Rear Fuselage Research Articles

Transition Flight Control and Simulation of a Novel Tail-Sitter UAV With Varying Fuselage Shape

This article studies the model, control and simulation on transition flight of a novel tail-sitter with vertical takeoff and landing capability. The proposed tail-sitter adopts an innovative varying fuselage shape design to break the technical bottleneck in the balance of efficient horizontal flight and agile vertical flight. Overset grids and computational fluid dynamic methods are used to explore the vehicle’s prestall aerodynamics, which are relative to not only the angle of attack and sideslip angle but also to the varying angle between two rear fuselage’s parts. High angle of attack aerodynamics based on the improved Viterna and Corrigan stall model is also established for this novel tail-sitter. Meanwhile, an accurate model of the propeller is tested in a wind tunnel. Combining the forces and moments generated by propellers, aerodynamics and gravity, a 6DoF nonlinear time-varying dynamic model is built. A robust controller based on incremental dynamic inversion method is designed for this tail-sitter, which is good at dealing with uncertainties and external forces. Nominal and model mismatch conditions are simulated to verify the controller’s performance. Different varying strategies for the mechanism are analyzed during the transition flight. Simulation results show that this novel tail-sitter can transform between vertical and horizontal flight mode easily and the varying strategy related to pitch angle is a prior choice for transition flight.

FUNDAMENTAL LESSONS OF THE FIRST HUMAN VERTICAL ROCKET FLIGHT

ABSTRACT This paper examines the historic but tragic first human vertical rocket flight which took place in south-western Germany on 1 March 1945. The primary lesson learned from the flight was that, as a result of the combination of psychological and physiological stresses, a human pilot could not be expected to fly a vertically launched rocket manually. An autopilot would be essential for the guidance of the Natter rocket interceptor up to its operational altitude. No further human vertical rocket flights took place until 1961 when Major Yuri Gagarin was launched into orbit. In early April 1945 a fully operational Natter flew successfully into the lower stratosphere under the control of a three-axis autopilot and crewed with a dummy pilot. Both dummy pilot and rear fuselage were recovered successfully under separate parachutes. In less than a year the engineers and scientists in collaboration with aviation physicians and physiologists at research institutions across Germany had laid down the basic principles which still apply to human rocket flight today.

Influence of interference between the fuselage and the propeller ring on the maximum thrust of the air pusher propeller

The geometrical parameters of ring profiles were presented and the estimation of its influence on the peak values of propeller thrust was described. The results of computational studies using software based on Reynolds-Averaged Navier-Stokes (RANS) equations to investigate the influence of the rear fuselage and the shape of the ring profile on the thrust of pusher propeller were presented. The pressure distribution and the velocity field changed depending on the shape of the rear fuselage and the ring profile, and their influence on the maximum thrust was shown.

Effect of a fuselage boundary layer ingesting propulsor on airframe forces and moments

The aim of this research project was to investigate the benefits of hybrid-distributed propulsion and boundary layer ingestion applied to an airliner similar in size, range and cruise velocity to an Airbus A320. The power system selected consisted of two under-wing mounted conventional turbofans and an electrically driven boundary layer ingesting fan, located at the rear fuselage. The power required to drive the fan is extracted from both turbofans. The worked carried out and presented in this paper offers a new approach to modelling boundary layer ingestion configurations, consisting of using the sliding mesh method to simulate the rotation of the blades within an airflow. The simulations were performed in two phases. The first stage corresponded to the analysis of an isolated configuration, where fuselage and electric fan were considered to be far away from each other such that there is no aerodynamic interference between the fuselage and the electric fan. During the second phase, the simulations were conducted with the electric fan integrated with the airframe, ingesting the boundary layer around the fuselage. The simulations indicate that ingesting the boundary layer results in remarkable improvements reducing drag and increasing propulsive force, reaffirming the potential of this technology to reduce fuel consumption. However, the results also registered a substantial increment in the pitching, rolling and yawing moments in the integrated configuration with respect to the isolated arrangement. This gain needs to be accounted for during early stages of the design process to maintain the stability and control of the aircraft within acceptable margins.

Flow Characterization of Heavy-Lift Helicopter Rear Fuselage with Pulsed Jets for Separation Control

An experimental study on the flow over the rear body of a 1:7-scaled fuselage model equipped with pulsed jets (PJs) for separation control is conducted in wind-tunnel experiments at length-based Reynolds number of by means of planar particle image velocimetry. The model is representative of an existing heavy-lift helicopter fuselage. The flowfield in correspondence of the loading ramp is investigated along both the longitudinal and transversal directions either at fixed incidence angle , 0°, and 4.5° or monotonically varying it in the range between . Two cases, that is, either without control or under the periodic excitation induced by PJs, have been considered. A statistical analysis of the case without control reveals the presence of a massive separation in the inspected region characterized by crosswise eddies, predominant at positive incidences, and streamwise vortices. Even though the separation mainly occurs slightly downstream the upsweep shape of the rear fuselage, two preferred reattachment regions are recognized either along the ramp or the boom tail for and 0°. When the incidence angle is increased, the separation points move upstream, and the reattachment points are predominantly located at the ramp end. The effectiveness of the control is achieved through a momentum addition of at a reduced frequency for the range of the investigated incidence angles. Benefits in aerodynamic performances result in a drag reduction from 5 to 24% for the incidence angle sweep. Under the periodic excitation, the separation is completely suppressed and the streamwise vortices move closer to the tail boom of the model. The vortical pattern preserves its symmetric spatial organization with respect to the longitudinal direction. A modal analysis using proper orthogonal decomposition (POD) indicates the dominance of periodic crosswise vortical structures evolving along the ramp of the fuselage in phase with the control input. A phase-average reconstruction, based on the POD analysis, highlights that these vortices travel along the ramp and then dissipate by the time they reach the tail boom of the helicopter fuselage.

Substructuring verification of a rear fuselage mounted twin-engine aircraft

Dynamic substructuring allows for the reduction of large complex structures into substructures to increase computational efficiency and to isolate the local dynamic behaviors of concern. However, errors such as truncation, continuity and rigid body mode errors still limit the applicability of this method experimentally. Additionally, the feasibility of implementing substructuring techniques on finite element models of multi-component fuselage structures has yet to be shown in the literature. The objective of this paper is twofold: first to introduce a feasible substructuring methodology that mitigates the experimental and multi-component limitations with current methods; and secondly, to investigate the modal properties and applicability of substructuring analysis on a rear fuselage mounted twin-engine aircraft. This configuration is not well understood in the literature despite having been shown to have increased interior cabin noise and vibration levels. Experimental validation of the computational model was first performed. A substructuring analysis of the validated computational model produced natural frequencies of the local and global modes that agreed within 6.47% on average, and pseudo-orthogonality terms greater than 0.89 for all modes considered. This methodology proved to be useful for generating an accurate representation of local modes within a global structure for the aircraft configuration studied. This will allow for future work to more thoroughly investigate the local modes using innovative design methods with confidence that the local modes will correlate with the global modes.

Wind Tunnel Support System Influence on NASA Common Research Model at Low-Speed Conditions

Computational studies were conducted in order to gain an understanding and to quantify the impact of the model support system and aeroelastic wing deformation on the Common Research Model at subsonic inflow conditions at . Comparisons between results obtained in the European Transonic Windtunnel and from simulations using TAU reveal the extent of the modeling error introduced by neglecting the support structure. RANS simulations with and without the support system show that it has a near-field effect on the model’s rear fuselage, significantly altering the pressure field and thereby impacting the force coefficients, most significantly drag and pitching moment. Aeroelastic wing deformation was also included in the investigation, showing non-negligible impact on the forces. The inclusion of the support system and wing deformation, as well as the use of advanced turbulence models, was shown to promise considerable error reduction potential in comparative studies between wind tunnel and simulation data.

Simulation and Experiment Investigation on Superplastic Forming/Diffusion Bonding Process of a Ti-6Al-4V Alloy Rear Fuselage Part

Superplastic forming/diffusion bonding (SPF/DB) process is widely used in aviation titanium alloy parts forming. PAM-STAMP was utilized to simulate the SPF/DB process of a Ti-6Al-4V double layer structure part in rear fuselage at 920°C and reach the thickness distribution of the part. Then the part was formed based on simulation. The thickness distribution of the practical part was measured and compared with the simulation result. The results show that the thickness distributions of practical and simulated part fit well with each other.

Aerodynamic optimization of helicopter rear fuselage

An optimization process for the rear helicopter fuselage is presented using Genetic Algorithms and Kriging surrogate models. Shape parameterization is carried out with the super ellipse technique employed for the well-known ROBIN fuselage. The simulations were based on the RANS equations solved using the HMB CFD code. It is shown that a decrease of fuselage drag around 2.5% is possible without compromising the structure and the functionality of the design. Combined with an optimization of the helicopter skids, benefits of up to 4.6% were possible. The demonstrated method can be applied to fuselages of any shape during the initial design phase.

The relevance of reliability-based topology optimization in early design stages of aircraft structures

This paper aims to compare the structural designs derived from Deterministic Topology Optimization (DTO) and Reliability-Based Topology Optimization (RBTO). The target is to prove that including uncertain data in preliminary stages of the design process such as topology optimization can help to reduce the weight of the component to be conceived. To compare both approaches the same Reliability-Based Design Optimization (RBDO) sizing problem is formulated over the bar structures emerged from the DTO and RBTO results. The RBTO is performed through the RBDO algorithm Sequential Optimization and Reliability Assessment (SORA) based on its robustness, accuracy and decoupled nature, which allows to use an external optimization software when dealing with the topology optimization problem. Three application examples, including two 2D bar structures and a 3D simplified aircraft rear fuselage are studied. The results show that the designs based on RBTO are lighter than the designs based on DTO when both structures are compared in a subsequent reliability-based size optimization including the same probabilistic information.

Some Special Aircraft

In this study, the authors want to present a few more distinct aircraft from a constructive and functional point of view. Their role has often been determined by the need to achieve or fulfill certain more or less strategic objectives. Such ships will also be built in the future more and more often in order to be able to respond to all new flight requirements and to meet, under optimal conditions, the new and increasingly demanding requirements. Many of the new special aircraft have been built so far to achieve special tasks, or at the request of the defense ministry in some highly developed countries, even with the United States of America. The PA-23 was the first twin-engine design from Piper and was developed from a proposed Twin design inherited when Piper bought the Stinson Division of the Consolidated Vultee Aircraft Corporation. The prototype PA-23 was a four-seater low-wing all-metal monoplane with a twin tail, powered by a two 125 hp Lycoming O-290-D piston engines the prototype first flew 2 March 1952. The aircraft performed badly and it was redesigned with a single vertical stabilizer and an all-metal rear fuselage and more powerful 150 hp Lycoming O-320-A engines. Two new prototypes of re-designed aircraft now named Apache were built in 1953 and entered production in 1954; 1,231 were built. In 1958, the Apache 160 was produced by upgrading the engines to 160 hp (119 kW) and 816 were built before being superseded by the Apache 235, which went to 235 hp (175 kW) engines and swept tail surfaces (119 built). In 1958 an upgraded version with 250 hp (186 kW) Lycoming O-540 engines and adding a swept vertical tail was produced as the PA-23-250 and was named Aztec. These first models came in a five-seat configuration which became available in 1959. In 1961 a longer nosed variant the Aztec B entered production. The later models of the Aztec were equipped with IO-540 fuel injected engines and six-seat capacity and continued in production until 1982. There were also turbocharged versions of the later models, which were able to fly at higher altitudes. The US Navy acquired 20 Aztecs, designating them UO-1, which changed to U-11A when unified designations were adopted in 1962. In 1974, Piper produced a single experimental PA-41P Pressurized Aztec concept.

Experimental Study of Helicopter Fuselage Drag

Experimental data are presented for the parasite drag of various helicopter fuselage components, such as skids, external fuel tanks, and tailplane. The experiments were conducted at the Kazan National Research Technical University (Kazan Aviation Institute) T-1K wind tunnel, investigating four versions of a fuselage similar to the Ansat helicopter. It was found that, for the range of pitch angles , the skids added 80% to the drag of the bare fuselage, whereas the tailplane increased the drag by 20%. At the same conditions, external fuel tanks were found to add 48% to the clean fuselage drag. A simple rotor hub with a tail support added 74% to the bare fuselage in the range of pitch angles . Streamlining the rear fuselage was found to reduce the drag by 16% over the range of pitch angles . Apart from the parasite drag, ideas for drag reduction are also discussed.

Preliminary sizing correlations for the rear-end of transport aircraft

Purpose – The purpose of this study of which this work is only the first part, is the development of conceptual design tools to perform an optimized design of the rear fuselage and tail surfaces. The development of a new and extensive database of transport aircraft and an analysis of certain general, rear fuselage and horizontal stabilizer parameters of the aircraft are presented in this paper. Design/methodology/approach – In addition to the development of a comprehensive high accurate database, linear and non-linear correlations between different parameters of the aircraft have been established. Data were analyzed using comparison criteria between aircraft database based on the mission, the number of engines installed or arrangement of the tail surfaces. Findings – It has been possible to obtain very relevant, linear and non-linear correlations for critical design parameters to optimize the design of the rear fuselage and horizontal tail. Research limitations/implications – In the case of the tail cone,...

FATIGUE LIFE ESTIMATION OF REAR FUSELAGE STRUCTURE OF AN AIRCRAFT

Integrity of the airframe structure is achieved through rigorous design calculations, stress analysis and structural testing. Finite element method (FEM) is widely used for stress analysis of structural components. Each component in the airframe becomes critical based on the load distribution, which in-turn depends on the attitude of the aircraft during flight. Fuselage and wing are the two major components in the airframe structure. The current study includes a portion of the fuselage structure. Empennage is the rear portion of the aircraft, which consists of rear fuselage, Horizontal tail and vertical tail. The air loads acting on the HT also get transferred to rear portion of the fuselage. First step in ensuring the safety of the structure is the identification of critical locations for crack initiation. This can be achieved through detailed stress analysis of the airframe In this project one of the major stress concentration areas in the fuselage is considered for the analysis. Rear fuselage portion with a cargo door cutout region will be analysed. The structure considered for the stress analysis consists of skin, bulkheads and longerons, which are connected to each other through rivets. Aerodynamic load acting on the aircraft components is a distributed load. Depending on the mass distribution of the fuselage structure the inertia forces will vary along the length of the fuselage. The inertia force distribution makes the fuselage to bend about wing axis. During upward bending, bottom portion of the fuselage will experience tensile stress. A cutout region in the tensile stress field will experience high stress due to concentration effect. These high stress regions will be probable fatigue crack initiation locations in the current work, fatigue damage calculation will be carried out to estimate the fatigue life of the structure under the fluctuating loads experienced during flight. Miner’s rule will be adopted for fatigue damage calculation.

Online programming technique for flexible assembly of fuselage

In order to meet the deep flexible demands of "one fixture with more models"and "one fixture with more states"for digital assembly of fuselage,enable the "bridge-frame"flexible tooling to have better capacity for independent adaptability,its kinematic model must be set up and its movement controls strategy must be planned properly by online programming technique. For the locating styles of fuselage and the characteristics of flexible tooling,the kinematic model of flexible tooling was set up,then simulation model-based online programming disciplines,the way of building reference points for assembly and the technology of movement controls strategy planning were studied. The way of building a process-driven simulation environment based on structure function was proposed. In order to generate the codes quickly to control motion,the online programming system for flexible assembly of fuselage was developed. With the assembly of rear fuselage,for example,the accuracy,high efficiency and strong independent adaptability of this system were verified.

Understanding and Reduction of Cruise Jet Noise at Aircraft Level

Within the LINFaN research project, Airbus and Rolls-Royce have jointly investigated the acoustic impact of supersonic jet noise on the fuselage of an Airbus A340 equipped with Rolls-Royce Trent 500 engines in cruise conditions. The main results are presented in this paper. The influence of chevron fan nozzles designed to reduce cruise jet noise on the rear fuselage is investigated. The characterization of both the acoustic field and the aerodynamic flow is carried out. The acoustic data are obtained from a microphone array on the rear fuselage. After specific data de-noising, the acoustic spectra radiated on the fuselage by only the right inner engine are estimated. Beamforming acoustic maps are also computed to localize the jet noise sources. Besides, the supersonic jet flows are characterized by RANS CFD approach. For the baseline round nozzle, the acoustic results show the presence of two distinct broadband shock-associated noise contributions on the fuselage. The first pattern is observed in the far aft section of the fuselage for Strouhal numbers based on the jet mixing diameter and velocity between St = 4 and 6. The other pattern, around St = 1, radiates preferably forward. In addition, the high frequency source is located between 4 and 5 mixed jet diameters past the secondary nozzle exhaust plane, while the low frequency source is located farther downstream between 7 and 8 mixed jet diameters. Shock-associated noise is usually associated with the interaction between shock cells and turbulent shear-layers. We postulate that the high-frequency noise component results from the interaction between the shock-cells in the secondary flow and the inner shear-layer, while the low-frequency component is due to the interaction between shock cells and the outer shear-layer. This interpretation is consistent with two observations. The first observation is a strong modification of the measured high frequency noise when the engine is pushed to high power. At this regime shock cells are present in the primary flow and the inner shear layer interacts with two supersonic flows, hence modifying emitted noise. The second observation is the reduction of low-and high-frequency noise by the introduction of chevron fan nozzles. The flow field results confirm the strong impact of chevrons on the structure of the shock cells in the fan stream: they are consistent with the experimental observations of the characteristic shock noise frequencies. For the Trent 500 engine, appropriate immersive fan chevrons may bring noise reductions by as much as 6dB SPL integrated over a relevant Strouhal number range [0.7; 2]. However, the reduction level due to the chevrons varies with the engine operating conditions. A given chevron nozzle may be found more efficient at reducing noise levels at low engine operating regime than at high regime.

Application of sensor technologies for local and distributed structural health monitoring

SUMMARY The work presented in this paper deals with the application of different sensor technologies for fatigue crack damage monitoring of metallic helicopter fuselages. A test programme has been conducted, consisting of seven fatigue crack propagation tests on aluminium panels (skin with riveted stringers) representative of the rear fuselage of a helicopter. Electrical crack gauges and comparative vacuum monitoring sensors have been used locally to monitor propagating cracks. A network of optical fibre Bragg gratings is presented as a valid possibility for distributed monitoring (based on strain field dependence on damage), alternative to consolidated electrical resistance-based strain gauges. A Smart Layer based on piezoelectric transducers that emit and receive Lamb wave signals has also been analysed in this paper for distributed monitoring. Two damages have been considered: a skin crack artificially initiated on a panel bay and a skin crack propagating from a rivet hole after stringer failure. Damage sensitivity has been evaluated and compared among the considered technologies, thus providing useful recommendations for each considered system, together with advantages and drawbacks concerning their suitability for on-board installation and monitoring. The damage index sensitivity to operative condition, load in particular, is verified and compared with damage effect for distributed networks. A finite element model able to describe any selected feature sensitivity to the monitored damage is also presented as a useful tool for the optimization of the structural health monitoring system design process. Copyright © 2013 John Wiley & Sons, Ltd.

Advanced Techniques of Stress Analysis

This article aims to check the stress analysis technique based on 3D models also making a comparison with the traditional technique which utilizes a model built directly into the stress analysis program. This comparison of the two methods will be made with reference to the rear fuselage of IAR- 99 aircraft, structure with a high degree of complexity which allows a meaningful evaluation of both approaches. Three updated databases are envisaged: the database having the idealized model obtained using ANSYS and working directly on documentation, without automatic generation of nodes and elements (with few exceptions), the rear fuselage database (performed at this stage) obtained with Pro/ ENGINEER and the one obtained by using ANSYS with the second database. Then, each of the three databases will be used according to arising necessities. The main objective is to develop the parameterized model of the rear fuselage using the computer aided design software Pro/ ENGINEER. A review of research regarding the use of virtual reality with the interactive analysis performed by the finite element method is made to show the state- of- the-art achieved in this field. Key Works: analysis stress, safety margins, parametric modeling

Rear fuselage stiffness design of T-tail

The present work is mainly involved in the fuselage stiffness design. It investigates a relationship between the flutter speed of a T-tail and fuselage support stiffness through optimization, which maximizes the critical flutter speed with the weight constraint. Circumferentially Uniform Stiffness and Circumferentially Asymmetrical Stiffness composite thickness of the fuselage skin are taken as design variables. According to the flutter characteristics of a T-tail, 1-order complex respond surface of the generalized additional unsteady aerodynamic forces is constructed by least squares. It enters flutter solutions and improves computational efficiency. And the optimal results show CAS skin of the fuselage more obviously benefits aeroelastic tailoring than that of CUS, under the same weight. When the flutter speed is greater than 200m/s, the stiffness distributions of the rear fuselage are suitable for the practical structures. This present research provides a reference for the detailed stiffness design of the rear fuselage.

T-Tail Flutter Analysis with Simulated Fuselage Stiffness

AbstractThe current work is a contribution to the potential effect of the aft fuselage stiffness modifications on the flutter characteristics of a T-tail. The rear fuselage is simulated by beam elements, based on a fixed T-tail structure. Their stiffness ranges from 1×105 N-m2 to infinite. The present analyses and discussion predict that a combination of out-of-plane or in-plane bending and torsion of the fuselage, fin, and stabilizer affects modes, and critical flutter and divergence speeds and flutter mechanisms, when the fuselage stiffness continuously varies. With the distinct segmental stiffness of the simulated fuselage, the different flutter results are compared and investigated. This paper also points out that the increasing joint stiffness of the stabilizer and fin can benefit the flutter characteristics. The calculated and analyzed database should be available to use for the structural design of the empennage in engineering