Real Flight Environment Research Articles

Experimental study on shock interaction control of double wedge in high-enthalpy hypersonic flow subject to plasma synthetic jet

The hypersonic shock-shock interaction flow field at double-wedge geometries controlled by plasma synthetic jet actuator is experimentally studied in a Ma = 8 high-enthalpy shock tunnel with the purpose of exploring a novel technique for reducing surface heat flux in a real flight environment. The results demonstrate that increasing the discharge energy is advantageous in eliminating the shock wave, shifting the shock wave interaction point, and shortening the control response time. The oblique shock wave can be completely removed when the actuator's discharge energy grows from 0.4 J to 11.5 J, and the displacement of the shock wave interaction point increases by 124.56%, while the controlled response time is shortened by 30 μs. Besides, the reduction in diameter of the jet exit is firstly proved to have a negative impact on energy deposition in a working environment with incoming flow, which reduces the discharge energy and hence decreases the control effect. The shock wave control response time lengthens when the jet exits away from the second wedge. Along with comparing the change in wall heat flux at the second wedge over time, the control effect of plasma synthetic jet actuator with and without inflation is also analyzed. When plasma synthetic jet works in inflatable mode, both the ability to eliminate shock waves and the shifting effect of the shock wave interaction point are increased significantly, and the wall heat flux is also reduced.

Assessment of drag prediction techniques based on radar data for supersonic vehicles

Flight testing is probably the most accurate approach for defining the aerodynamic characteristics of a flying vehicle. This is justified by the fact that compared to other approaches, either engineering or computational ones, flight testing is indeed the one that perfectly resembles the real flight environment. Flight data are either obtained by measuring the vehicle’s kinematics via onboard sensors or by tracking the vehicle’s flight via (commonly) radars. The advantages of the latter approach are evident in cases where modifying the vehicle design is not possible or in cases where rival/enemy vehicles are examined. The key issue for this approach is the quality and demands of the technique by which vehicle aerodynamic characteristics are reduced from flight-tracked data. In the open literature, different techniques are used to analyze radar data of vehicle position and utilize them to predict vehicle drag coefficient. Each technique has its strengths and pitfalls. In this paper, the three well-used techniques of drag estimation from radar data namely, Least Square (LS), Maximum Likelihood Estimation (MLE), and Stepwise regression (SR), are considered. The underlying principle, the output, and the range of validity for each technique are addressed with emphasis on what differentiates each of them. The viability and validity of the three techniques are addressed based on the own flight testing of a free-flight supersonic vehicle and using the point-mass flight model. Meteorological data are also recorded and flight conditions are used to enhance the resulting calculations. Based on experimental data available from the literature, a comparison is conducted for the techniques examined. Considering the lack of flight data utilized, and in conjunction with data from the literature, it has been concluded that the SR technique outperforms for a higher sample rate, however, the MLE is more feasible when there is a lack of data.

Assessment of Drag Prediction Techniques for a Flying Vehicle Based on Radar-Tracked Data

As far as the aerodynamic characterization of a flying vehicle is concerned, flight testing is probably the most accurate approach as it perfectly resembles the real flight environment. Flight data are obtained by tracking the vehicle via radars if modifying the vehicle design is not recommended/attainable. In the open literature, different techniques are used to analyze radar data; the key issue is the computational demands of each technique and the quality of the resulting aerodynamic characteristics. In this paper, three techniques are considered namely, Least-Square (LS), Maximum-Likelihood Estimation (MLE), and Stepwise Regression (SR), with focus on the prediction of the drag coefficient of a case study vehicle. Features for each technique are addressed based on brief previous published data. A new variant of the MLE method is proposed based on the physical segmentation of the available dataset. Predicted point-mass trajectories are compared with own comprehensive flight test to assess the techniques in concern. It is concluded that Stepwise-regression outperforms with a large dataset, while Maximum-Likelihood Estimation is more feasible considering the lack of data. The proposed variant of the MLE method yields more accurate drag prediction compared to the basic one.

Novel heat flux controlled surface cooling for hypersonic flight

This paper presents a new method in theory and experiment to adjust the transpiration cooling based on the actual measured heat flux. This is particularly useful in extreme heating environments, e.g. atmospheric entry flight or combustion chamber applications. In such environments, usually cooling is set constant based on the vehicle design, yet a mass efficient and performant cooling is sought after. We present a method with real-time surface heat flux determination of the transpiration cooled wall and an automatic adjustment of the cooling. The heat flux is determined based on a system identification process. The heat flux measurement itself is derived non-intrusively from the measurement of pressure inside the plenum, i.e. the region between mass flow controller and porous wall. The particular advantage of this system is that the heat shield material is not weakened by any sensor system and its performance is optimized with respect to cooling needed at a certain heating level. Another new feature of the pressure heat flux transformation is the attenuation of a destabilizing positive feedback loop, where the transpiration cooling controller’s output (i.e. mass flow rate) strongly influences its input (i.e. plenum pressure). We describe the identification of the model parameters for the heat flux determination, which are found and verified by a calibration approach. The controlled cooling was demonstrated in a hot air plasma flow with a reference heat flux of up to 1.4 MW/m^2. The results show the performance and verify the applicability in a real flight environment.

Terminal Explosion Power Test Flight Platform of Warhead

Aiming at the shortcomings of high risk, high cost, low landing speed and low attitude accuracy of warhead terminal dynamic performance test, a low cost, recyclable and modular warhead terminal explosion power test flight platform is proposed. The overall scheme of two-stage flight is constructed, and the aerodynamic load of the flight process is analysed. In addition, the flight trajectory is designed. The numerical simulation results show that the flight platform can send the warhead to the specified test window, accurately control the drop velocity and attitude of the warhead in the test window, simulate the real flight environment of the warhead and the motion law of the terminal point. Therefore, the flight platform can realize the terminal explosion power test and test of the subsonic and supersonic warheads, providing technical support for the development of the warhead.

Closed-Loop Software Architecture for Spacecraft Optical Navigation and Control Development

A software architecture is discussed to develop, run, and test novel autonomous visual spacecraft navigation and control methods in a realistic simulation. This architecture harnesses two main components: a high-fidelity, faster-than-real-time, astrodynamics simulation framework; and a sister software package to dynamically visualize the simulation environment. Maneuvers such as fly-bys and orbit insertions occur over short periods of time and must occur autonomously. Yet, there are no open-source software packages that provide fully coupled spacecraft environments and Flight Software (FSW) enabling Optical Navigation (OpNav) mission scenarios. The presented tool consists of the Basilisk∗ astrodynamics framework interfacing with a Unity-based visualization Vizard that provides a synthetic image stream of a camera sensor. This modular and extensible setup allows optical guidance, navigation and control (GNC) algorithms to be run in a closed-loop format purely in software. The optical measurements are generated in the visualization and passed to the simulation, allowing for real-time control and decision making. This Vizard software has the ability to import shape-models, planet maps, and move into an instrument point-of-view. Paired with open-source image processing libraries, these combined components provide all the necessary pieces to fully simulate autonomous, closed-loop, OpNav scenarios in a faster-than-real-time configuration. This allows for progress in the autonomy sector, as full-fledged FSW can be tested in a real flight environment. Furthermore, this enables more realistic and extensive testing of the software, which in turn increases reliability of the GNC methods as they are refined. This paper presents the Basilisk and Vizard interface architecture, its performance, and develops a example scenario. The image processing methods are displayed and the visualization scenes are validated for pointing purposes, which in turns allows to develop an autonomous pointing algorithm developed in this software environment.

Toward Noise Certification during Design: Airframe Noise Simulations for Full-Scale, Complete Aircraft.

An overview of a recent, NASA-sponsored effort to substantially advance simulation-based airframe noise prediction is presented. An accurate characterization of this component of aircraft noise requires a high-fidelity representation of the finer geometrical details associated with the landing gear and wing high-lift devices, such as slats and flaps, which constitute major noise sources. To achieve this ambitious goal, a systematic approach was followed to extend our current state-of-the-art computational tools to a full-scale, complete aircraft in landing configuration within a realistic flight environment. The work involved several phases: high-fidelity, large-scale, unsteady flow simulations; model-scale experiments in ground-based facilities; and farfield noise prediction for a full-scale, complete aircraft. The comprehensive aeroacoustic database generated during the course of the effort provided a wealth of relevant information for full validation and benchmarking of the advanced computational tools used in the present work. The database will also foster the development of simulation methodologies with improved predictive capabilities.

Study of aerodynamic uncertainty on the complex slender vehicle at high angle of attack

The phenomenon of lateral aerodynamic uncertainty is discovered in the wind tunnel test on the slender vehicle at high angle of attack. To analyze the mechanism of the phenomenon, the numerical simulation method with random pulsation added in the free stream is used, and the aerodynamic uncertainty of the vehicle under the wind tunnel test condition is reproduced by the simulation. The analysis of the flow field leads us to a conclusion that the mechanism of aerodynamic uncertainty is the asymmetry separation, which has a high sensitivity to the disturbance of the flow around the missle. The uncertainty with respect to the angle of attack for a “+” configuration of the main wing is given and the difference of the uncertainty between the “+” and “×” configurations are then analyzed. At last, the action of pulling up through the high angle of attack and the process of temporary atmospheric turbulence are simulated to study the effect of the real flight environment rather than the wind tunnel condition to the aerodynamic uncertainty. Based on these results, it can be concluded that the uncertainty is not looked upon as quite severe in real flight environment as that in the wind tunnel condition, and the maneuverability of flight can be controlled.

Sound Analysis and Synthesis for Audio Simulation System of Flight Simulator

Through simulating the feelings of vision, hearing, motion and force during flight, the flight simulator can build a realistic flight environment on the ground. Therefore, Flight simulators have an important role in pilot training field. Audio simulation system is a key component of the flight simulator, it has direct impact on the realism and immersion of flight simulation. In this paper, we introduced the development procedure of audio simulation system shortly. Software implementation and its key development technology were mainly focused for expressing the necessity of extracting sound sources from original cockpit recordings. Based on the development method and practical recording conditions, we dicussed sound short-time Fourier analysis and synthesis technology, and sound linear prediction analysis and synthesis technology in detail. The objective of using these technologies was to extract sound sources from original recordings and this was also the preparation step of sound simulation.

Multistrand, Fast Reaction, Shape Memory Alloy System for Uninhabited Aerial Vehicle Flight Control

This paper details an investigation of shape memory alloy (SMA) filaments which are used to drive a flight control system with precision control in a real flight environment. An antagonistic SMA actuator was developed with an integrated demodulator circuit from a JR NES 911 subscale UAV actuator. Most SMA actuator studies concentrate on modeling the open-loop characteristics of such a system with full actuator performance modeling. This paper is a bit different in that it is very practically oriented and centered on development of a flight-capable system which solves the most tricky, practical problems associated with using SMA filaments for aircraft flight control. By using well-tuned feedback loops, it is shown that intermediate SMA performance prediction is not appropriate for flight control system (FCS) design. Rather, capturing the peak behavior is far more important, along with appropriate feedback loop design. To prove the system, an SMA actuator was designed and installed in the fuselage of a 2 m uninhabited aerial vehicle (UAV) and used to control the rudder through slips and coordinated turns. The actuator was capable of 20 degrees of positive and negative deflection and was capable of 7.5 in-oz (5.29 N cm) of torque at a bandwidth of 2.8 Hz.

전환제어법칙 설계 및 검증에 관한 연구

The flight control law designed for prototype aircraft often leads to degraded stability and performance, although developed control law verify by non-real time simulation and pilot based evaluations. Therefore, the proper evaluation methods should be applied such that flight control law designed can be verified in real flight environment. The one proposed in this paper is IFS (In-Flight Simulator). Currently, this system has been implemented into the F-18 HARV (High Angle of Attack Research Vehicle), SU-27 and F-16 VISTA (Variable stability In flight Simulation Test Aircraft) programs. The IFS necessary switching control law such as fader logic and integrator stand-by mode to reduce abrupt transient and minimize the integrator effect for each flight control laws switching. This paper addresses the concept of switching mechanism with fader logic of TFS (Transient Free Switch) and stand-by mode of feedback type based on SSWM (Software Switching Mechanism). And the result of real-time pilot evaluation reveals that the aircraft is stable for inter-conversion of flight control laws and transient response is minimized.

HBS-SWMC 환경에서의 전환장치 설계 및 검증에 관한 연구

Although non-real time simulation and pilot based evaluations are available for the development of flight control computer prior to real flight tests, there are still many risky factors. The control law designed for prototype aircraft often leads to degraded performance from the initial design objectives, therefore, the proper evaluation methods should be applied such that flight control law designed can be verified in real flight environment. The one proposed in this paper is IFS(In-Flight Simulator). Currently, this system has been implemented into the F-18 HARV(High Angle of Attack Research Vehicle), SU-27 and F-16 VISTA(Variable stability. In flight Simulation Test Aircraft) programs. This paper addresses the concept of switching mechanism for FLCC(Flight Control Computer)-SWMC(Switching Mechanism Computer) using 1553B communication based on flight control law of advanced supersonic trainer. And, the fader logic of TFS(Transient Free Switch) and stand-by mode of reset '0' type are designed to reduce abrupt transient and minimize the integrator effect in pitch axis control law. It hans been turned out from the pilot evaluation in real time that the aircraft is controllable during the inter-conversion process through the flight control computer, and level 1 handling qualities are guaranteed. In addition, flight safety is maintained with an acceptable transient response during aggressive maneuver performed in severe flight conditions.

Biological component of life support systems for a crew in long-duration space expeditions

Creation of effective life support systems (LSSs) is one of the main tasks of medico-biological support of long-duration space flight. Principles of development of such an LSS will be defined on the basis of number of parameters, including mass-overall and energetic limitations of interplanetary spacecraft, duration of expedition and crew size. It is obvious that including biological subsystems in LSS of long-duration interplanetary space flights will help to form a full-fledged environment for humans in the spacecraft. It would be an appropriate solution for long-term biological needs of humans and important for elimination of possible negative consequences of their long stay under artificial (abiogenous) environment. Experiments with higher plants, conducted on board “MIR” orbital complex and Russian segment of ISS, showed that plant organisms are capable of long-duration normal growth, full development and reproduction without deviations under real space flight environment. These results allow us to assume that greenhouses are potential candidates to be a biological subsystem to be included in the LSS for interplanetary space flight. Inclusion of greenhouse equipment in the spacecrafts will require a number of corrective actions in functional schemes of the existing LSS, i.e. it will lead to redistribution of material streams inside an LSS and increase in functional load of authorized systems. Furthermore, involvement of greenhouse in an LSS of an interplanetary spacecraft requires a number of technical tasks to be cleared. In the present review, we discuss the constructive, technological and mass-transfer characteristics of greenhouse as a component part of the LSS for crews of long-term interplanetary missions, in particular, Mars expedition.

저가형 GPS/IMU를 이용한 UAV 비행 제어 시스템 개발

UAVs(Unmanned Aerial Vehicles) have many applications in military and commercial areas. The flight control system of UAVs is more important than manned aircraft's because the mission of UAVs must be operated without a human pilot. But very heavy and expensive navigation system makes it difficult to develop UAV flight control system. In this research, GPS/IMU integrated navigation filter was developed for light weight/low cost flight control system of small UAVs. With this navigation filter, full flight control system which has real time operating capability has been developed. The performance of the flight control system is basically checked by HILSIM (Hardware In the Loop SIMulation). Finally, the flight control system is verified by showing performance test result under real flight environment.

Pilots Strategically Compensate for Display Enlargements in Surveillance and Flight Control Tasks

Experiments were conducted to assess the impact of display size on flight control, airspace surveillance, and goal-directed target search. Research of 3-D displays has shown that display scale compression influences the perception of flight path deviation, though less is known about the causes that drive this effect. In addition, research on attention-based tasks has shown that information displaced to significant eccentricities can amplify effort, but it is unclear whether the effect generates a performance difference in complex displays. In Experiment 1, 16 pilots completed a low-fidelity flight control task under single- and dual-axis control. In Experiment 2, the control task from Experiment 1 was scaled up to a more realistic flight environment, and pilots performed hazard surveillance and target search tasks. For flight control, pilots exhibited less path error and greater stick activity with a large display, which was attributed both to greater enhanced resolution and to the fact that larger depictions of error lead to greater urgency in correcting deviations. Size did not affect hazard surveillance or search, as pilots were adaptive in altering scanning patterns in response to the enlargement of the displays. Although pilots were adaptive to display changes in search and surveillance, display size reduction diminished estimates of flight path deviation and control performance because of lowered resolution and control urgency. Care should be taken when manipulating display size, as size reduction can diminish control performance.

BAE Systems/EPSRC integrated research programme in aeronautical engineering

AbstractBAE Systems and the Engineering and Physical Sciences Research Council (EPSRC) have recently formed a partnership to invest in strategic research in aerospace and defence. The framework which has been developed as part of this new alliance places a strong emphasis on collaboration. This contrasts with the conventional approach to industry led research which is normally based on the establishment of ‘centres of excellence’ in specific subject areas. By using a collaborative approach, the funding partners aim to benefit from inter-disciplinary collaboration which will take place during the project, giving rise to a more effective use of the invested funds.The objective of the research programme is to develop technologies which would support the design of low cost (both to acquire and operate) flapless unmanned aerial vehicles (UAVs). This work includes fundamental aerodynamic research to provide control forces without the use of conventional flaps, coupled with developments in the areas of control systems, manufacturing engineering, structural engineering, the electromagnetic behaviour of these structures and design optimisation. The output from the research has been extended to include the design, manufacture, assembly and flight of a demonstrator vehicle, which will provide the research teams with a platform on which to evaluate the performance of their technology in a realistic flight environment. A total of fourteen research groups at ten universities are involved in the five year programme, which has a total value of £6·5M (€9·75M).

History of In-Flight Simulationat General Dynamics

General Dynamics‐Advanced Information Systems (GDAIS) (formerly Veridian) has been the primary innovator, developer, and operator of in-flight simulators in the United States, as well as the rest of the world, since 1947. Though other agencies and countries have developed their own in-flight simulators, the focus is on GDAIS accomplishments in this field. In-flight simulation puts the pilot in the real flight environment and has been used in the development of new aircraft, research of flying qualities and flight-control systems, and training of pilots and engineers in these areas. More recent uses have been in the field of display systems and as avionics test beds. Early technologies that led to the development of variable stability aircraft and their earlier applications, are described first, followed by GDAIS’s history in the development and utilization of in-flight simulation, starting in 1949 with the first flight of the F4U-5 and its auxiliary rudder surface, up to the present with the five-degree-of-freedom F-16 Variable Stability In-Flight Simulator and Test Aircraft. Specific case studies are presented that describe the development and distinctive features of each of the GDAIS in-flight simulators and some of the more significant applications of these unique tools are highlighted.

In-Flight Parity Vector Compensation for FDI

The performance of a failure detection and isolation (FDI) algorithm applied to a redundant strapdown inertial measurement unit (IMU) is limited by sensor errors such as input axis misalignment, scale factor errors, and biases. A techique is presented for improving the performance of FDI algorithms applied to redundant strapdown IMUs. A Kalman filter provides estimates of those linear combinations of sensor errors that affect the parity vector. These estimates are used to form a compensated parity vector which does not include the effects of sensor errors. The compensated parity vector is then used in place of the uncompensated parity vector to make FDI decisions. Simulation results are presented in which the algorithm is tested in a realistic flight environment that includes vehicle maneuvers, the effects of turbulence, and sensor failures. The results show that the algorithm can significantly improve FDI performance, especially during vehicle maneuvers.

Increased Piloting Tasks and Performance of X-15A-2 in Hypersonic Flight

In past years, flying a basic X-15 from launch through a dead stick landing on a precise speed and altitude profile was considered a demanding piloting task. The success of each mission was determined by the pilot's ability to perform under the physiological and psychological stresses of flight into a hostile flight environment of high Mach numbers, high accelerations, high altitude, and limiting temperatures. Today the aircraft has been extensively modified to qualify the X-15 as a test bed vehicle for a ramjet engine, which will be carried in place of the movable ventral, providing real time flight environment for testing and developing an operational prototype ramjet for use on future hypersonic cruise vehicles. The additional systems pictured in Fig. 1 (external tanks, ramjet, external window, and ablative coating), their associated emergency procedures and the degraded flying characteristics of the X-15A-2 have increased the pilot's workload significantly over the standard X-15 mission.