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形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2021-06-23, 資料番号: PA2210035000

This paper focuses on the prediction of disturbance effects of the vertical acceleration of an aircraft flying in atmospheric turbulence. To this end, 5-hole probes with high-dynamic differential pressure sensors are installed in front of a fixed-wing unmanned aircraft system (UAS) and a manned experimental aircraft to measure the local airspeed and angle of attack of the airflow. Test flights are performed in light, moderate and severe turbulence to assess the anticipating character and the accuracy of the predicted acceleration. Thereby, depending on the flown airspeed, anticipation times up to 0.1 s are observed. For the UAS the prediction accuracy is assessed to be 71.19% for moderate turbulence and 71.05% for severe turbulence, where vertical acceleration disturbances higher than 30 m/s2are measured. The first manned test flight revealed a prediction accuracy of 61.97%.

This paper describes short procedure for preliminary estimation of power capacity for the air cushion landing gear (ACLG) lift equipment and provides statistical and design data on lift equipment used for the range of existing experimental aircraft. Examples of how the ACLG design affects power capacity required for operation of the ACLG, are considered.

The adoption of conventional “black metal” composites in aircraft structures has lead to an improvement in performance over previous metallic designs. With the maturation of automated fiber placeme...

Electric and hybrid-electric propulsion offer unique possibilities to improve aircraft efficiency and performance. The potential advantages are fuel savings, lower emissions, and reduced noise. Bec...

Fault diagnosis typically consists of fault detection, isolation and identification. Fault detection and isolation determine the presence of a fault in a system and the location of the fault. Fault identification then aims at determining the severity level of the fault. In a practical sense, a fault is a conditional interruption of the system ability to achieve a required function under specified operating condition; degradation is the deviation of one or more characteristic parameters of the component from acceptable conditions and is often a main cause for fault generation. A fault occurs when the degradation exceeds an allowable threshold. From the point a new aircraft takes off for the first time all of its components start to degrade, and yet in almost all studies it is presumed that we can identify a single fault in isolation, i.e. without considering multi-component degradation in the system. This paper proposes a probabilistic framework to identify a single fault in an aircraft fuel system with consideration of multi-component degradation. Based on the conditional probabilities of sensor readings for a specific fault, a Bayesian method is presented to integrate distributed sensory information and calculate the likelihood of all possible fault severity levels. The proposed framework is implemented on an experimental aircraft fuel rig which illustrates the applicability of the proposed method.

В статье рассматриваются научные достижения ученых ЛИИ ГВФ и их вклад в развитие отечественной авиации в предвоенные годы. Авторы обращают внимание на основные научные исследования, труды и разработки в области авиационной науки и техники. В статье рассказано о наиболее известных образцах самолетов, созданных в институте до начала Великой Отечественной войны, их уникальность и новаторство., The article discusses the scientific achievements of scientists of the LII GVF and their contribution to the development of domestic aviation in the prewar years. The authors draw attention to the basic scientific research, works and developments in the field of aviation science and technology. The article describes the most famous aircraft models created at the institute before the start of World War II, their uniqueness and innovation., №1 (2020)

Vectored thrust is a powerful flow control technology that has long been used in rockets, launch vehicles, and other space systems. The technology has been more recently applied to atmospheric flig...

The volume of passenger air travel has increased rapidly since the first commercially viable passenger aircraft, the DC-3. Over this period, commercial aircraft increasingly appear to have consolid...

A trajectory generation approach for the radius to fix leg type is presented, which uses clothoids to ensure smooth transitions between straight and circular sections of the flight path. The derived algorithm is implemented in an integrated modular mission management system for manned and unmanned aerial systems. The trajectory generation module generates inputs for a nonlinear dynamic inversion based 3D trajectory controller using second order error dynamics. By using clothoid segments at the turn-in and turnout transitions, continuous curvature paths can be computed. The capability of the proposed method is demonstrated in flight with a CS-23 experimental aircraft.

This paper summarizes the status of ongoing NASA research supported over the past eight years to advance computational capabilities for modeling civil aircraft loss of control due to airframe damage or wing stall. The research is motivated by a desire to exploit the capabilities of computational methods to create augmented flight simulation models that improve pilot training for such loss-of-control scenarios. Flight of aircraft with either airframe damage or operating near and beyond the stall boundary encounters additional nonlinear aerodynamic influences on stability and control from dynamic motions that, if not included in flight simulation models, may lead to incorrect pilot responses. In the present work, both low- and high-fidelity computational methods are explored for analyzing such nonlinearities. The challenge of creating nonlinear reduced-order models from high-fidelity computational data is also addressed. At the beginning, few guidelines were available for computing or modeling the dynamic s...

With the development of system reliability for aircraft, human error is becoming one of the main factors that leading to flight accidents, which has been a key and difficult challenge for the design and airworthiness authorized departments. This paper analyzes the airworthiness certification process based on CCAR 25, including the selection of regulations relevant to human factor based on SHELL model and T.E.S.T matrix, the means of compliance verification for experimental aircraft based on MOC5 and MOC6; the compliance test program is determined, and a data analysis example based on fuzzy evaluation method is shown. The thought for airworthiness certification of human factor and verification method proposed in this paper has been applied preliminarily in ARJ21-700, and it can provide the technical support for the design and airworthiness certification of human factor for other civil aircraft such as C919.

This paper presents the modeling, system identification, simulation and flight testing of the airbrake of an unmanned experimental aircraft in frame of the FLEXOP H2020 EU project. As the aircraft is equipped with a jet engine with slow response an airbrake is required to increase deceleration after speeding up the aircraft for flutter testing in order to remain inside the limited airspace granted by authorities for flight testing. The airbrake consists of a servo motor, an opening mechanism and the airbrake control surface itself. After briefly introducing the demonstrator aircraft, the airbrake design and the experimental test benches the article gives in depth description of the modeling and system identification referencing also previous work. System identification consists of the determination of the highly nonlinear (saturated and load dependent) servo actuator dynamics and the nonlinear aerodynamic and mechanical characteristics including stiffness and inertia effects. New contributions relative to the previous work are a unified servo angular velocity limit model considering opening against the load or closing with it, the detailed construction and evaluation of airbrake normal and drag force models considering the whole deflection and aircraft airspeed range, the presentation of a unified aerodynamic - mechanic nonlinearity model giving direct relation between airbrake angle, dynamic pressure and servo torque and the transfer function-based modeling of stiffness and inertial effects in the mechanism. The identified servo dynamical model includes system delay, inner saturation, the aforementioned load dependent angular velocity limit model and a transfer function model. The servo model was verified based-on test bench measurements considering the whole opening angle and dynamic load range of the airbrake. New, unpublished measurements with gradually increasing servo load as the servo moves are also considered to verify the model in more realistic circumstances. Then the full airbrake model is constructed and tested in simulation to check realistic behavior. In the next step the airbrake model integrated into the nonlinear simulation model of the FLEXOP aircraft is tested by flying simulated test trajectories with the baseline controller of the aircraft in software-in-the-loop (SIL) Matlab simulation. First, the standalone airbrake simulation is compared to the SIL results to verify flawless integration of airbrake model into the nonlinear aircraft simulation. Then deceleration times with and without airbrake are compared underlining the usefulness of the airbrake in the test mission. Finally, real flight data is used to verify and update the airbrake model and show the effectiveness of the airbrake.

International audience; Quantification and improvement of safety is one of the most important objectives among the General Aviation community. In recent years, machine learning techniques have emerged as an important enabler in the data-driven safety enhancement of aviation operations with a number of techniques being applied to flight data to identify and isolate anomalous (and potentially unsafe) operations. Energy-based metrics provide measurable indications of the energy state of the aircraft and can be viewed as an objective currency to evaluate various safety-critical conditions across a heterogeneous fleet of aircraft and operations. In this paper, a novel method of identifying instantaneous anomalies for retrospective safety analysis in General Aviation using energy-based metrics is proposed. Each flight data record is processed by a sliding window across the multi-variate time series of evaluated metrics. A Gaussian Mixture Model using energy metrics and their variability within each window is fit in order to predict the probability of any instant during the flight being nominal. Instances during flights that deviate from the nominal are isolated to identify potential increased levels of risk. The identified anomalies are compared with traditional methods of safety assessment such as exceedance detection to highlight the benefits of the developed method. The methodology is demonstrated using flight data records from two representative aircraft for critical phases of flight.

We present a model-based approach to wind velocity profiling using motion perturbations of a multirotor unmanned aircraft system (UAS) in both hovering and steady ascending flight. A state estimation framework was adapted to a set of closed-loop rigid body models identified for an off-the-shelf quadrotor. The quadrotor models used for wind estimation were characterized for hovering and steady ascending flight conditions ranging between 0 and 2 m/s. The closed-loop models were obtained using system identification algorithms to determine model structures and estimate model parameters. The wind measurement method was validated experimentally above the Virginia Tech Kentland Experimental Aircraft Systems Laboratory by comparing quadrotor and independent sensor measurements from a sonic anemometer and two SoDARs. Comparison results demonstrated quadrotor wind estimation in close agreement with the independent wind velocity measurements. Wind velocity profiles were difficult to validate using time-synchronized SoDAR measurements, however. Analysis of the noise intensity and signal-to-noise ratio of the SoDARs proved that close-proximity quadrotor operations can corrupt wind measurement from SoDARs., Comment: 30 pages, 19 figures

Light sport aircraft builders widely construct experimental aircraft for their personal enjoyment for both building and flying purposes. A variety of engines are used as powerplants for their aircraft. The Volkswagen (VW) automotive engine is one of the most reliable engine conversions for amateur-built aircraft. The aim of this paper is to investigate the cylinder head temperature of the VW-Aero conversion engine during flight conditions by using CFD. The models of the front and rear cylinder barrels include a cylinder head which is constructed with the addition of a baffle plate and air inlet design shape based on internal air flow patterns. The simulations were performed at a velocity of between 50–100 kt (58–115 mph) at a flying altitude of 6,000 feet as per typical flight conditions. The results showed that the rear cylinder lacked cooling air because it was obstructed by a front cylinder. The rear cylinder encountered heat soak conditions which resulted in higher temperature when compared with what the front cylinder experienced. At a velocity of below 80 kt (92 mph), the incoming air velocity through the engine air duct was inadequate to maintain the engine cylinder’s head temperature so that it rises above a normal operating temperature range. At a velocity of above 80 kt (92 mph), the internal baffle plate was capable of feeding air through the controlled path between baffle plate and cylinder fin, which consequently resulted in a controllable engine cylinder barrel and head temperature to be within the normal operating temperature range. The aircraft flying at a velocity of less than 80 kt (92 mph) and equipped with a VW engine would be likely to experience an overheated engine problem due to insufficient cooling air.

Morkūnas Aurimas. „Eksperimentinio ultralengvojo elektriniais varikliais varomo orlaivio projektavimas ir jo stiprumo savybių tyrimas“. Magistro baigiamasis projektas / vadovas prof. dr. Vaidas Lukoševičius; Kauno technologijos universitetas, mechanikos inžinerijos ir dizaino fakultetas. Studijų kryptis ir sritis: Aeronautikos inžinerija (E14) inžinerijos mokslai., Morkūnas Aurimas. “Experimental ultralight aircraft powered by electric engines designing and strength characteristics testing“. Master's Final Degree Project / supervisor prof. dr. Vaidas Lukoševičius; Faculty of Mechanical Engineering and Design, Kaunas University of Technology. Study field and area (study field group): Aeronautical Engineering (E14), Engineering Science.

Electric-motor powered aircraft possess the ability to operate with efficient energy delivery, but lack the operational range of internal combustion engine powered aircraft. This range limitation requires the use of high aspect ratio, thin-chord wings to minimize aerodynamic drag losses, which results in highly loaded composite spar structures. High aspect ratio wings are required to increase mission durations for a NASA-developed experimental multi-rotor electric powered aircraft denoted as the Scalable Convergent Electric Propulsion Technology and Operations Research (SCEPTOR) or “X-57”. This paper examines the structural performance of the composite main wing spars to validate spar strength using ply-based laminate finite element methods. Geometric scaling of a main spar test-section was initially proposed for proof-testing but sacrificed stability. Ply-based structures modeling with local structural features was implemented as a risk-reduction methodology. Ply-based modeling was selected to augment the conventional “building block” approach to reduce risk, and leverage a “performance-based” approval processes encouraged in Federal Aviation Administration (FAA) design guidance. Therefore, ply-based laminate modeling of the full-scale main spar and forward spar shear-web attachments were subsequently undertaken to determine load path complexity with predicted flight loads. Ply-based modeling included stress concentrations and interlaminate behavior at interface locations that can be obscured in traditional finite element sizing models. Analysis of the wing spar laminate ply-based models compared with bearing test coupon performance was used to reduce future wing assembly proof-testing burden and facilitate performance-based flight hardware safety for the X-57 experimental aircraft.

DESIGNED BY AEROSPACE and defence company Lockheed Martin, the X-59 Quiet SuperSonic Technology (QueSST) aircraft is Nasa's first large-scale, piloted X-plane (or experimental aircraft) in more than three decades.

This paper investigates the design and flight test of two fixed wing aircraft connected at the wing tips. Connecting multiple aircraft introduces flexible modes into a typically rigid body system. These flexible modes make manual control of the entire system extremely difficult if not impossible. An autopilot system that seeks to keep this aircraft system wings level and a constant pitch angle is investigated here. The autopilot system is shown to work in an example simulation for a two body aircraft connected at the wing tips. An experimental aircraft system is also designed, built and flown with reasonable success proving the implementation of said controller on a real system.

The human pilot is a crucial component of the pilot/vehicle system. This article focuses on the in-flight ability of an adaptive receding horizon optimal controller to specify favorable flying qualities. Moreover, in the case of a deteriorated vehicle, the controller serves to protect the desirable flying qualities by automatically compensating for the new aircraft characteristics. Although the receding horizon optimal controller is able to protect the handling qualities as long as adequate physical control authority exists, particularly severe failures can result in heavily limited authority where encounters with actuator rate limits may occur. The controller itself is robust to rate-limit encounters, however, the handling qualities suffer because the authority no longer exists to deliver the performance expected by the pilot. To protect handling qualities and mitigate the risk of pilot-induced oscillation past this exhaustion of control authority, the controller is integrated into the Smart Adaptive Fli...

This paper describes the design, implementation and flight testing of flight control laws based on Incremental Nonlinear Dynamic Inversion (INDI). The method compares com- manded and measured accelerations to compute increments on the current control de ec- tions. This results in highly robust control solutions with respect to model uncertainties as well as changes in aircraft dynamic characteristics of failure cases during flight. At the same time, the complexity of the algorithms is similar to classical ones. The key for practical implementation is in ensuring synchronization between angular acceleration and control de ection measurements or estimates. This paper describes the design and vali- dation of manual attitude control functions for a Cessna Citation II experimental aircraft, covering control structure design, application of INDI, design optimization, robustness analyses, software implementation, ground and flight testing. For comparison, also control laws based on classical Nonlinear Dynamic Inversion were implemented and flown. The flight tests were highly successful and marked the first successful demonstration of INDI on a CS-25 certified aircraft. The flight test results proved that INDI clearly outperforms NDI and provided valuable lessons-learnt for future applications.

A flight controller based on Simple Adaptive Control (SAC) method was implemented to multipurpose experimental aircraft MuPAL-α developed by JAXA (Japan Aerospace Exploration Agency) to demonstrate the effectiveness of SAC flight controller in the serious fault case where the longitudinal stability was suddenly reduced in the air. A parallel feed-forward compensator was added to make the whole system meet the ASPR (Almost Strictly Positive Real) condition that is necessary for the plant to be stabilized by the SAC controller. In order to represent the fault case in a real flight test, a “pseudo-fault term” was added to the original controller equipped with MuPAL-α. Through numerical simulations, hardware in the loop tests and flight experiments, it was revealed that the SAC could stabilize the aircraft with such a fatal fault while a conventional PID control system could not stabilize the plane.

System wide information management (SWIM) is a recent concept for providing efficient and safe information sharing in civil aviation. The goal in implementing SWIM is to achieve seamless air traffic management (ATM) services worldwide through the efficient integration of the current ATM systems. In this paper, we consider the wireless media to access SWIM. In the International Civil Aviation Organization (ICAO), the aeronautical mobile airport communications system (AeroMACS) is standardized for the next generation airport surface communication system. It is effective to share various types of information among air traffic controllers, pilots, and airport operators, not only on airport surface, but also in taking off and landing by AeroMACS. However, AeroMACS is intended to be used at speeds of 120km/h or below. In this paper, we study the possibility of using AeroMACS while an aircraft is moving at over 120km/h, as would occur in takeoff and landing situations. We conducted flight tests using an AeroMACS standard-compliant prototype system and an experimental aircraft. In the experiments, we obtained a received signal strength indicator (RSSI) to investigate 5GHz ground-to-air propagation. The results revealed that an onboard antenna configuration involving vertical placement at the top and bottom of the aircraft body is effective in achieving multiple input multiple output (MIMO) operation. Additional throughputs measurements showed that it was possible for AeroMACS to access SWIM both in takeoff and landing situations at an aircraft speed of 200km/h.

This paper presents a design methodology for the rejection of wind gusts during flight. The proposed application is based on the Simple Adaptive Controller (SAC) methodology, but differs from other SAC applications in the sense that the reference model utilised is an Unknown Input Observer (UIO). As a consequence, the adaptive controller is fed by an estimate of the wind gust and will only act in the event of such a disturbance occurring, retaining nominal performance otherwise. Although stability guarantees of the flight control system proposed follows closely previous work, some differences are found and commented upon in this work. Success of the proposed scheme is demonstrated via simulation using a linear model of the MuPAL-α experimental aircraft of the Japanese Aerospace Exploration Agency (JAXA), and the results are compared with an implementation of the original SAC formulation.

Recently, much attention has been paid to unmanned aircraft systems (UAS), especially for civilian purposes such as disaster prevention and transportation. This paper considers big long-flight-type UASs remotely piloted from ground station by radio communication link, since these types of UASs require higher reliability. The purpose of this paper is to consider a communication system which transmits real-time video of the surroundings of aircraft from unmanned aircraft to the ground pilot. In this paper, the aeronautical mobile airport communications system (AeroMACS) is considered as a command and non-payload communications (CNPC) link between unmanned aircraft and ground pilot. Flight tests using the experimental aircraft and AeroMACS prototype system are conducted to obtain the air-to-ground propagation profile. In addition, the placements of airborne antennas are considered, because the positional relationship between airborne antenna and ground fixed antenna is not constant during the flight. The results of flight tests show the effectiveness of our system, and the received signal strength indicator (RSSI) obtained by flight tests has the same tendency as the free space propagation model. Finally, the results of flight tests show that a vertical placement on top and bottom of aircraft body is effective for the multiple inputs and multiple outputs (MIMO) antenna configuration of UAS.

Reconfigurable aircraft are capable of in-flight shape change to increase mission performance or provide multimission capability. This paper examines a fundamental issue raised by aircraft reconfigurability: design of an optimal, lightweight structure to support all of the loads experienced by each configuration. This paper reviews current methods used to design such structures and shows that simultaneous optimization can produce a lower-weight structure than other techniques. The developed optimization routine combines simultaneous optimization with a method based on Liang’s performance-based optimization for determining the optimum location of the structural members of the substructure. An overpopulated finite-element model serves as a starting point. The initial set of structural members includes ribs, spars, and a series of cross members that diagonally connect the ribs and spars. Using an iterative process, the optimization routine removes members from the design if they do not carry enough load to j...

This paper describes the power and command system architecture of the X‑57 Maxwell flight demonstrator aircraft. The X-57 is an experimental aircraft designed to demonstrate radically improved aircraft efficiency with a 3.5 times aero-propulsive efficiency gain at a “high‑speed cruise” flight condition for comparable general aviation aircraft. These gains are enabled by integrating the design of a new, optimized wing and a new electric propulsion system. As a result, the X‑57 vehicle takes advantage of the new capabilities afforded by electric motors as primary propulsors. Integrating new technologies into critical systems in experimental aircraft poses unique challenges that require careful design considerations across the entire vehicle system, such as qualification of new propulsors (motors, in the case of the X-57 aircraft), compatibility of existing systems with a new electric power distribution bus, and instrumentation and monitoring of newly qualified propulsion system devices.

An L 1 adaptive controller is proposed to control the short period dynamics of the F-18 HARV experimental aircraft when affected by bounded uncertainties, bounded external disturbance and time delay. The reference model is improved with a linear quadratic regulator (LQR) applied to the short period nominal dynamics model. The L 1 adaptive technique is used to attenuate the effects of the uncertainties, disturbance and time delay and to track the response of the augmented reference model. Simulation results verify that the proposed L 1 adaptive controller ensures smooth transient behaviour for the control and output signals simultaneously, in addition to asymptotic tracking.

Landing high-drag, low-inertia light sport and experimental aircraft requires significant pilot cognitive workload, causing this to be one of the most dangerous phases of flight for these types of aircraft. Reducing pilot workload through a standardized landing procedure will help maximize safety during landing. The Federal Aviation Administration (FAA) provides many commercial aircraft with standardized procedures for landing, but the same level of support does not exist for light sport and experimental aircraft. In order to calculate a landing approach that can be customized to multiple light sport and experimental aircraft, a model has been developed to test various speed and power combinations with the given runway parameters to determine the best approach(es) for that given aircraft and runway. The simulation model is used to minimize the workload required by ensuring that only the engine power (revolution per minute-RPM) or the approach airspeed (miles per hour) variable remains constant throughout the landing sequence. Since the pilot needs to adjust the speed and power settings in order to follow the landing procedure, keeping one of these parameters constant will minimize pilot workload.

For the next generation of aircraft, new engine and fuselage concepts are investigated because of the increasing demand for fuel efficiency and the CO2/NOx restrictions declared by the European Union. A lightweight design of civil aircraft combined with efficient engines like the counter-rotating open rotor (CROR) is a key technology to reach these ambitious goals. Especially in the lower frequency range, where the CROR blade-passing frequencies (BPF’s) are dominating the acoustic response, the transmission loss of lightweight structures is typically poor. Active control techniques was tested on plates and beams to increase the system performance in terms of global vibration or sound power reduction. Yet to the authors’ best knowledge, only very little work has been conducted to establish sensor or actuator position optimisation on a complex aircraft structure. The presented paper investigates the potential of a sensor position optimisation on a Dornier 728 experimental aircraft of the German Aerospace center (DLR). This experimental aircraft provides a real fuselage structure and the multi-sinusoidal CROR excitation is induced by a 112-channel loud speaker array. The system model of a fuselage segment is experimentally identified, a sensor optimisation is established and compared to empirically placed sensor configurations. The sensor optimisation includes the simulation of an multiple-input multipleoutput (MIMO) active feedforward vibration controller and the resulting sound radiating vibration patterns. It is shown in the experiments that the feedforward controller using optimised sensor locations reduces the radiated sound power by 7 dB, whereas the empirically placed sensors achieve only 4.5 dB. Furthermore, the potential of the proposed optimisation method regarding sound power reduction performance and the number of required sensors is shown for a real aircraft structure and a complex acoustic excitation.

Landing Test. An extraordinary array of aircraft were tested during his tenure, including the Douglas Skyrocket D-558-2; X-series airplanes from X-1E through Lifting Bodies; most of the Century Series of fighters; numerous test-bed aircraft incorporating unique technology such as the digital fly-by-wire system, integrated propulsion control system, and supercritical wing; two variable-stability aircraft (F-100C and Lockheed JetStar); the Lunar Landing Research Vehicle; remotely piloted vehicle research, including drones for aerodynamic and structural testing, 3=8-scaled F-15 spin vehicle, and highly maneuverable aircraft technology; and several radio-controlled experimental aircraft now called mini remotely piloted vehicles.

PurposeThe main targets of the work are analysis and simulation of flying laboratory performance. In particular, synthesis of control system for handling qualities change and evaluation in flight are taken into consideration.Design/methodology/approachModification of handling qualities is obtained by applying indirect flight control system (FBW). The properties of the optimal controller are calculated through the indirect (implicit) model‐following method. In particular, the modified version based on the computer simulations is used.FindingsCalculation and simulation concern the synthesis of desired handling qualities of the general aviation aircraft PZL‐M20 “Mewa” equipped with indirect (FBW) experimental flight control system. Results of the simulation show that the flying laboratory has the same properties as modeled aircraft, and it is possible to say that handling properties concern attitude orientation of the experimental aircraft is similar to modeled commuter aircraft.Practical implicationsThe result of research can be implemented on a project of the flying laboratory based on general aviation aircraft PZL M20 “Mewa”.Originality/valueThe paper presents the practical approach for synthesis of the “Simplified total in flight simulator” performance which can be used for analysis of handling qualities of general aviation aircraft equipped with FBW. Research of this type focuses on military and transport airplanes however, there are no published works in the area of small aircraft so far.

Mogas has been an alternative to leaded fuel since 1964 when Experimental Aircraft Association (EAA) began testing on it. However, in order for mogas to be used in aircraft engines and air frame modification, approval via the Supplemental Type Certificate (STC) authorization from Federal Aviation Administration (FAA) is mandatory. Cessna on 01.06.2010 evaluated alternative fuels with ethanol based fuels approved by FAA STCs for use in some single engine airplanes. However, Cessna's tests discovered that ethanol based gasoline cannot be viewed as an option to 100LL avgas. The test likewise proposed that operational safety might be in jeopardy if usage of these fuels containing ethanol is continued. Cessna outlined a few problems in MOGAS; MOGAS needs fuel flow increase of 40% compared to AVGAS, MOGAS fuel is incompatible with some fuel system components, possible hazardous influence of electric fuel pumps by adding internal wear causing unexpected spark generation, MOGAS is incompatible with some fuel gauging systems and cause be able to incorrect fuel amount signs on the indicator, dissolve large amounts of water at conditions down to -77°F, impeding detection and removal of water from the fuel system, possible blockage of fuel filters and fuel flow and possible heavy losses from evaporation. This paper reviews concerns when using MOGAS in aircraft.

The silent-aircraft experimental aircraft are balanced by generating lift near the aircraft nose through leading-edge carving of the centerbody. The use of leading-edge carving over the centerbody is novel, in that previous blended-wing-body aircraft have balanced the aircraft by downloading the centerbody (via reflex camber) to achieve the effect of a tail. This paper decomposes the aerodynamic forces into contributions from spanwise sections to explain how three-dimensional flow effects are beneficial in allowing the silent-aircraft experimental aircraft to be both statically stable and to have an elliptical lift distribution over a large range of angles of attack. By analyzing the results in this manner, rationale is also given as to why, unlike other blended-wing-body-type configurations, the silent-aircraft-experimental design can use supercritical unstable-outer-wing airfoil profiles to generate a balanced and stable aircraft. The results are then used to develop a methodology to aid the aircraft designer in determining the amount of leading-edge carving that is necessary to achieve static stability for blended-wing-body-type aircraft.

T HIS Note presents a real-time flight trajectory optimization method. The flight trajectory that minimizes the fuel consumption or flight time of an aircraft can be solved with optimization. Because the optimization is time-consuming, the optimal flight trajectory needs to be obtained before flight. This, however, cannot cope with unexpected situations in flight. Thus, the real-time optimization, which optimizes the flight trajectory with the transition of the flight state, is important. The final goal of our study is to establish the real-time trajectory generation applicable to emergency landing approaches. Though the present flight management system provides the optimal flight path for a flight plan in normal operations, it cannot operate in emergency situations. This Note presents a fundamental real-time trajectory optimization algorithm and its flight validation. Many studies about real-time flight-path generation have produced a flight path by connecting the trim conditions in prestored databases [1,2]. In these studies, computer-assisted simulations were performed, but there were few actual flight experiments. Additionally, it is difficult to generate nonstational trajectories and the wind effects were not considered. In Japan, the Society of Japanese Aerospace Companies has promoted research on the development of a fault-tolerant flight control system. In a proceeding study, Suzuki et al. proposed a flight trajectory search method composed of a realtime A* algorithm and a random tabu search method [3]. They also conducted flight experiments. This random-based method, however, lacks convergent stability, so it often does not produce the appropriate trajectory. We have taken over this study, and in this Note we apply a direct collocation method including some schemes to solve the preceding problems. Actual flight experiments are also conducted to prove the effectiveness of the method. The experimental aircraft is MuPAL, the Multi-Purpose Aviation Laboratory, developed by JAXA (Japan Aerospace Exploration Agency) [4]. It has a high-precision Global Positioning System/inertial navigation system and a three-axis airspeed sensor. The system allows the estimation of the wind velocity and direction during flight. MuPALalso has a tunnel-in-the-sky display, which was developed by JAXA [5]. The experiments were conducted under manual tracking control using the display. In the future, the trajectory generated by the real-time optimization method will be tracked by an automaticflight control systembeing developed in the research on the fault-tolerant flight control system.

With the growing intolerance to failures within systems, the issue of fault diagnosis has become ever prevalent. Information concerning these possible failures can help to minimise the disruption to the functionality of the system by allowing quick rectification. Traditional approaches to fault diagnosis within engineering systems have focused on sequential testing procedures and real-time mechanisms. Both methods have been predominantly limited to single fault causes. Latest approaches also consider the issue of multiple faults in reflection to the characteristics of modern day systems designed for high reliability. In addition, a diagnostic capability is required in real time and for changeable system functionality. This paper focuses on two approaches which have been developed to cater for the demands of diagnosis within current engineering systems, namely application of the fault tree analysis technique and the method of digraphs. Both use a comparative approach to consider differences between actual system behaviour and that expected. The procedural guidelines are discussed for each method, with an experimental aircraft fuel system used to test and demonstrate the features of the techniques. The effectiveness of the approaches is compared and their future potential highlighted.

This paper considers three different methods for low-order anti-windup (AW) compensator design and compares their application to a realistic flight control problem, where the dominant actuator nonlinearity is aileron rate saturation. The compensator design methods all rigorously enforce at least local exponential stability via absolute stability results, but differ in both the construction of the AW compensator itself and the performance requirements used in the design. In particular, the paper compares a low order AW “optimal” design method in which the compensator poles and zeros are chosen by the designer and the accompanying gains synthesised optimally; a new low order method in which the optimisation procedure optimally chooses both the zeros and the gains; and a recently introduced classical design method where loopshaping is used to completely determine the gains and dynamics of the compensator. The methods are compared using a high-order nonlinear model of the lateral dynamics of an experimental aircraft, on which similar compensators have recently been flight tested.

The fuel consumption of future civil aircraft needs to be reduced because of the CO2 restrictions declared by the European Union. A consequent lightweight design and a new engine concept called counter-rotating open rotor are seen as key technologies in the attempt to reach this ambitious goals. Bearing in mind that counter-rotating open rotor engines emit very high sound pressures at low frequencies and that lightweight structures have a poor transmission loss in the lower frequency range, these key technologies raise new questions in regard to acoustic passenger comfort. One of the promising solutions for the reduction of sound pressure levels inside the aircraft cabin are active sound and vibration systems. So far, active concepts have rarely been investigated for a counter-rotating open rotor pressure excitation on complex airframe structures. Hence, the state of the art is augmented by the preliminary study presented in this paper. The study shows how an active vibration control system can influence the sound transmission of counter-rotating open rotor noise through a complex airframe structure into the cabin. Furthermore, open questions on the way towards the realisation of an active control system are addressed. In this phase, an active feedforward control system is investigated in a fully equipped Dornier 728 experimental prototype aircraft. In particular, the sound transmission through the airframe, the coupling of classical actuators (inertial and piezoelectric patch actuators) into the structure and the performance of the active vibration control system with different error sensors are investigated. It can be shown that the active control system achieves a reduction up to 5 dB at several counter-rotating open rotor frequencies but also that a better performance could be achieved through further optimisations.

This paper presents the experimental validation of an adaptive total energy control strategy to regulating the airspeed and the flight path angle of a fixed-wing aircraft. The control strategy is developed from a nonlinear description of the total energy control system approach. An open source computational fluid dynamics software is used to estimate the aerodynamic characteristics of the experimental aircraft. The stability properties are analyzed considering the linearized closed-loop dynamics around an equilibrium point.