Flying Test Bed Research Articles

Progressive Test and Evaluation Strategy for Verification of KF-X AESA Radar Development

This paper describes a progressive test and evaluation strategy for verification of Korean Fighter eXperimental (KF-X) AESA(Active Electronically Scanned Array) radar development. Three progressive stages of development test and evaluation were officially performed from simulated test conditions to actual operating conditions according to standards: radar function/performance and avionics integration. KF-X AESA radar development is repeatedly verified by progressive stages consisting of five tests: Roof-lab ground test, System Integration Laboratory(SIL) ground test, Flying Test Bed(FTB) test, KF-X ground test, and KF-X flight test. As a result, the risk factor decreases as stages and tests progress. Therefore, development test and evaluation of KF-X AESA radar are successfully performed at low development risk.

The engine mission simulation system: coupled simulation of aircraft and engine in the mission

AbstractWhen designing future aircraft, the entire system including airframe, propulsion system, energy management and mission control must be considered in a coupled approach at an early design stage. To model and study this complex interaction, the Engine Mission Simulation System (EMSS) environment was created in cooperation with the Bundeswehr Technical Center for Aircraft and Aircraft Equipment. EMSS is intended to serve as a digital, flying test bed at the Institute of Jet Propulsion for the rapid and low-cost evaluation of new concepts for propulsion and power supply systems for new and legacy aircraft designs. This tool chain allows a detailed engine simulation under consideration of its flight condition while using an external propulsion and power control system. First, the software architecture of EMSS is discussed, which involves a propulsion simulation in NPSS® coupled with the X-Plane 11 flight simulator via MATLAB®/Simulink®. The aircraft itself is modeled as a six degrees of freedom simulation using the blade element theory, allowing both complex flight maneuvers and unusual aircraft to be easily simulated. The EMSS achieves real-time capability and enables user-friendly pre-design as well as quick trend detection through the ability to perform focused parametric studies. To evaluate the accuracy of the system’s results, the engine model used is validated against experimental data. Subsequently, the aerodynamic characteristics of the aircraft model are verified with a CFD simulation. Finally, a parametric study is used to demonstrate the mission as well as propulsion evaluation capabilities of the Engine Mission Simulation System.

Flight-to-Wind-Tunnel Correlation Study using JAXA’s Flying Test Bed “Hisho”

Predicting aircraft flight performance using ground tests is important for development but difficult, and flight-to-wind-tunnel correlation methods have been developed and improved over many years. This paper examines the flight-to-wind-tunnel correlation of the aerodynamic characteristics of the Japan Aerospace Exploration Agency’s (JAXA) “Hisho” Flying Test Bed by applying typical methods to flight test data. It was found that the minimum drag coefficient of a business jet geometry test model follows the gradient of the Schoenherr equation well for forced turbulent flows. After that, the reference condition for the correlation was set, but discrepancies in the lift and drag coefficients remained between the flight and wind tunnel tests. A preliminary study identified force differences parallel to the axis of the engine nacelle as being caused by spillage drag due to the difference in the capture area ratio (CAR) between the flight and wind tunnel tests. However, the test data were insufficient to reasonably attribute the difference in the normal force component to lift yielded by the engine nacelle. This was due to the fact that the flight test parameters showed an interdependency and the effects of relevant parameters were not fully investigated in wind tunnel tests.

Strain gauge placement optimization methodology to measure multiaxial loads of complex structure

The use of a strain gauge to measure loads is, in some respects, similar to its use in determining stress, but a different approach is required. In load measurement, it is necessary to compile a suitably selected configuration of strain gauges, which can be used to measure often very complex loads of the structure. For designing the engine mount instrumentation for the Flying Test Bed, an optimization tool has been developed. The algorithm and the theory behind the instrumentation design are described in detail. The basic principle is to find the strain gauge configuration that eliminates the measurement error due to the noise in the measured signal as much as possible. The input for optimization is the strain response of the structure to the applied loads analyzed using the FE model. In contrast to the common strategy using purely stochastic methods, this developed tool uses a hybrid approach based on a combination of a heuristic approach with repeated deterministic local optimization. The optimization is focused on the connection of a simple uni-axial strain gauge to a quarter-bridge and a T-rosette to a half-bridge that provides temperature compensation. Furthermore, an approach is proposed that takes into account the possibility of failure of some strain gauges. The instrumentation is thus robust and allows to obtain quality data even in the event of failure of some of the strain gauges.

Flight demonstration of aircraft wing monitoring using optical fiber distributed sensing system

We conduct flight demonstrations of aircraft monitoring by using optical fiber distributed sensing technique. We monitor a main wing of a flying test bed that is a middle-sized passenger jet. We use optical frequency domain reflectometry and long-length fiber Bragg gratings (FBGs). The sensing system measures the strain distribution profile within the FBGs with a mm order spatial resolution in real-time. Thanks to the high spatial resolution, we could observe local strain distributions due to the ribs of the wing. We also monitored strain distribution variations corresponding to various maneuvers during flight. We discussed the correlation with flight maneuvers and the interpretation of the monitoring data.

Flight demonstration of aircraft fuselage and bulkhead monitoring using optical fiber distributed sensing system

We have developed an optical fiber distributed sensing system based on optical frequency domain reflectometry (OFDR) that uses long-length fiber Bragg gratings (FBGs). This technique obtains strain data not as a point data from an FBG but as a distributed profile within the FBG. This system can measure the strain distribution profile with an adjustable high spatial resolution of the mm or sub-mm order in real-time. In this study, we applied this OFDR-FBG technique to a flying test bed that is a mid-sized jet passenger aircraft. We conducted flight tests and monitored the structural responses of a fuselage stringer and the bulkhead of the flying test bed during flights. The strain distribution variations were successfully monitored for various events including taxiing, takeoff, landing and several other maneuvers. The monitoring was effective not only for measuring the strain amplitude applied to the individual structural parts but also for understanding the characteristics of the structural responses in accordance with the flight maneuvers. We studied the correlations between various maneuvers and strains to explore the relationship between the operation and condition of aircraft.

Airframe noise measurements on JAXA Jet Flying Test Bed “Hisho” using a phased microphone array

In 2015, the Japan Aerospace Exploration Agency launched the Flight demonstration of QUiet technology to Reduce nOise from High-lift configurations project to verify by flight demonstration the feasibility of practical noise-reducing aircraft modification concepts. In order to serve as a baseline for comparison before modification, airframe noise sources of the JAXA Jet Flying Test Bed “Hisho” were measured with a 30 m diameter array of 195 microphones mounted on a wooden platform built temporary beside the runway of Noto Satoyama Airport in Japan. A classical Delay and Sum in the time domain beamforming algorithm was adapted for the present study, with weight factors introduced to improve the low-frequency resolution and autocorrelations eliminated to suppress wind noise at high frequencies. In the landing configuration at idle thrust, the main landing gear, nose landing gear, and side edges of the six extended flap panels were found to be the dominant “Hisho” airframe noise sources. Deconvolution by the DAMAS and CLEAN-SC algorithms provided clearer positions of these sound sources at low frequencies. Integration of acoustical maps agreed well with the sound pressure level measured by a microphone placed at the center of the microphone array and gave detailed information about the contribution of each noise source.

スペースプレーン技術実証のための,気球を利用した飛行実験システム( 先進的空力技術 第2部:流れの制御と実験・計測 第22回)

スペースプレーン技術実証のための,気球を利用した飛行実験システム( 先進的空力技術 第2部:流れの制御と実験・計測 第22回)

JAXA実験用航空機「飛翔」飛行試験で取得された実機空力係数の系統不確かさ解析

Time series data of aerodynamic characteristics obtained by flight tests have large variances compared with data obtained by ground testing tools such as wind tunnel tests and numerical simulations. It is necessary to examine the uncertainty of the data obtained by flight tests in order to minimize random and systematic errors. We investigated on-board measuring instruments of the aircraft, and analyzed systematic uncertainties of the aerodynamic coefficients obtained by the flight tests using the JAXA Flying Test Bed ``Hisho''. This paper discusses the result of the systematic uncertainty analysis and the contributing factors to the uncertainty of aerodynamic coefficients.

Aerodynamic performance analysis of a winged re-entry vehicle from hypersonic down to subsonic speed

The paper deals with the aerodynamic performance analysis at a phase-A design level of a reusable and unmanned flying laboratory designed to perform a high lift return from low Earth orbit. The flying test bed aims to provide experimental data in the framework of re-entry technologies. The vehicle concept is a wing body with rather sharp leading edges and slender aerodynamic configuration. Several design approaches, ranging from engineering-based to computational fluid dynamics analyses, have been addressed in this work. In particular, vehicle aerodynamic performances for a wide range of freestream flow conditions, from subsonic to hypersonic regime, including reacting and non-reacting flow and different angles of attack have been provided and compared. Highlights on aerodynamic control surfaces' effects, vehicle longitudinal and lateral–directional static stability are provided as well. Computational fluid dynamics results confirm that high temperature real gas effects seem to be fundamental for the assessment of the concept aerodynamics, especially concerning pitching moments evaluation.

時間領域Beamforming法に基づく実機騒音源解析

A microphone array comprised of 195 microphones was expanded at Noto Airport, and flyover tests on noise source localization around the aircraft were carried out with JAXA flying test bed ‘Hisho’. Analysis condition of time domain beamforming method is adjusted properly, which improves both resolution of noise source and signal-to-noise ratio. In the landing configuration, a nose gear, main gears, and edges of flaps are proved to be main noise sources. Integration of noise source maps obtained by beamforming gives good agreement with sound pressure level measured by a microphone placed at the center of the microphone array.

重力天体着陸 FTB の基本性能検証試験

重力天体着陸 FTB の基本性能検証試験

Assessment of the aerodynamic and aerothermodynamic performance of a high-lift reentry vehicle

This paper deals with the aerodynamic and aerothermodynamic trade-off analysis of a hypersonic flying test bed. Such vehicle will have to be launched with an expendable launcher and shall reenter the Earth atmosphere allowing to perform several experiments on critical re-entry phenomena. The demonstrator under study is a re-entry space glider characterized by a relatively simple vehicle architecture able to validate hypersonic aerothermodynamic design database and passenger experiments, including thermal shield and hot structures. A summary review of the aerodynamic characteristics of two flying test bed concepts, compliant with a phase-A design level, has been provided hereinafter. Several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper.

Development of a Small-scale Supersonic Flight Experiment Vehicle as a Flying Test Bed for Future Space Transportation Research

Development of a Small-scale Supersonic Flight Experiment Vehicle as a Flying Test Bed for Future Space Transportation Research

Antares DLR-H2 – Flying Test Bed for Development of Aircraft Fuel Cell Systems

Fuel cell systems provide a zero-emission and low noise solution for on-board power generation in aircraft. To assure safe and reliable fuel cell operation in aircraft, system testing has to be conducted under realistic operating conditions. On the one hand these conditions can hardly be simulated in laboratory; on the other hand evaluation of the systems with a wide-body aircraft like an Airbus A320 is inflexible, time and cost consuming. Therefore the aerodyne Antares DLR-H2 has been developed as a flying test bed to provide a system carrier for time and cost effective airborne fuel cell system testing. In this presentation the test platform Antares DLR-H2 will be described with the new fuel cell system generation and latest results, comprising laboratory measurements (fuel consumption, power output) and insights on the system design, especially focusing on the concept of direct hybridization, allowing a simple and therefore reliable possibility of coupling fuel cells and batteries.

Hypersonic environment assessment of the CIRA FTB-X re-entry vehicle

This paper deals with design activities performed by CIRA in the frame of its flying test bed vehicle development, named FTB-X. The re-entry scenario with the corresponding loading environment for the proposed vehicle, devoted to research activities in re-entry technologies, is reported and analyzed. The hypersonic environment characteristics of FTB-X are investigated on a trajectory-based design approach by means of several engineering and CFD analyses. Being a phase A design, the air is modeled only as a perfect gas and the analyses were based on laminar flow conditions, while only a preliminary analysis of real gas and turbulence effects has been performed in time. The results show that the aerodynamics derived from engineering design approach is sufficiently accurate for preliminary analysis purpose; moreover, surface heat loads were computed on the vehicle configuration for thermal shield design scopes.

Aerodynamic and aerothermodynamic trade-off analysis of a small hypersonic flying test bed

This paper deals with the aerodynamic and aerothermodynamic trade-off analysis aiming to design a small hypersonic flying test bed with a relatively simple vehicle architecture. Such vehicle will have to be launched with a sounding rocket and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry technologies such as boundary-layer transition and shock–shock interaction phenomena. The flight shall be conducted at hypersonic Mach number, in the range 6–8 at moderate angles of attack. In the paper some design analyses are shown as, for example, the longitudinal and lateral-directional stability analysis. A preliminary optimization of the configuration has been also done to improve the aerodynamic performance and stability of the vehicle. Several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper. The aerodynamic model of vehicle is also provided.

Analysis of Aerodynamic Performances of Experimental Flying Test Bed in High-Altitude Flight

An aerodynamic, computational study has been performed on the 3.9.2_FW50 configuration of the winged experimental flying test bed (FTB-X) of an experimental unmanned space vehicle in the altitude interval 90—110 km, where the vehicle is in transitional regime. The range of the angle of attack was 0—40° and the side slip angle was 15°. The flow field has been solved by the three-dimensional (3D) direct simulation Monte Carlo (DSMC) code: DS3V. The results showed better aerodynamic behaviour both in symmetric and in side-slip flights, but worst longitudinal stability in symmetric flight with respect to the previous version of FTB-X (1.1.2). In fact, both the aerodynamic efficiency and the derivative of pitching moment coefficient in symmetric flight increased. Furthermore, a preliminary analysis about the possibility of an aerodynamic control of the vehicle by deflection of a trailing edge flap has been fulfilled. This analysis has been carried out in terms of the lift and drag forces and pitching moment at the altitude of 70 km in the range of the angle of attack 0—30° and flap deflection 0—30°. The flow field has been solved by a 2D DSMC code (DS2V) and computational fluid dynamic code (H3NS). A thermal analysis has been also carried out for evaluating the heat flux on the flap. This heat flux is comparable with that at the nose stagnation point, and therefore a thermal protection system should be necessary also on the flap. The effect of the flap deflection on the flow separation has been also evaluated. In particular, at high flap deflection angle, the shock wave boundary layer interaction produces a decrease of the airfoil aerodynamic efficiency. Therefore, the increases of lift and drag of the aerodynamic force, as functions of the flap deflection angle, encourage performing similar tests considering the whole vehicle.

Unscented Kalman Filtering for Reentry Vehicle Identification in the Transonic Regime

Parameter identification methods for processing flight data are frequently used to validate and improve a preflight aerodynamic database and, specifically, to reduce the associated uncertainties. In this framework, the paper describes an identification methodology developed for the first flying test bed of the Italian Aerospace Research Center, a demonstrator of technologies relevant to future reusable launch vehicles. The analysis is focused on aerodynamic modeling of the reentry vehicle configuration in the transonic flow regime, in which flight control system performance is affected by a significant level of parameter uncertainty. The parameter estimation is formulated as a nonlinear filtering problem and solved through a multistep approach, in which the aerodynamic coefficients are identified first and, in a following phase, a set of model parameters is updated. In each step, an unscented Kalman filter is used as a recursive estimation algorithm. The methodology is applied to the flight data of the Dropped Transonic Flight Test mission of the vehicle, carried out during the winter of 2007. The reported results demonstrate the good characteristics of the technique in terms of convergence, reduction of uncertainty of the a priori aerodynamic model, and capability of extracting the information content from a rather limited set of flight data on vehicle response.

DTFT-1: Analysis of the first USV flight test

The first dropped transonic flight test (DTFT) of the USV Program, performed with Castor, the first of the two spacecrafts developed within the USV Program, was performed on Saturday 24th February 2007, from Tortolì Airport in Sardinia. At 8:30 a.m. the 340 000 m 3 stratospheric balloon lifted off from the East coast of Sardinia, bringing the flying test bed (FTB) up to 20.2 km before release within the isolated sea polygon controlled by Italian air force test range in Salto di Quirra (PISQ). The mission ended at 10:30 a.m. with the splash-down of the space vehicle. The flight itself was very good, with a nose-up manoeuvre under transonic conditions, reaching a maximum Mach as high as 1.08. The mission target was completely achieved as some 2 million measures were taken related to flight data, housekeeping, as well as 500 aerodynamic and structural experimental sensors. Unfortunately, the vehicle has been damaged more than expected during splash-down. Many national and international institutions and industries contributed to the mission carrying out, under the supervision and technical guide of CIRA: Italian Space Agency, Italian Air Force, Italian Navy, Italian Civil Aviation Authority, Italian Company for Air Navigation Services, Port Authorities, European Space Agency, Techno System Dev., Vitrociset, Carlo Gavazzi Space, Space Software Italia, Alcatel Alenia Space Italy, ISL-Altran Group. The paper reports the actual status of post-flight data analysis.