Full-scale Airplane Research Articles

Development of a Modularized Virtual Flight Simulator based on Multiple Discipline Coupled Method

The problem of aircraft control involves multiple subjects and has always been extremely complicated in mechanics modelling. Virtual flight simulation was put forward in order to reduce the cost of full-scale airplane flying test and wind tunnel experiment. A typical virtual flight simulator includes a computational fluid dynamics (CFD) solver, a structure dynamics (FD) function, and a flight controller. However, existed virtual flight solver usually ignores the universal modelling of dynamics for complex aircraft configuration and the modularization of CFD, FD and flight controller module, which makes it very inconvenient for many in-house virtual flight codes for engineering application and commercial software development. A new framework of virtual flight simulation platform is established with a parallel Reynolds Averaged Navier-Stokes (RANS) solver, a structured moving overlapping grid function, a generic multiple body dynamics (MBD) solver, and a plug-in type flight controller platform. A universal data exchange interface between MBD and flight controller is presented. An X-wing missile flying forward and pitching-up maneuvering validation case is conducted with the current virtual flight simulator, the result is compared with Matlab Simulink solver based on steady aerodynamic database and good agreement is obtained.

Experimental research on impact loading characteristics by full-scale airplane impacting on concrete target

To investigate the impact loading characteristic during an airplane impacting process, an experiment of a full-scale airplane impacting on a movable concrete target was carried out, using a rocket slid propulsion facility. In the experiment, an on-off signal measurement system was used to monitor the airplane’s impacting velocity; the total impacting process as well as the fragment dispersion of the airplane’s wreckage were recorded by a high-speed photography system. Moreover, five accelerometers were installed on the reverse side of the target to measure its acceleration, from which the impact force-time history can be obtained. Moreover, two storage airborne accelerometers were placed on the airplane’s tail to measure the acceleration history of the airplane, from which both the crushing force-time history and the crushing force distribution along the fuselage can be obtained. It is found that the measured impact force-time history is in good agreement with that calculated using the modified Riera impact loading model. This verifies the reliability of using the modified Riera model to predict the impact force acting on the target during the airplane impacting process, and the value of the correction coefficient α in the Riera model can also be determined. In addition, it is observed that, except at the initial moment, the ratio of the crushing impulse to the impacting one remains roughly a constant, which makes it possible to further simplify the Riera model by getting rid of the crushing force distribution curve.

Fatigue and Damage Tolerance Substantiation Approach for a Regional Jet

The overview of the fatigue and damage tolerance substantiation approach for a 70-90 seat class regional aircraft named Mitsubishi Regional Jet, which is now under development by Mitsubishi Aircraft Corporation, is presented. To comply with the new regulatory requirement for prevention from widespread fatigue damage, full-scale airplane fatigue test will be performed with simulating actual airplane structural configuration and typical loading spectra expected in service. In addition to widespread fatigue damage, damage tolerance evaluations assuming manufacturing defect or in-service damage are being performed. Hundreds of structural tests from coupon level to sub-component level are to be used to verify the material properties and analysis methodology.

Different World Views: STEM Students and Student Affairs

Different World Views: STEM Students and Student Affairs

Airplane hot spot monitoring using integrated impedance and guided wave measurements

In this paper, an integrated impedance and guided wave (IIG) damage detection technique using permanently embedded lead zirconate titanate (PZT) transducers is advanced specifically for online monitoring of critical hot spots within a composite airplane wing structure. A full-scale airplane wing segment is designed and manufactured to validate the developed online structural health monitoring system. First, structural hot spots, which are vulnerable to damage due to high stress concentration, are determined through finite element analysis during the design stage of the airplane wing segment. Next, the parameters related to PZT design, size, number, cabling, and location are determined, and the PZTs are embedded near the identified hot spots. Then, the applicability of the IIG technique to detection of debonding on a composite skin and bolt loosening on an aluminum fitting lug is experimentally examined under varying temperature and static loading conditions. Finally, the long-term durability of the PZTs is experimentally investigated using a PZT self-diagnosis technique over 2 years. Copyright © 2012 John Wiley & Sons, Ltd.

Simulation of fluid-solid interaction on water ditching of an airplane by ALE method

Ditching is considered as one of the important aspects of safety performances of airplanes. It is related primarily with the fluid-solid interaction, whose studies mainly depend on experiments at the present time. Numerical and analytical methods for fluid-solid interaction by using 3-D full scale airplane's model will reduce the dependence on the expensive model tests. Numerical studies can be used to estimate the safety of ditching and provide a reference for the crashworthiness design. This article proposes a 3-D dynamical structural model after the real shape of an airplane and an Arbitrary Lagrange-Euler (ALE) fluid-field model, to simulate the fluid-solid interactions caused by low speed ditching. The simulation is based on interaction computational methods, within LS-DYNA nonlinear finite-element code. The results of pressure distributions and accelerating time histories of the airplane's subfloor are discussed in the context of the safety of ditching, and the simulation results and the analytical methods are verified.

Elastic wind-tunnel models for predicting longitudinal stability derivatives of elastic airplanes.

The objective of this investigation was the development of elastic wind-tunnel models to be used in predicting stability derivatives for elastic airplanes, in particular the SST. A novel feature of these models is that they were tested at full-scale dynamic pressures. Scaling laws, manufacturing problems, and testing problems encountered during this development are brought out. The differences between jig shape and design shape of an elastic airplane are illustrated, using the Boeing SST as an example. Application and use of elastic model data to the prediction of full-scale airplane lift and pitching-moment characteristics are presented. Correlation data between theoretical predictions for longitudinal stability and control derivatives and experimental measurements on elastic models are given. These results clearly demonstrate the practicality of elastic models for tests at full-scale dynamic pressures.

Subsonic aerodynamic performance of nozzle installations in supersonic airplanes.

The external aerodynamic performance of supersonic airplane nozzle systems at subsonic speeds is studied and some significant results of analytical and experimental investigations of exhaust system subsonic drag are shown. Experimental techniques are described which can minimize model scaling uncertainties in wind-tunnel tests. An analysis method for successfully predicting drag on axisymmetric bodies at subsonic speeds combines axisymmetric potential flowfield predictions with laminar and turbulent boundary-layer computation procedures. The results for certain applications are compared with experiment, and application of the analysis to the design of asymmetric models for detailed experimental investigations is demonstrated. For accurate extension of model test results to full-scale airplane performance predictions, an analytical method is proposed. This discussion concludes with a mention of some limitations of the analysis and of areas where further work is needed.

Full-scale fatigue test of twin-turboprop transport

When the design work was commenced on the YS-11 transport airplane in 1959, the fatigue behavior of the structure was one of the most cardinal problems in the structural design, then the various design principles were introduced in order to overcome the fatigue problems, and the fatigue test of the full-scale airplane structure was planned to guarantee a fatigue life of 30,000 flight hours.

A full-scale structural fatigue test program

Test methods are described for an experimental fatigue evaluation of a full-scale airplane structure. Specific reference is made to a fatigue test program conducted on the Lockheed P3A wing and gear. A flight-by-flight test loading spectrum comprising 7500 flight hours was constructed to represent the total service experience of the airplane. In the construction of the spectrum, emphasis is directed to the air-ground-air transitions that are provided during a flight-by-flight simulation. Previous studies have shown that these large transitions adversely affect fatigue life. Loads, preprogramed on magnetic tape, were applied to the specimen by hydraulic jacks through a closed-loop electrohydraulic servo system. The program employed 51 servo channels to control the force output of 105 jacks. The test methods devised for this program provided an automated loading system that faithfully duplicated service conditions on the specimen in a relatively short test-time span with a small operating crew. The span for the P3A program was 20 weeks, operating on a three-shifts-per-day basis with a three-man test crew per shift. The methods described can be readily adaptable to supersonic aircraft testing.

Effect of Slipstream on the Longitudinal Stability of a Low Wing Monoplane

T following investigation was carried out last year at the California Institute of Technology. The one-sixth scale self-propelled model used by J. S. Russell and H. M. McCoy was and it was a relatively simple matter to convert it from a high wing to a low wing configuration. A three-bladed fixed pitch propeller having Hamilton Standard 1A1-0 blade form with 10% of each blade cut off at the tip was used, blades being set at 29° at three-quarter radius. Fig. 1 illustrates the important features of the model. The propeller was driven electrically by a motor of approximately one thirty-sixth full scale power. Propeller revolutions were six times their full scale value and therefore the velocity of each blade element closely approximated its full scale amount. Power was measured by counting the revolutions over a ten second interval and simultaneously reading the iorque. The torque mechanism consisted of an Alternating Current Wheatstone Bridge, propeller torque acting against springs and moving a soft iron armature relative to two coils, the impedances of which were balanced one against the other. The springs, armature, and coils are shown in Fig. 2. For this investigation, accurate calculation of the was only necessary insofar as it was in controlling the range to be used in carrying out the tests. The power required and power available curves of a comparable full scale airplane were used as a guide in making each run, on the model being varied to give several readings from a value less than that to more than that available.

Effects of Compressibility on High Speed Flight

T T has been known tha t the air forces exerted -** on aerodynamic bodies follow the so-called square law except for the effects of scale and compressibility. The effects of dynamic scale, or Reynolds number, have become rather well known during the last 10 years, largely because of research conducted in the N.A.C.A. variabledensity wind tunnel. The effects of compressibility have commonly been neglected because until the relatively recent development of the last Schneider trophy aircraft the speeds have been low as compared with the velocity of sound, and the consequent local pressures over the surfaces of high speed airplanes have differed but slightly from atmospheric pressure. At the present time, however, the speeds associated with the fastest airplanes approach 60 percent of the velocity of sound, and the induced velocities over their exposed surfaces lead to local pressures tha t differ appreciably from the pressure of the atmosphere. When this condition exists air can no longer be regarded as an incompressible medium. The effects of compressibility on the aerodynamic characteristics of airfoils have been under investigation by the N.A.C.A. in the high speed wind tunnel, and it is the purpose of this paper to examine the possibility of further increases in speeds in the light of this relatively recent research. The following computations are made for a hypothetical airplane which, however, is not beyond the limits of possibility. The airplane is a scaled-up model which has been tested in the variable-density wind tunnel as part of an extensive investigation of wing-fuselage interference, and represents one of the best wingfuselage combinations thus far produced. The geometric scale factor applied to the model dimensions was determined from the ratio of the model fuselage diameter to the diameter of a fuselage sufficiently large to house a 2300horsepower Rolls-Royce R type engine. The model from which the airplane is derived is a mid-wing cantilever monoplane having a fuselage of circular cross section, and a straight tapered wing having the N.A.C.A. 0018 section at the center and the N.A.C.A. 0009 section at the tip. The principal full-scale airplane dimensions are given in Table I. An outline drawing of the airplane for which computations will be made is given in Fig. 1.