Rear Fuselage Research Articles

Civil aircraft cabin noise resource acoustic characteristic and transmission path analysis

Typically, there are mainly 4 different noise resources critical to cabin +F1184noise: auxiliary power unit (APU) noise, environment control system (ECS) noise, engine noise and turbulent boundary layer (TBL) noise. Because of the different acoustic characteristic and transmission path for each resource, their impacts to cabin noise level are not the similar. A vibration and noise test under ground and flight status of an in-service civil aircraft was conducted. Based on the test results, comparing the data under different test status, the acoustic characteristic and transmission path are analyzed for the 4 noise resources in this paper, including distribution characteristic, spectrum characteristic and transmission path. APU noise mainly affects the rear fuselage, ECS noise transmits by ducts, engine noise and TBL noise transmit through side panel.

노즐 특성에 따른 후방동체 온도 변화 연구

In order to enhance the aircraft survivability, infrared signatures emitted by engine parts should be diminished. For its reduction it is necessary for the rear fuselage temperature to be decreased. In this study, numerical modeling of flow fields and heat transfer of nozzle is performed and its temperature distribution along each component wall is predicted. The effects of material characteristics and shape of nozzle wall and radiation shield on the heat transfer are also investigated. Through this numerical analysis, design parameters related to the susceptibility of aircraft are examined.

Prediction of the Rear Fuselage Temperature with Radiation Shield

� Abstract—In order to enhance the aircraft survivability, the infrared signatures emitted by hot engine parts should be determined exactly. For its reduction it is necessary for the rear fuselage temperature to be decreased. In this study, numerical modeling of flow fields and heat transfer characteristics of an aircraft nozzle is performed and its temperature distribution along each component wall is predicted. The radiation shield is expected to reduce the skin temperature of rear fuselage. The effect of material characteristic of radiation shield on the heat transfer is also investigated. Through this numerical analysis, design parameters related to the susceptibility of aircraft are examined. Keywords—Infrared signature, Nozzle flow, Radiation shield, Rear fuselage temperature, Susceptibility I. INTRODUCTION URVIVABILITY of aircraft has been a one of the important factors in modern battlefield in which the standard of technology determines the supremacy. Military forces are engaged in equipping a stealth function to the fighter plane at the early stage of design. The importance of susceptibility problem has been strengthened related to detection and tracking by enemy missiles or seekers. It is well known that aircraft survivability is threatened by detection using radar and thermal infrared(IR) sensor as well as by primary ways such as visual and sound sources. Several varieties of detection systems have been developed and their technology has been more sensitive, therefore signature reduction technology is excessively in need of development. Advanced countries lead the survivability and susceptibility research (1, 2), but they strongly control the outflow of these technologies overseas. High level of infrared signature emitted by aircraft rear fuselage is an easy target for IR detectors, therefore reduction of IR signature is essential for the improvement of survivability. IR signature can be predicted by figuring out the distribution of the rear fuselage temperature, because IR signature is a function of temperature. The numerical analysis of nozzle flow needs to be preceded to estimate the skin temperature precisely,

Bombardier throws down the gauntlet with CSeries airliner

Bombardier Aerospace says its CSeries will deliver a 20% fuel burn advantage, 15% lower operating costs and 25% lower maintenance costs compared to other aircraft in its class. The aircraft's low-weight airframe features extensive use of advanced structural materials, including composites in the wing, empennage, nacelle and rear fuselage. George Marsh reports on the project's progress so far.

A comprehensive PIV measurement campaign on a fully equipped helicopter model

The flow field around a helicopter is characterised by its inherent complexity including effects of fluid–structure interference, shock–boundary layer interaction, and dynamic stall. Since the advancement of computational fluid dynamics and computing capabilities has led to an increasing demand for experimental validation data, a comprehensive wind tunnel test campaign of a fully equipped and motorised generic medium transport helicopter was conducted in the framework of the GOAHEAD project. Different model configurations (with or without main/tail rotor blades) and several flight conditions were investigated. In this paper, the results of the three-component velocity field measurements around the model are surveyed. The effect of the interaction between the main rotor wake and the fuselage for cruise/tail shake flight conditions was analysed based on the flow characteristics downstream from the rotor hub and the rear fuselage hatch. The results indicated a sensible increment of the intensity of the vortex shedding from the lower part of the fuselage and a strong interaction between the blade vortex filaments and the wakes shed by the rotor hub and by the engine exhaust areas. The pitch-up phenomenon was addressed, detecting the blade tip vortices impacting on the horizontal tail plane. For high-speed forward flight, the shock wave formation on the advancing blade was detected, measuring the location on the blade chord and the intensity. Furthermore, dynamic stall on the retreating main rotor blade in high-speed forward flight was observed at r/R = 0.5 and 0.6. The analysis of the substructures forming the dynamic stall vortex revealed an unexpected spatial concentration suggesting a rotational stabilisation of large-scale structures on the blade.

Engine integration based on multi-disciplinary optimisation technique

The paper presents the work related to the engine integration of a Rear Fuselage Mounted Engine configuration in the frame of the DLR/Onera project MDOrmec (“Multi-Disciplinary Optimisation of Rear-fuselage Mounted Engine Configuration”). The developed multidisciplinary optimisation process is based on a mixed-fidelity approach, where the aerodynamics is simulated using CFD methods while the weight, aircraft stability and control are assessed using preliminary design approach. The Breguet mission range is used as the objective value in order to find the best compromise between the aerodynamic performance of the aircraft and its structural weight. The stability and control qualities of the aircraft are also considered in order to ensure similar flight properties as the jet airliner Dornier Do-728, used as reference configuration. The engine-fuselage distance, the longitudinal and the circumferential engine position are used as the design parameters. Since the displacement of the engine impacts the centre of gravity of the aircraft, the longitudinal position of the wing is additionally used as parameter to provide similar longitudinal stability as the reference configuration. In order to ensure the control requirements, the tail planes are automatically sized during the process. The drag coefficient at required lift and zero pitching moment of the full configuration are obtained by using Euler computations at cruise conditions and adapting automatically the angle of attack and angle of incidence of the tail planes. The static margin of the aircraft is also evaluated using Euler computation. The resulting process chain is implemented in the ModelCenter optimisation framework. Finally, the results of the optimisation are presented and discussed in detail.

Study of sunshine, skyshine, and earthshine for aircraft infrared detection

Investigations are reported on infrared (IR) signature studies from internal sources of aircraft and helicopters, e.g. engine heated surfaces and exhaust plume. Due to the increasing sensitivities of IR sensors, irradiances from external sources are investigated here, namely sunshine, skyshine, and earthshine. Their role in determining the IR contrast in the 1.9–2.9, 3–5, and 8–12 µm bands is studied. Estimation of earthshine is emphasized because of its tactical significance in low altitude missions and IR imaging studies. A comparison of IR radiance from aircraft's rear fuselage due to internal and external sources shows different roles in the 3–5 and 8–12 µm bands. The important roles of earthshine and temperature-based radiance of aircraft surfaces are identified in the 8–12 µm band.

Analytical estimation of solid angle subtended by complex well-resolved surfaces for infrared detection studies

The solid angle (Omega) subtended by the hot power-plant surfaces of a typical fighter aircraft, on the detector of an infrared (IR) guided missile, is analytically obtained. The use of the parallel rays projection method simplifies the incorporation of the effect of the optical blocking by engine surfaces, on Omega-subtended. This methodology enables the evaluation of the relative contribution of the IR signature from well-resolved distributed sources, and is important for imaging infrared detection studies. The complex 3D surface of a rear fuselage is projected onto an equivalent planar area normal to the viewing aspect, which would give the same Omega-subtended.

3D fatigue crack propagation analysis of a helicopter component

In this paper, the analysis of fatigue crack propagation on the folding beams mounted on the rear fuselage of a helicopter is performed under complex flight and folding load spectra; a preliminary cumulative damage fatigue analysis is used to select the zone of the structure where cracking is most likely to occur. A damage tolerance analysis is then carried out in the selected zone using an FEM-based approach in conjunction with a Three-Dimensional (3D) automatic crack propagation algorithm developed by the authors. The submodelling technique is used to obtain accurate results in the assessment area with a limited increase in the amount of time required for calculations. The numerical approach is validated by the experimental results of full-scale fatigue tests on the rear fuselage of the helicopter.

Noise of Dual-Stream Beveled Nozzles at Supercritical Pressure Ratios

The nozzles of a dual-stream turbofan engine are operated invariably at supercritical pressure ratios at cruise, thereby producing shocks in the jet plume. Consequently, broadband shock-associated noise is generated in addition to turbulent mixing noise. The jet exhaust noise impinges on the fuselage and is then transmitted into the interior of the aircraft cabin. In a companion paper, the beveled nozzle has been shown to provide a community noise benefit during takeoff. (Viswanathan, K., Elegant Concept for Reduction of Jet Noise from Turbofan Engines, Journal of Aircraft, Vol. 43, No. 3, 2006, pp. 616-626.) Here, the aeroacoustic characteristics of two beveled nozzles, with bevel angles of 24 deg (bevel24) and 45 deg (bevel45), operated at supercritical pressure ratios, are presented under static conditions as well as in the presence of a Mach 0.32 flight stream. The intended application is for the control of aft-cabin noise. The velocity of the primary jet plays a strong role in the reduction of noise at the aft polar angles, with a substantial benefit at higher jet velocities. The magnitude of reduction in overall levels for bevel45 at the aft angles ranges from ∼5 to ∼12 dB under static conditions, depending on the Mach number of the secondary stream. In the presence of a flight stream, the jet conditions, as well as the bevel angle, control the flight effects at different azimuthal angles. It is demonstrated that at typical cruise power settings, the current concept with bevel24 yields a reduction of ∼4 dB in overall levels; the reduction occurs over a wide range of lower frequencies without any increase at the higher frequencies. Given this reduction in impingement levels on the rear fuselage, this design could yield a cabin noise benefit in addition to the benefit for community noise at takeoff conditions.

Infrared Signatures of Low-Flying Aircraft and Their Rear Fuselage Skin's Emissivity Optimization

This paper analyzes the main contributors of infrared (IR) signature in a typical aircraft on a low-altitude mission. Various computational models are used to predict IR radiation from the aircraft. The bands within IR spectrum in which aircraft are susceptible to a typical IR-guided surface-to-air and air-to-air missile, for typical cases of tactical relevance, are identified. Lock-on range for aircraft against a typical missile is also computed. The feasibility of a low-altitude mission against a ground-based IR-guided threat is analyzed. The technique of emissivity optimization of aircraft rear fuselage skin, for reducing its infrared signature, is introduced and compared with other IR signature suppression techniques. The effectiveness of this technique in enlarging the safe flight envelope of aircraft, with respect to threat from heat-seeking missiles, for both surface-to-air and air-to-air missiles, is demonstrated. It is found that earthshine reflected off the aircraft surface plays a crucial role in the effectiveness of this technique against a surface-to-air missile (SAM) in 8‐12 μm band. Nomenclature A = area, m 2 H = spectral irradiance, μW/μm · m 2 h = aircraft altitude, km I = spectral radiant intensity, W/Sr · μm · m 2 J = spectral radiance: comprising emission and earthshine, W/Sr · μm · m 2 L = length, m M = Mach number N = number of discretized elements NEI = noise equivalent irradiance, W/m 2 Rma = distance separating missile and aircraft, km

Skin-Temperature Prediction of Aircraft Rear Fuselage with Multimode Thermal Model

The temperature distribution of the exposed rear fuselage of fighter aircraft, which is necessary for estimating passively emitted infrared signature levels, is predicted. Hot combustion gases flowing through the gas-turbine engine and aerodynamic heating of the rear fuselage at high freestream Mach numbers are the two heat sources. Turbulent forced convection and radiation are the dominant heat-transfer modes, and hence two thermal models are developed: only turbulent forced convection, and combined turbulent forced convection and radiation. Variation in transport and thermophysical properties of fluids is also modeled. Gas-dynamic equations are used to obtain local flow parameters, considering the role of friction, heat transfer, and area variation. A radiation-shape-factor analysis for the complete layout is presented that accounts for mutual, partial, or complete visibility of emitter and receiver elements. A closed-form solution for shape factors between ring elements is derived for the annulus. The error in temperature estimation by neglecting parameters such as aerodynamic heating, variation in transport, and thermophysical and local-flow properties of fluid is found to be significant. We conclude with a discussion of the role of altitude and flight Mach number.

Tailscrape on rear fuselage

Tailscrape on rear fuselage

Mode localization and frequency loci veering in disordered engineering structures

Vibration mode localization and frequency loci veering in disordered structures of aircraft engineering are discussed in this paper. Three T-type beam models of disordered horizontal tails with rear fuselage, a model of two disordered coupled control surfaces and a model of an aircraft with disordered external stores are adopted for the theoretical analysis. The matrix perturbation technique and the geometric theory are used to study the occurring mechanism of the two phenomena. Several simple criteria for predicting the occurrence of the mode localization and frequency loci veering are given. The theoretical prediction of the phenomena is in great agreement with that obtained through numerical calculations of characteristic solutions of the disordered structures.

FEM submodelling fatigue analysis of a complex helicopter component

This paper deals with the analysis of fatigue damage to upper and lower folding beams on the rear fuselage of a naval helicopter which may result from flight and folding loads; an original FEM-based analytical approach is used and is found to be applicable both together with and as an alternative to experimental tests.The finite element model of the helicopter part is created by means of an ABAQUS/Standard finite element programme combined with advanced submodelling techniques, thus enabling assessment of structural stresses. Thanks to submodelling, it is possible to obtain more precise results with only a limited increase in the amount of time required for calculations.The finite element model defines the stress values in points of the beams most affected by flight and folding loads previously applied in experimental tests. Linear damage summation (Miner's rule) and multiaxial stress criterion enable us to evaluate the fatigue damage to the beams under investigation.

Rear Fuselage Transonic Flow Characteristics for a Complete Wing-Body Configuration

The study presents results for experiments conducted in the transonic pressurized T2 wind tunnel of Centre d'Etudes et de Recherches de Toulouse. The 1/80th scale model of a modern subsonic transport aircraft was mounted in the test section using a fin-sting. It consisted of a fuselage, a detachable horizontal rear stabilizer, a belly fairing, and tip-truncated wings. The tip-truncated wings were designed by Aerospatiale to reproduce the correct downwash in the rear fuselage region. Tests were conducted at a freestream Mach number of 0.82 and a Reynolds number evaluated with the model chord length of approximately 2.5 x 10 6 . Various measurements were performed, mainly in the vicinity of the fuselage downstream rear part: oil-flow visualizations, pressure distributions, boundary-layer surveys along the fuselage symmetry lines and some lateral lines using a laser Doppler anemometry system, and near-wake surveys with both pressure and velocity measurements. These experiments were designed to allow a fine description of the flows developing along the rear part of a fuselage in a configuration similar to actual flight applications.

Simulation and analysis of V-22 tiltrotor aircraft forward-flight flowfield

The flowfield about a wing-fuselage-nacelle configuration of the V-22 tiltrotor aircraft at a freestream Mach number of 0.345 and angles of attack ranging from 7 to 16 deg is simulated by using a multizone, thin-layer NavierStokes method. The flow conditions yield a Reynolds number of 18.2 x 10 based on the wing chord for flight at an altitude of 10,000 ft. Major flow features including the three-dimensional flow separation from viscous-vortex interactions observed experimentally are captured by the simulation. The massive separation from the wing and the fore fuselage-sponson region for flows at high angles of attack creates free vortices that pass by the rear fuselage region, including the tail section of the aircraft.

Rear fuselage flow studies on a modern transonic transport aircraft

An experimental program was set up in the CERT/ONERA's T2 wind tunnel, which is transonic, pressurized, and has self-adaptive walls. A l/80th-scale model of a modern transonic transport aircraft was tested in two configurations: 1) without horizontal and 2) with horizontal stabilizer. Various measurements were performed: oil-flow visualizations, pressure distributions, boundary-layer surveys along the symmetry lines using a three-dimensional laser Doppler anemometry system, and wake surveys with both pressure and velocity measurements in a plane downstream of the base. Moreover, both inviscid and viscous computations were carried out separately under experimental conditions. This article reports a detailed analysis of the computations and experiments conducted on the model in both configurations; the objective of this study was to reach a better understanding of the flows developing along the rear part of a fuselage. ESIGNERS have given considerable attention to threedimensional flow separation on various parts of aircraft. Although rear-fuselage separation is less important on a commercial transport aircraft than on one with military-style, rearloading cargo doors, it is still worth investigating because it may be a significant source of drag. It should be emphasized that three-dimensional flow separation cannot be accurately detected by observing tuft stuff during flight tests or woollen threads in a wind tunnel. Generally speaking, because threedimensional flow separation is associated with the storage of fluid in the wall region, it causes an important vortex motion that develops near the lower symmetry line of the and, consequently, spreads downstream of the upswept base. The experiments reported here were conducted in the test section of CERT's T2 wind tunnel, which is transonic, pressurized, and has self-adaptive walls. Tests were performed at a stagnation pressure close to 2 bar, at ambient temperature and at a freestream Mach number of 0.82. The Reynolds number, based on the aerodynamical chord length of the model, was close to 2.5 x 10 6. That figure is comparable to Reynolds numbers reached for industrial wind-tunnel applications, but of course is less than the flight Reynolds number, which is close to 5 x 10 7. For each model configuration (fuselage without stabilizer and fuselage with stabilizer), a number of measurements were performed: oil-flow visualizations, pressure distribu

Meeting the Need — The A330/A340 Programme

UNDERGOING structural and vibration tests at Airbus Industrie's Toulouse facility is the A340 which is being prepared for its first flight in the autumn. It is the first European four‐engine aircraft and the most recent product of the Airbus consortium, with the forward fuselage and centre section being built by Aerospatiale, wing produced by British Aerospace, horizontal stabilizer by CASA and rear fuselage and fin as well as furnishing by Deutsche Aerospace.

Insights on the whirl-flutter phenomena of advanced turboprops and propfans

The new aircraft configurations that have two pusher propellers or propfans located at the rear fuselage are susceptible to a whirl-flutter-related instability characterized by a complex dynamic coupling involving not only the propeller-nacelle whirl modes but also the natural modes of the supporting backup structure. It is shown that the single-engine classical whirl-flutter formulation does not correctly predict these particular situations of instability, which were called whirl induced flutter. The problem is treated in this paper in the form of a parametric study using a 15-degree-of-freedom model. The main stiffness properties of both the supporting backup structure and the engine mounting system are systematically changed and the mechanism of the aforementioned phenomenon is explained on simple physical grounds. a b d h hy,hz Nomenclature = distance between fuselage and propeller centerlines = engine-propeller center-of-mas s distance from the pylon elastic axis = propeller distance from the pylon elastic axis = pylon elastic axis flatwise displacement = aft fuselage displacements in the lateral and vertical directions, respectively = engine polar and transverse moments of inertia, respectively = propeller polar moment of inertia = aft fuselage polar moment of inertia = nacelle polar and transverse moments of inertia, respectively = V/QR - shock mount system lateral and vertical stiffnesses relative to the nacelle, respectively = aft fuselage lateral, vertical, and torsional stiffnesses, respectively