation pattern, which will allow developing technical means for monitoring the dynamics of internal deformations of controlled units in real time. As promising areas of research in the field of creating new sensors with new physical properties, fiber-optic Bragg sensors with an inclined grating should be considered.

Algorithm for Calculating the Flight Time of an Unmanned Aerial Vehicle for Aerial Photography

The paper shows the possibility of using unmanned aerial vehicles with additional outboard equipment for aerial photography. It is noted that some parameters of the drone and additional suspension equipment are not taken into account by the software when calculating the drone flight time for one battery charge. (Research purpose) To develop an algorithm for calculating the flight time of an unmanned aerial vehicle with installed outboard equipment. (Materials and methods) The technical characteristics of DJI Phantom 4 pro and DJI Matrice 200v2 were studied, as well as the specifications of Parrot Sequoia, MicaSense Altum multispectral cameras mounted on the drone. The existing research results are used to calculate the flight mission time depending on the route length and the battery capacity. (Results and discussion) It is found that the maximum flight time of a drone with additional outboard equipment is reduced due to the additional equipment mass, increased power consumption, the preflight preparation time, the need to return to the take-off point and the necessity to preserve the battery life and save the flight load. The maximum flight time calculated for DJI Phantom 4 pro and DJI Matrice 200v2 with multispectral cameras Parrot Sequoia, MicaSense Altum is 8 minutes and 18 minutes, respectively, with a minimum flight load. A method for calculating the number of batteries for aerial photography with additional outboard equipment is determined. (Conclusions) An algorithm for calculating the flight time of an unmanned aerial vehicle with additional outboard equipment is developed, the parameters ignored by the standard software in the flight time calculation are taken into account.

A Flight Parameter-Based Aircraft Structural Load Monitoring Method Using a Genetic Algorithm Enhanced Extreme Learning Machine

High-precision operational flight loads are essential for monitoring fatigue of individual aircraft and are usually determined by flight parameters. To tackle the nonlinear relationship between flight loads and flight parameters for more accurate prediction of flight loads, artificial neural networks have been widely studied. However, there are still two major problems, namely the training strategy and sensitivity analysis of the flight parameters. For the first problem, the gradient descent method is usually used, which is time-consuming and can easily converge to a local solution. To solve this problem, an extreme learning machine is proposed to determine the weights based on a Moore–Penrose generalized inverse. Moreover, a genetic algorithm method is proposed to optimize the weights between the input and hidden layers. For the second problem, a mean impact value (MIV) method is proposed to measure the sensitivity of the flight parameters, and the neuron number in the hidden layer is also optimized. Finally, based on the measured dataset of an aircraft, the proposed flight load prediction method is verified to be effective and efficient. In addition, a comparison is made with some well-known neural networks to demonstrate the advantages of the proposed method.

Static aeroelasticity of the propulsion system of ion propulsion unmanned aerial vehicles

“Ionic wind” generators are used as the main propulsion system in ion propulsion unmanned aerial vehicles (UAVs). Owing to the large size and poor stiffness of the electrode array in the propulsion system, the electrode array is prone to deformation under the flight load. In this work, the thrust characteristics and static aeroelastic properties of “ionic wind” propulsion systems were analyzed in detail. The simulation model for an “ionic wind” propulsion system was established by coupling a two-dimensional gas discharge model with a gas dynamics model. The influences of electrode voltage, spacing, size, and shape on the performance of the propulsion system were investigated. The fluid-solid interaction method was used to solve static aeroelastic characteristics under deformation. The aerodynamic and thrust performances of the elastic state and the rigid state were compared. It was found that the operating voltage, the distance between two electrodes, and the emitter radius had greater impacts on the thrust of the propulsion system. The propulsion system had a small contribution to the lift but a large contribution to the drag. In the elastic state, the lift coefficient accounted for 12.2%, and the drag coefficient accounted for 25.8%. Under the action of the downwash airflow from the wing, the propulsion system formed an upward moment around the center of mass, which contributed greatly to the pitching moment derivative of the whole aircraft. In the elastic state, the pitching moment derivative accounted for 29.7%. After elastic deformation, the thrust action point moved upward by 28.7 mm. Hence, the no lift pitching moment is reduced by 0.104 N·m, and the pitching moment coefficient is reduced by 0.014, causing a great impact on the longitudinal trimming of the whole aircraft.

Validation of the structural integrity of an aircraft external fuel tank through a structural test under flight-load conditions

An external fuel tank, one of the main components of an aircraft, is a component that extends the range of the aircraft and that should be separable from the aircraft in emergency situations. During flight, if structural damage occurs in an external fuel tank, the survivability of the aircraft may be catastrophically affected. Therefore, the structural integrity of such external fuel tanks must be proven under flight-load conditions applied to the external fuel tank during flight. This study presents the results of a structural test conducted to verify the structural integrity of an external fuel tank for flight-load conditions. In the main text, the process of converting the six-component flight load into a test load is described, and a whiffle tree application plan is presented for efficient testing. Then, the reliability of the test load was confirmed by comparing it with the required flight load. As a result of the structural test, the test loads were stably controlled. It was also confirmed that the internal pressure of the test specimen measured in real time during the test was maintained well within the required pressure range and that the behavior of the EFT was suitably predicted through a numerical analysis. After the structural test, it was found through an inspection that the test specimen did not undergo serious structural damage under the required flight-load condition. In conclusion, through structural testing, it was validated that the newly developed external fuel tank had sufficient structural integrity under all required flight-load conditions.

UltraFLoads: A Simulation Suite and Framework for High-Fidelity Flight Loads

With the progress of high-performance computing, computationally expensive high-fidelity methods can be applied early in the design process of an aircraft. This enables Computational Fluid Dynamics (CFD) for the assesment of flight performance, handling qualities, flight loads due to manoeuver or gusts, and stability such as flutter. Those aeroealstic analyses require coupling of disciplines to adequately address the interplay of flexibility, aerodynamic forces and inertia forces. Previous work at the German Aerospace Center have demonstrated the proof of concept for those multidisciplinary analyses. Now, UltraFLoads is a simulation suite in the FlowSimulator ecosystem to offer standardized, multidisciplinary scenarios. UltraFLoads can be used as a tool, as a library and as a framework. This work describes the architecture, the integration layers, which scenarios are available and which equations are used. Three different applications are presented. The first case involves a steady fluid–structure-coupled simulation with geometrically nonlinear deformation. The second case compares the deformation of a simulated elastic, free-flying aircraft with actual flight test data. The last application demonstrates the flight dynamic capabilities with a bank to bank maneuver.

Estimating the Economic Viability of Advanced Air Mobility Use Cases: Towards the Slope of Enlightenment

While different vehicle configurations enter the AAM market, airlines declare different ticket fares for their operations. This research investigates the operating cost of an airline and the economic viability with the announced fare per km rates. For this purpose, three use cases in the metropolitan area of Hamburg showcase representative applications of an AAM system, whereby a flight trajectory model calculates a flight time in each case. The direct operating cost are investigated for each use case individually and are sub-classified in five categories: fee, crew, maintenance, fuel and capital costs. Here, each use case has its own cost characteristics, in which different cost elements dominate. Additionally, a sensitivity analysis shows the effect of a variation of the flight cycles and load factor, that influences the costs as well as the airline business itself. Based on the occurring cost, a profit margin per available seat kilometer lead to a necessary fare per km, that an airline has to charge.

Application of electrolaminates for the development of biomimetic morphing unmanned aerial vehicles

Morphing aircraft offer the promise of performance that can be optimized for a range of flight profiles. But such capability places challenging demands on the morphing system that cannot easily be met with conventional actuators or smart materials. Electrolaminates use electrostatic forces controlled by the application of a voltage (typically high-voltage but extremely low-current) to controllably bond or unbond layers of a laminated composite structure together. Such on-demand bonding can be used to effectively change the stiffness of the structure or control elongation or twist of the structure. This technology has multiple benefits compared to other smart material or conventional actuator-driven approaches; it is thin and light, has very low energy requirements, and offers rapid response capabilities controlled by simple electronics. These capabilities can enable bio-inspired morphing with large numbers of degrees of freedom and high spatial resolution. We designed, fabricated, and tested a fully functional morphing-wing unmanned aerial vehicle (UAV) with telescoping wings and a splaying, bird-like tail enabled by electrolaminates. Key features of the current UAV configuration include: a variable wingspan of 1.2–2.4 m with a corresponding change in the wing area of nearly a factor of four; no vertical tail; and a horizontal tail area variable by a factor of three. The ability to asymmetrically control the tail area allows for independent pitch and body circulation control. The tail area is changed by a separate electrolaminate clamping mechanism. The variable-area tail uses ‘feathers’ that can be overlapped or splayed as needed. The practical shape-changing is enabled by electrolaminate materials that can rapidly lock the orientation of the feathers. The electrolaminate clutches support more than 5 nm of torque and are sufficient to resist expected flight loads. We also designed, fabricated, and tested other morphing-wing designs enabled by electrolaminate technology to create a wing with smoothly sliding skin that is capable of changing chord and camber with a single linear actuator. Our results suggest that electrolaminates can practically enable bio-inspired, small (Group 1 and 2) morphing UAVs.

Hypersonic shock impingement studies on a flat plate: flow separation of laminar boundary layers

Interactions between shock waves and boundary layers produce flow separations and augmented pressure/thermal loads in hypersonic flight. This study provides details of Mach 7 impinging-shock-flat-plate experiments conducted in the T4 Stalker Tube. Measurements were taken at flow conditions of Mach 7.0 (2.44 MJ kg $^{-1}$ ) and Mach 7.7 (2.88 MJ kg $^{-1}$ ) flight enthalpies with a range of freestream unit Reynolds numbers from $1.43 \times 10^{6}$ m $^{-1}$ to $5.01 \times 10^{6}$ m $^{-1}$ . A shock generator at $12^{\circ }$ or $16^{\circ }$ to the freestream created an oblique shock which impinged on a boundary layer over a flat plate to induce flow separation. The flow field was examined using simultaneous measurements of wall static pressure, heat transfer and schlieren visualisation. Measured heat transfer along the flat plate without the shock impingement indicated that the boundary layer remained laminar for all flow conditions. The shock impingement flow field was successfully established within the facility test duration. The onset of separation was observed by a rise in wall pressure and a decrease in heat transfer at the location corresponding to the stem of the separation shock. Downstream of this initial rise, an increased pressure and higher heating loads were observed. The heat-transfer levels also indicated an immediate boundary layer transition due to the shock impingement. The separation data of the present work showed good agreement with our previous work on shock impingement on heated walls (Chang et al., J. Fluid Mech., vol. 908, 2021, pp. 1–13). A comparison with the previous scaling indicated that the separation also relates to the pressure ratio and the wall temperature parameter.

Results of measurements carried out for the purpose of determining the center of gravity of a convertiplane-type UAV

Currently, due to the increasing accuracy required to determine the center of mass and gravity, the development of methods and technologies in this direction remains a high priority. In the research work, the characteristics of the available methods and tools for determining the flight load and centering, as well as the analysis of the research work on determining the center of mass of the aircraft, calculating the weight and centering cost were carried out. When designing and preparing aircraft, in order to increase the efficiency and safety of their flights, it is required to correctly determine their loading, centering and center of gravity. The methodology of using the "balance-mass" method, which is widely used in determining the center of gravity of aircraft, was developed in our research conducted for this purpose - in the example of a convertiplane-type UAV. It is known that it is possible to determine the coordinates of the center of gravity of gliders beforehand based on survey books. This data only corresponds to the state of the glider without load and not on the final delivery. The article also describes the essence and capabilities of the developed method for determining the center of gravity of the aircraft with higher accuracy. The results of theoretical and experimental studies on the creation of a system for determining the weight and center of gravity of the UAV in practice are shown. The proposed system construction model is presented.

A Novel Reverse Hinge Spoiler for Flight Loads Control

This paper presents the reverse hinge spoiler, a novel spoiler concept, for flight load control. The reverse hinge spoiler is a control surface mounted on the upper surface of the wing. Unlike conventional hinged spoilers that are hinged at their front and rotate forward toward the leading edge of the wing, the proposed spoiler concept is hinged at its rear and rotates backward toward the trailing edge of the wing. The aerodynamic performance of the proposed spoiler is compared and contrasted with that of a conventional hinged spoiler for different flight conditions and hinge locations using the two-dimensional Reynolds-Averaged Navier–Stokes (RANS) with the k-omega SST turbulence model-based computational fluid dynamic solver. The results show that the proposed spoiler results in a larger increase in drag and a sharper reduction in the lift for a wide range of spoiler angles and flight conditions. Reversing the spoiler is found to cause a higher adverse pressure gradient in front of the spoiler compared to a conventional spoiler, as it ‘traps’ more flow, thereby increasing drag and reducing lift.