WorldWideScience

Sample records for ailerons

  1. Calculation of transonic aileron buzz

    Science.gov (United States)

    Steger, J. L.; Bailey, H. E.

    1979-01-01

    An implicit finite-difference computer code that uses a two-layer algebraic eddy viscosity model and exact geometric specification of the airfoil has been used to simulate transonic aileron buzz. The calculated results, which were performed on both the Illiac IV parallel computer processor and the Control Data 7600 computer, are in essential agreement with the original expository wind-tunnel data taken in the Ames 16-Foot Wind Tunnel just after World War II. These results and a description of the pertinent numerical techniques are included.

  2. Nonclassical aileron buzz in transonic flow

    Science.gov (United States)

    Bendiksen, Oddvar O.

    1993-01-01

    A computational study of inviscid, transonic aileron and trailing-edge buzz instabilities is presented. A mixed Eulerian-Lagrangian formulation is used to model the fluid-structure system and to obtain a system of space-discretized equations that is time-marched to simulate the aeroelastic behavior of the wing-aileron system. Results obtained suggest that shock-induced separation may not be an essential driving force behind all buzz phenomena. Several examples are shown where the shock motion interacts with the aileron motion to extract energy from the flow. If the trailing-edge region is sufficiently flexible and the shocks are at the trailing edge, a trailing-edge buzz instability appears possible.

  3. Numerical design of an adaptive aileron

    Science.gov (United States)

    Amendola, Gianluca; Dimino, Ignazio; Concilio, Antonio; Magnifico, Marco; Pecora, Rosario

    2016-04-01

    The study herein described is aimed at investigating the feasibility of an innovative full-scale camber morphing aileron device. In the framework of the "Adaptive Aileron" project, an international cooperation between Italy and Canada, this goal was carried out with the integration of different morphing concepts in a wing-tip prototype. As widely demonstrated in recent European projects such as Clean Sky JTI and SARISTU, wing trailing edge morphing may lead to significant drag reduction (up to 6%) in off-design flight points by adapting chord-wise camber variations in cruise to compensate A/C weight reduction following fuel consumption. Those researches focused on the flap region as the most immediate solution to implement structural adaptations. However, there is also a growing interest in extending morphing functionalities to the aileron region preserving its main functionality in controlling aircraft directional stability. In fact, the external region of the wing seems to be the most effective in producing "lift over drag" improvements by morphing. Thus, the objective of the presented research is to achieve a certain drag reduction in off-design flight points by adapting wing shape and lift distribution following static deflections. In perspective, the developed device could also be used as a load alleviation system to reduce gust effects, augmenting its frequency bandwidth. In this paper, the preliminary design of the adaptive aileron is first presented, assessed on the base of the external aerodynamic loads. The primary structure is made of 5 segmented ribs, distributed along 4 bays, each splitted into three consecutive parts, connected with spanwise stringers. The aileron shape modification is then implemented by means of an actuation system, based on a classical quick-return mechanism, opportunely suited for the presented application. Finite element analyses were assessed for properly sizing the load-bearing structure and actuation systems and for

  4. Summary of NASA/DOE Aileron-Control Development Program for Wind Turbines

    Science.gov (United States)

    Miller, D. R.

    1986-01-01

    The development of aileron-control for wind turbines is discussed. Selected wind tunnel test results and full-scale rotor test results are presented for various types of ailerons. Finally, the current status of aileron-control development is discussed. Aileron-control was considered as a method of rotor control for use on wind turbines based on its potential to reduce rotor weight and cost. Following an initial feasibility study, a 20 percent chord aileron-control rotor was fabricated and tested on the NASA/DOE Mod-0 experimental wind turbine. Results from these tests indicated that the 20 percent chord ailerons regulated power and provided overspeed protection, but only over a very limited windspeed range. The next aileron-control rotor to be tested on the Mod-0 had 38 percent chord ailerons and test results showed these ailerons provided overspeed protection and power regulation over the Mod-0's entire operational windspeed range.

  5. Evaluation of aileron actuator reliability with censored data

    Directory of Open Access Journals (Sweden)

    Li Huaiyuan

    2015-08-01

    Full Text Available For the purpose of enhancing reliability of aileron of Airbus new-generation A350XWB, an evaluation of aileron reliability on the basis of maintenance data is presented in this paper. Practical maintenance data contains large number of censoring samples, information uncertainty of which makes it hard to evaluate reliability of aileron actuator. Considering that true lifetime of censoring sample has identical distribution with complete sample, if censoring sample is transformed into complete sample, conversion frequency of censoring sample can be estimated according to frequency of complete sample. On the one hand, standard life table estimation and product limit method are improved on the basis of such conversion frequency, enabling accurate estimation of various censoring samples. On the other hand, by taking such frequency as one of the weight factors and integrating variance of order statistics under standard distribution, weighted least square estimation is formed for accurately estimating various censoring samples. Large amounts of experiments and simulations show that reliabilities of improved life table and improved product limit method are closer to the true value and more conservative; moreover, weighted least square estimate (WLSE, with conversion frequency of censoring sample and variances of order statistics as the weights, can still estimate accurately with high proportion of censored data in samples. Algorithm in this paper has good effect and can accurately estimate the reliability of aileron actuator even with small sample and high censoring rate. This research has certain significance in theory and engineering practice.

  6. Aircraft wing structural detail design (wing, aileron, flaps, and subsystems)

    Science.gov (United States)

    Downs, Robert; Zable, Mike; Hughes, James; Heiser, Terry; Adrian, Kenneth

    1993-01-01

    The goal of this project was to design, in detail, the wing, flaps, and ailerons for a primary flight trainer. Integrated in this design are provisions for the fuel system, the electrical system, and the fuselage/cabin carry-through interface structure. This conceptual design displays the general arrangement of all major components in the wing structure, taking into consideration the requirements set forth by the appropriate sections of Federal Aviation Regulation Part 23 (FAR23) as well as those established in the statement of work.

  7. Reflection plane tests of a wind turbine blade tip section with ailerons

    Science.gov (United States)

    Savino, J. M.; Nyland, T. W.; Birchenough, A. G.; Jordan, F. L.; Campbell, N. K.

    1985-01-01

    Tests were conducted in the NASA Langley 30 by 60 foot Wind Tunnel on a full scale 7.31 m (24 ft) long tip section of a wind turbine rotor blade. The blade tip section was built with ailerons on the trailing edge. The ailerons, which spanned a length of 6.1 m (20 ft), were designed so that two types could be evaluated: the plain and the balanced. The ailerons were hinged on the suction surface at the 0.62 X chord station behind the leading edge. The purpose of the tests was to measure the aerodynamic characteristics of the blade section for: an angle of attack range from 0 deg to 90 deg aileron deflections from 0 deg to -90 deg, and Reynolds numbers of 0.79 and 1.5 x 10 to the 6th power. These data were then used to determine which aileron configuration had the most desirable rotor control and aerodynamic braking characteristics. Tests were also run to determine the effects of vortex generators, leading edge roughness, and the gaps between the aileron sections on the lift, drag, and chordwise force coefficients of the blade tip section.

  8. Experimental characterization of an adaptive aileron: lab tests and FE correlation

    Science.gov (United States)

    Amendola, Gianluca; Dimino, Ignazio; Amoroso, Francesco; Pecora, Rosario

    2016-04-01

    Like any other technology, morphing has to demonstrate system level performance benefits prior to implementation onto a real aircraft. The current status of morphing structures research efforts (as the ones, sponsored by the European Union) involves the design of several subsystems which have to be individually tested in order to consolidate their general performance in view of the final integration into a flyable device. This requires a fundamental understanding of the interaction between aerodynamic, structure and control systems. Important worldwide research collaborations were born in order to exchange acquired experience and better investigate innovative technologies devoted to morphing structures. The "Adaptive Aileron" project represents a joint cooperation between Canadian and Italian research centers and leading industries. In this framework, an overview of the design, manufacturing and testing of a variable camber aileron for a regional aircraft is presented. The key enabling technology for the presented morphing aileron is the actuation structural system, integrating a suitable motor and a load-bearing architecture. The paper describes the lab test campaign of the developed device. The implementation of a distributed actuation system fulfills the actual tendency of the aeronautical research to move toward the use of electrical power to supply non-propulsive systems. The aileron design features are validated by targeted experimental tests, demonstrating both its adaptive capability and robustness under operative loads and its dynamic behavior for further aeroelastic analyses. The experimental results show a satisfactory correlation with the numerical expectations thus validating the followed design approach.

  9. Performance and power regulation characteristics of two aileron-controlled rotors and a pitchable tip-controlled rotor on the Mod-O turbine

    Science.gov (United States)

    Corrigan, Robert D.; Ensworth, Clinton B. F., III; Miller, Dean R.

    1987-01-01

    Tests were conducted on the DOE/NASA mod-0 horizontal axis wind turbine to compare and evaluate the performance and the power regulation characteristics of two aileron-controlled rotors and a pitchable tip-controlled rotor. The two aileron-controlled rotor configurations used 20 and 38 percent chord ailerons, while the tip-controlled rotor had a pitchable blade tip. The ability of the control surfaces to regulate power was determined by measuring the change in power caused by an incremental change in the deflection angle of the control surface. The data shows that the change in power per degree of deflection angle for the tip-controlled rotor was four times the corresponding value for the 2- percent chord ailerons. The root mean square power deviation about a power setpoint was highest for the 20 percent chord aileron, and lowest for the 38 percent chord aileron.

  10. Failure Analysis of T-38 Aircraft Burst Hydraulic Aileron Return Line

    Science.gov (United States)

    Martinez, J. E.; Figert, J. D.; Paton, R. M.; Nguyen, S. D.; Flint, A.

    2012-01-01

    During maintenance troubleshooting for fluctuating hydraulic pressures, a technician found that a right hand aileron return line, on the flight hydraulic side, was ruptured (Fig. 1, 2). This tubing is part of the Hydraulic Flight Control Aileron Return Reducer to Aileron Manifold and is suspected to be original to the T-38 Talon trainer aircraft. Ailerons are small hinged sections on the outboard portion of a wing used to generate rolling motion thereby banking the aircraft. The ailerons work by changing the effective shape of the airfoil of the outer portion of the wing [1]. The drawing, Northrop P/N 3-43033-55 (6/1960), specifies that the line is made from 0.375 inch OD, aluminum 5052-0 tubing with a 0.049 inch wall thickness. WW-T-787 requires the tube shall be seamless and uniform in quality and temper [2]. The test pressure for this line is 3000 psi, and the operational pressure for this line is estimated to be between 45 psi and 1500 psi based on dynamic loading during flight. Examination of the fracture surface found evidence of arrest bands originating on the inner diameter (Fig 3). Ductile dimples are observed on the tube fractures (Fig. 4). The etched cross-section revealed thinning and work-hardening in the burst region (Fig. 5). The wall thickness just outside the work-hardened fracture region measured 0.035". Barlow's Formula: P = 2St/D, where P is burst pressure, S is allowable stress, t is wall thickness and D is the outer diameter of tube. Using the ultimate tensile strength of 28 ksi and a measured wall thickness of 0.035 inches at burst, P = 5.2 ksi (burst pressure). Using the yield of 13 ksi (YS) for aluminum 5052-0, plastic deformation will happen at P = 2.4 ksi suggesting plastic deformation occurred at a proof pressure of 3.0 ksi. Conclusion: The burst resulted from high stress, low-cycle fatigue. Evidence of arrest bands originating on the inner diameter. Fracture is predominately shear dimples, characteristic of high load ductile fractures

  11. Rapid non-contact inspection of composite ailerons using air-coupled ultrasound

    Science.gov (United States)

    Panda, Rabi Sankar; Karpenko, Oleksii; Udpa, Lalita; Haq, Mahmoodul; Rajagopal, Prabhu; Balasubramaniam, Krishnan

    2016-02-01

    This paper demonstrates an approach for rapid non-contact air-coupled ultrasonic inspection of composite ailerons with complex cross-sectional profile including thickness changes, curvature and the presence of a number of stiffeners. Low-frequency plate guided ultrasonic modes are used in B-scan mode for the measurements in pitch-catch mode. Appropriate probe holder angles suitable for generating and receiving lower order guided wave modes are discussed. Different embodiments of the pitch-catch tandem positions along and across stiffener and curved regions of the test sample enable a rapid test campaign capturing the feature-rich sample profile. Techniques to distinguish special features in the stiffener are presented.

  12. The feasibility analysis of aileron manpower control  in  a  fighter%某型歼击机副翼人力操纵可行性分析

    Institute of Scientific and Technical Information of China (English)

    王建平; 沈燕良; 曹克强

    2001-01-01

    In order to eliminate the suspense of aileron manpower control, the model of feasibility analysis of aileron manpower control was built. Make use of the flying mechanics base principle,this paper analyzed the required power when aileron hydromechanical servomechanism switch to manpower control because of the failure of normal hydraulic power supply in a fighter.Based on it, the article analysed the feasibility of aileron manpower control.The result is,take the suitable process, aileron manpower control is feasible.%为消除飞行人员对副翼人力操纵的疑虑,在建立副翼人力操纵可行性分析模型的基础上,应用飞行力学的基本原理,对某型歼击机因液压系统正常供压部分失效,副翼助力器转换为人力操纵后所需的压杆力进行了定量分析计算。结果表明,在应急情况下,只要处理得当,该型飞机副翼人力操纵是可行的。

  13. Wind tunnel research concerning lateral control devices, particularly at high angles of attack VII : Handley Page tip and full-span slots with ailerons and spoilers

    Science.gov (United States)

    Weick, Fred E; Wenzinger, Carl J

    1933-01-01

    Tests were made with ordinary ailerons and different sizes of spoilers on rectangular Clark Y wing models with Handley Page tip and full span slots. The tests showed the effect of the control devices on the general performance of the wings as well as on the lateral control and lateral stability characteristics.

  14. 带副翼三维机翼粘性绕流计算的嵌套网格方法%Chimera Grid Calculation of 3D Viscous Flow over Wing with Aileron

    Institute of Scientific and Technical Information of China (English)

    李孝伟; 乔志德

    2001-01-01

    严格考虑副翼端面和机翼切口流动而运用嵌套网格方法,生成了带副翼三维机翼的计算网格 。 流场计算采用雷诺平均Navier-Stokes方程和Johnson-King湍流模型。数值计算结果与实 验值吻合很好。%Failing to find details about chimera grid calculation in the open lecture, we researched and developed such details ourselves. In section 2 we discussed Johnson-King turbulence model in much detail and derivedeqs. (1) through (9 ) . This discussion and the nine equations formed the core of our research and dev elopment. With the chimera embedding technique, we divided the entire flowfield into small sub-regions with overlapping grids. In each sub-region, we, aided by eqs. (1) through (9) derived for the Johnson-King turbulence model, solved the Re ynolds-averaged Navier-Stokes equations by iteration; after each iteration, our algorithm provided automatic boundary information exchanges among different sub -regions. We used a certain delta wing as numerical example. Simulation results as shown in Fig.3, 4 and 5 show good agreement with experimental data.

  15. Morphing Wing-Tip Open Loop Controller and its Validation During Wind Tunnel Tests at the IAR-NRC

    Directory of Open Access Journals (Sweden)

    Mohamed Sadok GUEZGUEZ

    2016-09-01

    Full Text Available In this project, a wing tip of a real aircraft was designed and manufactured. This wing tip was composed of a wing and an aileron. The wing was equipped with a composite skin on its upper surface. This skin changed its shape (morphed by use of 4 electrical in-house developed actuators and 32 pressure sensors. These pressure sensors measure the pressures, and further the loads on the wing upper surface. Thus, the upper surface of the wing was morphed using these actuators with the aim to improve the aerodynamic performances of the wing-tip. Two types of ailerons were designed and manufactured: one aileron is rigid (non-morphed and one morphing aileron. This morphing aileron can change its shape also for the aerodynamic performances improvement. The morphing wing-tip internal structure is designed and manufactured, and is presented firstly in the paper. Then, the modern communication and control hardware are presented for the entire morphing wing tip equipped with actuators and sensors having the aim to morph the wing. The calibration procedure of the wing tip is further presented, followed by the open loop controller results obtained during wind tunnel tests. Various methodologies of open loop control are presented in this paper, and results obtained were obtained and validated experimentally through wind tunnel tests.

  16. 76 FR 41045 - Special Conditions; Gulfstream Aerospace LP (GALP) Model G250 Airplane, Design Roll-Maneuver...

    Science.gov (United States)

    2011-07-13

    ... flight control system that provides control through the pilot inputs to the flight computer. This novel... airplane is equipped with an electronic spoiler-control system and a mechanical aileron-control system that provide roll control of the aircraft through pilot inputs. An electronic control unit operates the......

  17. Selected advanced aerodynamics and active controls technology concepts development on a derivative B-747

    Science.gov (United States)

    1980-01-01

    The feasibility of applying wing tip extensions, winglets, and active control wing load alleviation to the Boeing 747 is investigated. Winglet aerodynamic design methods and high speed wind tunnel test results of winglets and of symmetrically deflected ailerons are presented. Structural resizing analyses to determine weight and aeroelastic twist increments for all the concepts and flutter model test results for the wing with winglets are included. Control law development, system mechanization/reliability studies, and aileron balance tab trade studies for active wing load alleviation systems are discussed. Results are presented in the form of incremental effects on L/D, structural weight, block fuel savings, stability and control, airplane price, and airline operating economics.

  18. Output feedback non-linear decoupled control synthesis and observer design for manoeuvring aircraft

    Science.gov (United States)

    Singh, S. N.; Schy, A. A.

    1980-01-01

    A study of the applicability of nonlinear decoupling theory to the design of control systems using output feedback for maneuvering aircraft is presented. The response variables chosen for decoupled control were angular velocity components along roll, pitch, and yaw axes, angle of attack (p), and angle of sideslip, using aileron, rudder, and elevator controls. An observer design for a class of nonlinear systems was presented and this method was used to estimate angle of attack and sideslip; an approximate observer was obtained by neglecting derivatives of p and aileron deflection angles and it was used in a simulation study. A simulation study showed that precise rapid combined lateral and longitudinal maneuvers can be performed; it was also demonstrated that a bank-angle-command outer loop could be designed for precise bank angles changes and simultaneous large lift maneuvers.

  19. Development of the Floating Centrifugal Pump by Use of Non Contact Magnetic Drive and Its Performance

    OpenAIRE

    Mitsuo Uno; Takaaki Masuzoe; Isamu Aotani; Shin Oba; Toshiaki Kanemoto

    2004-01-01

    This article focuses on the impeller construction, non contact driving method and performance of a newly developed shaftless floating pump with centrifugal impeller. The drive principle of the floating impeller pump used the magnet induction method similar to the levitation theory of the linear motor. In order to reduce the axial thrust by the pressure different between shroud and disk side, the balance hole and the aileron blade were installed in the floating impeller. Considering the above ...

  20. Comparison of tensile strength of different carbon fabric reinforced epoxy composites

    OpenAIRE

    Jane Maria Faulstich de Paiva; Sérgio Mayer; Mirabel Cerqueira Rezende

    2006-01-01

    Carbon fabric/epoxy composites are materials used in aeronautical industry to manufacture several components as flaps, aileron, landing-gear doors and others. To evaluate these materials become important to know their mechanical properties, for example, the tensile strength. Tensile tests are usually performed in aeronautical industry to determinate tensile property data for material specifications, quality assurance and structural analysis. For this work, it was manufactured four different l...

  1. Rudder Augmented Trajectory Correction for Unmanned Aerial Vehicles to Decrease Lateral Image Errors of Fixed Camera Payloads

    OpenAIRE

    Fisher, Thomas M.

    2016-01-01

    This thesis developed a Rudder Augmented Trajectory Correction (RATC) method for small unmanned aerial vehicles. The goal of this type of controller is to minimize the lateral image errors of body-fixed non-gimbaled cameras. This is achieved through both aggressive trajectory following and elimination of the roll angle present in current aileron only trajectory correction autopilots. The analytical derivation of the rudder augmented trajectory correction controller is presented. Using estimat...

  2. Wind tunnel tests of high-lift systems for advanced transports using high-aspect-ratio supercritical wings

    Science.gov (United States)

    Allen, J. B.; Oliver, W. R.; Spacht, L. A.

    1982-01-01

    The wind tunnel testing of an advanced technology high lift system for a wide body and a narrow body transport incorporating high aspect ratio supercritical wings is described. This testing has added to the very limited low speed high Reynolds number data base for this class or aircraft. The experimental results include the effects on low speed aerodynamic characteristics of various leading and trailing edge devices, nacelles and pylons, ailerons, and spoilers, and the effects of Mach and Reynolds numbers.

  3. Transonic high Reynolds number stability and control characteristics of a 0.015-scale remotely controlled elevon model (44-0) of the space shuttle orbiter tested in calspan 8-foot TWT (LA70)

    Science.gov (United States)

    Parrell, H.; Gamble, J. D.

    1977-01-01

    Transonic Wind Tunnel tests were run on a .015 scale model of the space shuttle orbiter vehicle in the 8-foot transonic wind tunnel. Purpose of the test program was to obtain basic shuttle aerodynamic data through a full range of elevon and aileron deflections, verification of data obtained at other facilities, and effects of Reynolds number. Tests were performed at Mach numbers from .35 to 1.20 and Reynolds numbers from 3,500,000 to 8,200,000 per foot. The high Reynolds number conditions (nominal 8,000,000/foot) were obtained using the ejector augmentation system. Angle of attack was varied from -2 to +20 degrees at sideslip angles of -2, 0, and +2 degrees. Sideslip was varied from -6 to +8 degrees at constant angles of attack from 0 to +20 degrees. Aileron settings were varied from -5 to +10 degrees at elevon deflections of -10, 0, and +10 degrees. Fixed aileron settings of 0 and 2 degrees in combination with various fixed elevon settings between -20 and +5 degrees were also run at varying angles of attack.

  4. Evaluating the Handling Qualities of Flight Control Systems Including Nonlinear Aircraft and System Dynamics

    Science.gov (United States)

    Lin, Raymond Chao

    The handling qualities evaluation of nonlinear aircraft systems is an area of concern in loss-of-control (LOC) prevention. The Get Transfer Function (GetTF) method was demonstrated for evaluating the handling qualities of flight control systems and aircraft containing nonlinearities. NASA's Generic Transport Model (GTM), a nonlinear model of a civilian jet transport aircraft, was evaluated. Using classical techniques, the stability, control, and augmentation (SCAS) systems were designed to control pitch rate, roll rate, and airspeed. Hess's structural pilot model was used to model pilot dynamics in pitch and roll-attitude tracking. The simulated task was simultaneous tracking of, both, pitch and roll attitudes. Eight cases were evaluated: 1) gain increase of pitch-attitude command signal, 2) gain increase of roll-attitude command signal, 3) gain reduction of elevator command signal, 4) backlash in elevator actuator, 5) combination 3 and 4 in elevator actuator, 6) gain reduction of aileron command signal, 7) backlash in aileron actuator, and 8) combination of 6 and 7 in aileron actuator. The GetTF method was used to estimate the transfer function approximating a linear relationship between the proprioceptive signal of the pilot model and the command input. The transfer function was then used to predict the handling qualities ratings (HQR) and pilot-induced oscillation ratings (PIOR). The HQR is based on the Cooper-Harper rating scale. In pitch-attitude tracking, the nominal aircraft is predicted to have Level 2* HQRpitch and 2 pitch tracking exercise was also conducted to validate the structural pilot model.

  5. Structural testing for static failure, flutter and other scary things

    Science.gov (United States)

    Ricketts, R. H.

    1983-01-01

    Ground test and flight test methods are described that may be used to highlight potential structural problems that occur on aircraft. Primary interest is focused on light-weight general aviation airplanes. The structural problems described include static strength failure, aileron reversal, static divergence, and flutter. An example of each of the problems is discussed to illustrate how the data acquired during the tests may be used to predict the occurrence of the structural problem. While some rules of thumb for the prediction of structural problems are given the report is not intended to be used explicitly as a structural analysis handbook.

  6. A ram-air-spoiler roll stabilization device for forward control cruciform missiles

    Science.gov (United States)

    Blair, A. B., Jr.; Sawyer, W. C.; Jackson, C. M., Jr.

    1978-01-01

    An experimental investigation has been made at supersonic Mach numbers to determine the feasibility of using a ram-air-spoiler roll control device on a typical canard control missile configuration. As a basis for roll control comparisons, conventional aileron controls on the tail fins were also tested. Results are presented which indicate that the addition of nacelles on the missile tail fins resulted in satisfactory roll control effectiveness and only small changes in basic missile stability. The ram-air-spoiler roll control effectiveness is relatively constant over the range of vehicle attitudes and Mach numbers investigated.

  7. Computational methods for unsteady transonic flows

    Science.gov (United States)

    Edwards, John W.; Thomas, J. L.

    1987-01-01

    Computational methods for unsteady transonic flows are surveyed with emphasis on prediction. Computational difficulty is discussed with respect to type of unsteady flow; attached, mixed (attached/separated) and separated. Significant early computations of shock motions, aileron buzz and periodic oscillations are discussed. The maturation of computational methods towards the capability of treating complete vehicles with reasonable computational resources is noted and a survey of recent comparisons with experimental results is compiled. The importance of mixed attached and separated flow modeling for aeroelastic analysis is discussed, and recent calculations of periodic aerodynamic oscillations for an 18 percent thick circular arc airfoil are given.

  8. Design and development of multi-lane smart electromechanical actuators

    CERN Document Server

    Annaz, Fawaz Yahya

    2014-01-01

    Design and Development of Multi-Lane Smart Electromechanical Actuators presents the design of electromechanical actuators in two types of architectures, namely, Torque Summed Architecture (TSA) and Velocity Summed Architecture, (VSA). It examines them in: * Hardware redundancy, where the architecture is made up of 3 or 4 lanes. * Digital Math Model redundancy, where a more compact two lanes architectures will be presented. The book starts with the very basic concepts and introduces the design process logically so that an understanding of the smart multi-lane systems that drive an aileron

  9. Results obtained during accelerated transonic tests of the Bell XS-1 airplane in flights to a Mach number of 0.92

    Science.gov (United States)

    Drake, Hubert M; Mclaughlin, Milton D; Goodman, Harold R

    1948-01-01

    Results are presented of tests up to a Mach number of 0.92 at altitudes around 30,000 feet. The data obtained show that the airplane can be flown to this Mach number above 30,000 feet. Longitudinal trim changes have been experienced but the forces involved have been small. The elevator effectiveness decreased about one-half with increase of Mach number from 0.70 to 0.87. Buffeting has been experienced in level flight but it has been mild and the associated tail loads have been small. No aileron buzz or other flutter phenomena have been noted.

  10. Exploring bird aerodynamics using radio-controlled models

    Energy Technology Data Exchange (ETDEWEB)

    Hoey, Robert G, E-mail: bobh@antelecom.ne [Air Force Flight Test Center, Edwards AFB, CA (United States)

    2010-12-15

    A series of radio-controlled glider models was constructed by duplicating the aerodynamic shape of soaring birds (raven, turkey vulture, seagull and pelican). Controlled tests were conducted to determine the level of longitudinal and lateral-directional static stability, and to identify the characteristics that allowed flight without a vertical tail. The use of tail-tilt for controlling small bank-angle changes, as observed in soaring birds, was verified. Subsequent tests, using wing-tip ailerons, inferred that birds use a three-dimensional flow pattern around the wing tip (wing tip vortices) to control adverse yaw and to create a small amount of forward thrust in gliding flight.

  11. Optimal Topology of Aircraft Rib and Spar Structures under Aeroelastic Loads

    Science.gov (United States)

    Stanford, Bret K.; Dunning, Peter D.

    2014-01-01

    Several topology optimization problems are conducted within the ribs and spars of a wing box. It is desired to locate the best position of lightening holes, truss/cross-bracing, etc. A variety of aeroelastic metrics are isolated for each of these problems: elastic wing compliance under trim loads and taxi loads, stress distribution, and crushing loads. Aileron effectiveness under a constant roll rate is considered, as are dynamic metrics: natural vibration frequency and flutter. This approach helps uncover the relationship between topology and aeroelasticity in subsonic transport wings, and can therefore aid in understanding the complex aircraft design process which must eventually consider all these metrics and load cases simultaneously.

  12. Prediction of the aeroelastic behavior An application to wind-tunnel models

    OpenAIRE

    Roucou, Mickaël

    2015-01-01

    The work of this paper has been done during a Master thesis at the ONERA and deals with the establish-ment of an aeroelastic state-space model and its application to two wind-tunnel models studied at the ONERA. The established model takes into account a control surface input and a gust perturbation. The generalized aerodynamic forces are approximated using Roger’s and Karpel’s methods and the inertia of the aileron is computed using a finite element model in Nastran. The software used during ...

  13. Flutter Analysis of a Morphing Wing Technology Demonstrator: Numerical Simulation and Wind Tunnel Testing

    Directory of Open Access Journals (Sweden)

    Andreea KOREANSCHI

    2016-03-01

    Full Text Available As part of a morphing wing technology project, the flutter analysis of two finite element models and the experimental results of a morphing wing demonstrator equipped with aileron are presented. The finite element models are representing a wing section situated at the tip of the wing; the first model corresponds to a traditional aluminium upper surface skin of constant thickness and the second model corresponds to a composite optimized upper surface skin for morphing capabilities. The two models were analyzed for flutter occurrence and effects on the aeroelastic behaviour of the wing were studied by replacing the aluminium upper surface skin of the wing with a specially developed composite version. The morphing wing model with composite upper surface was manufactured and fitted with three accelerometers to record the amplitudes and frequencies during tests at the subsonic wind tunnel facility at the National Research Council. The results presented showed that no aeroelastic phenomenon occurred at the speeds, angles of attack and aileron deflections studied in the wind tunnel and confirmed the prediction of the flutter analysis on the frequencies and modal displacements.

  14. Fuzzy Logic Decoupled Lateral Control for General Aviation Airplanes

    Science.gov (United States)

    Duerksen, Noel

    1997-01-01

    It has been hypothesized that a human pilot uses the same set of generic skills to control a wide variety of aircraft. If this is true, then it should be possible to construct an electronic controller which embodies this generic skill set such that it can successfully control different airplanes without being matched to a specific airplane. In an attempt to create such a system, a fuzzy logic controller was devised to control aileron or roll spoiler position. This controller was used to control bank angle for both a piston powered single engine aileron equipped airplane simulation and a business jet simulation which used spoilers for primary roll control. Overspeed, stall and overbank protection were incorporated in the form of expert systems supervisors and weighted fuzzy rules. It was found that by using the artificial intelligence techniques of fuzzy logic and expert systems, a generic lateral controller could be successfully used on two general aviation aircraft types that have very different characteristics. These controllers worked for both airplanes over their entire flight envelopes. The controllers for both airplanes were identical except for airplane specific limits (maximum allowable airspeed, throttle ]ever travel, etc.). This research validated the fact that the same fuzzy logic based controller can control two very different general aviation airplanes. It also developed the basic controller architecture and specific control parameters required for such a general controller.

  15. Estimated Benefits of Variable-Geometry Wing Camber Control for Transport Aircraft

    Science.gov (United States)

    Bolonkin, Alexander; Gilyard, Glenn B.

    1999-01-01

    Analytical benefits of variable-camber capability on subsonic transport aircraft are explored. Using aerodynamic performance models, including drag as a function of deflection angle for control surfaces of interest, optimal performance benefits of variable camber are calculated. Results demonstrate that if all wing trailing-edge surfaces are available for optimization, drag can be significantly reduced at most points within the flight envelope. The optimization approach developed and illustrated for flight uses variable camber for optimization of aerodynamic efficiency (maximizing the lift-to-drag ratio). Most transport aircraft have significant latent capability in this area. Wing camber control that can affect performance optimization for transport aircraft includes symmetric use of ailerons and flaps. In this paper, drag characteristics for aileron and flap deflections are computed based on analytical and wind-tunnel data. All calculations based on predictions for the subject aircraft and the optimal surface deflection are obtained by simple interpolation for given conditions. An algorithm is also presented for computation of optimal surface deflection for given conditions. Benefits of variable camber for a transport configuration using a simple trailing-edge control surface system can approach more than 10 percent, especially for nonstandard flight conditions. In the cruise regime, the benefit is 1-3 percent.

  16. Control system design for the MOD-5A 7.3 mW wind turbine generator

    Science.gov (United States)

    Barton, Robert S.; Hosp, Theodore J.; Schanzenbach, George P.

    1995-01-01

    This paper provides descriptions of the requirements analysis, hardware development and software development phases of the Control System design for the MOD-5A 7.3 mW Wind Turbine Generator. The system, designed by General Electric Company, Advanced Energy Programs Department, under contract DEN 3-153 with NASA Lewis Research Center and DOE, provides real time regulation of rotor speed by control of both generator torque and rotor torque. A variable speed generator system is used to provide both airgap torque control and reactive power control. The wind rotor is designed with segmented ailerons which are positioned to control blade torque. The central component of the control system, selected early in the design process, is a programmable controller used for sequencing, alarm monitoring, communication, and real time control. Development of requirements for use of aileron controlled blades and a variable speed generator required an analytical simulation that combined drivetrain, tower and blade elastic modes with wind disturbances and control behavior. An orderly two phase plan was used for controller software development. A microcomputer based turbine simulator was used to facilitate hardware and software integration and test.

  17. Root Locus Based Autopilot PID’s Parameters Tuning for a Flying Wing Unmanned Aerial Vehicle

    Directory of Open Access Journals (Sweden)

    Fendy Santoso

    2008-05-01

    Full Text Available This paper depicts the applications of classical root locus based PID control to the longitudinal flight dynamics of a Flying Wing Unmanned Aerial Vehicle, P15035, developed by Monash Aerobotics Research Group in the Department of Electrical and Computer Systems Engineering, Monash University, Australia. The challenge associated with our UAV is related to the fact that all of its motions and attitude variables are controlled by two independently actuated ailerons, namely elevons, as its primary control surfaces along with throttle, in contrast to most conventional aircraft which have rudder, aileron and elevator. The reason to choose PID control is mainly due to its simplicity and availability. Since our current autopilot, MP2028, only provides PID control law for its flight control, our design result can be implemented straight away for PID parameters’ tuning and practical flight controls. Simulations indicate that a well-tuned PID autopilot has successfully demonstrated acceptable closed loop performances for both pitch and altitude loops. In general, full PID control configuration is the recommended control mode to overcome the adverse impact of disturbances. Moreover, by utilising this control scheme, overshoots have been successfully suppressed into a certain reasonable level. Furthermore, it has been proven that exact pole-zero cancellations by employing Derivative control configuration in both pitch and altitude loop to eliminate the effects of integral action contributed by open loop transfer function of elevon-average-to- pitch as well as pitch- to- pitch- rate is impractical.

  18. Aircraft wing structure detail design

    Science.gov (United States)

    Sager, Garrett L.; Roberts, Ron; Mallon, Bob; Alameri, Mohamed; Steinbach, Bill

    1993-01-01

    The provisions of this project call for the design of the structure of the wing and carry-through structure for the Viper primary trainer, which is to be certified as a utility category trainer under FAR part 23. The specific items to be designed in this statement of work were Front Spar, Rear Spar, Aileron Structure, Wing Skin, and Fuselage Carry-through Structure. In the design of these parts, provisions for the fuel system, electrical system, and control routing were required. Also, the total weight of the entire wing planform could not exceed 216 lbs. Since this aircraft is to be used as a primary trainer, and the SOW requires a useful life of 107 cycles, it was decided that all of the principle stresses in the structural members would be kept below 10 ksi. The only drawback to this approach is a weight penalty.

  19. Simulation to Flight Test for a UAV Controls Testbed

    Science.gov (United States)

    Motter, Mark A.; Logan, Michael J.; French, Michael L.; Guerreiro, Nelson M.

    2006-01-01

    The NASA Flying Controls Testbed (FLiC) is a relatively small and inexpensive unmanned aerial vehicle developed specifically to test highly experimental flight control approaches. The most recent version of the FLiC is configured with 16 independent aileron segments, supports the implementation of C-coded experimental controllers, and is capable of fully autonomous flight from takeoff roll to landing, including flight test maneuvers. The test vehicle is basically a modified Army target drone, AN/FQM-117B, developed as part of a collaboration between the Aviation Applied Technology Directorate (AATD) at Fort Eustis, Virginia and NASA Langley Research Center. Several vehicles have been constructed and collectively have flown over 600 successful test flights, including a fully autonomous demonstration at the Association of Unmanned Vehicle Systems International (AUVSI) UAV Demo 2005. Simulations based on wind tunnel data are being used to further develop advanced controllers for implementation and flight test.

  20. Control Surface Fault Diagnosis for Small Autonomous Aircraft

    DEFF Research Database (Denmark)

    Hansen, Søren; Blanke, Mogens

    2011-01-01

    Small unmanned aerial vehicles require a large degree of fault-tolerance in order to fulfil their duties in an satisfactory way, both with respect to economy and safety in operation. Small aerial vehicles are commonly constructed without much redundancy in hardware, primarily for reasons of cost...... of distributions and change detection methods are employed to reach decisions about not-normal behaviour and it is shown how control surface faults can be diagnosed for a specific UAV without adding additional hardware to the platform. Only telemetry data from the aircraft is used together with a basic model...... of relations between signals within the aircraft. Frequency domain methods are shown to be robust in exploring relevant properties of the signals. The detection is shown to work on data from a real incident where an aileron gets stuck during launch of a UAV....

  1. Optimalisasi Proses Perakitan Pesawat Tanpa Awak dengan Metoda Critical Path Methods (CPM

    Directory of Open Access Journals (Sweden)

    Dendi Adi Saputra M

    2016-04-01

    Full Text Available The manufacturing of Unmanned Aerial Vehicles (UAV requires a design process that involves the design of aircraft’s components such as fuselage, wing, horizontal stabilizer, vertical stabilizer, ailerons, elevators, tail, and wing. The process takes a long time. Therefore, the distribution of structural works based on their characteristics and classifications by considering their design attributes and manufacturing processes is required. This research aims to find the optimal time and critical path of the assembly process of an UAV based on product work breakdown structure (PWBS and critical path method (CPM. The result reveals that the optimal assembly time is 139 minutes. Finally, the application of product-oriented structural work distribution and the optimization of the assembly activities involved in the critical path successfully minimize the duration of the assembly process.

  2. Development of the Floating Centrifugal Pump by Use of Non Contact Magnetic Drive and Its Performance

    Directory of Open Access Journals (Sweden)

    Mitsuo Uno

    2004-01-01

    Full Text Available This article focuses on the impeller construction, non contact driving method and performance of a newly developed shaftless floating pump with centrifugal impeller. The drive principle of the floating impeller pump used the magnet induction method similar to the levitation theory of the linear motor. In order to reduce the axial thrust by the pressure different between shroud and disk side, the balance hole and the aileron blade were installed in the floating impeller. Considering the above effect, floating of an impeller in a pump was realized. Moreover, the performance curves of a developed pump are in agreement with a general centrifugal pump, and the dimensionless characteristic curve also agrees under the different rotational speed due to no mechanical friction of the rotational part. Therefore, utility of a non contacting magnetic-drive style pump with the floating impeller was made clear.

  3. Composite NDE using full-field pulse-echo ultrasonic propagation imaging system

    Science.gov (United States)

    Hong, Seung-Chan; Lee, Jung-Ryul; Park, Jongwoon

    2016-04-01

    In this paper, a novel ultrasonic propagation imaging system, called a full-field pulse-echo ultrasonic propagation imaging (FF PE UPI) system is presented. The coincided laser beams for ultrasonic sensing and generation are scanned and pulse-echo mode laser ultrasounds are captured. This procedure makes it possible to generate full-field ultrasound in through-the-thickness direction as large as the scan area. The system nondestructively inspected targets with two-axis translation stages. Various structural inspection results in the form of full-field ultrasonic wave propagation videos are introduced, which are an aluminum honeycomb sandwich, ailerons and carbon fiber reinforced plastic (CFRP) honeycomb sandwich structures including various defects.

  4. Technology for vertical flight. 5. Flight control and autopilot; Helicopter kogaku no kiso to oyo. 5. soju sochi to jidoka

    Energy Technology Data Exchange (ETDEWEB)

    Ohashi, Y.; Yamada, H. [Mitsubishi Heavy Industries, Ltd., Tokyo (Japan)

    2000-03-05

    The paper explained a flight control of helicopter. Fundamental compositional elements of the flight control of helicopter are a pilot operating device, linkage, centering device/trimming gear, and actuator. The related device is an autopilot which is for controllability and reduction of work loads of pilot. In the fixed wing aircraft, the wing generating lift, engine giving thrust, and aileron/rudder/elevator in charge of control are playing each role. However, in helicopter, a rotor plays 3 roles: lift generation, going ahead, and control of fuselage. As to the control method, the control stick and pedal are operated in the fixed wing aircraft, and the cyclic stick and pedal are operated also in helicopter. In addition, another control stick, collective stick, is also operated. In this operation, lift of rotor increases/decreases to control the vertical movement of fuselage. (NEDO)

  5. Investigation of lateral-directional aerodynamic parameters identification method for fly-by-wire passenger airliners

    Institute of Scientific and Technical Information of China (English)

    Wu Zhao; Wang Lixin; Lin Jiaming; Ai Junqiang

    2014-01-01

    A new identification method is proposed to solve the problem of the influence on the loaded excitation signals brought by high feedback gain augmentation in lateral-directional aerody-namic parameters identification of fly-by-wire (FBW) passenger airliners. Taking for example an FBW passenger airliner model with directional relaxed-static-stability, through analysis of its signal energy distribution and airframe frequency response, a new method is proposed for signal type selec-tion, signal parameters design, and the appropriate frequency relationship between the aileron and rudder excitation signals. A simulation validation is presented of the FBW passenger airliner’s lat-eral-directional aerodynamic parameters identification. The validation result demonstrates that the designed signal can excite the lateral-directional motion mode of the FBW passenger airliner ade-quately and persistently. Meanwhile, the relative errors of aerodynamic parameters are less than 5%.

  6. Combined pitching and yawing motion of airplanes

    Science.gov (United States)

    Baranoff, A V; Hopf, L

    1931-01-01

    This report treats the following problems: The beginning of the investigated motions is always a setting of the lateral controls, i.e., the rudder or the ailerons. Now, the first interesting question is how the motion would proceed if these settings were kept unchanged for some time; and particularly, what upward motion would set in, how soon, and for how long, since therein lie the dangers of yawing. Two different motions ensue with a high rate of turn and a steep down slope of flight path in both but a marked difference in angle of attack and consequently different character in the resultant aerodynamic forces: one, the "corkscrew" dive at normal angle, and the other, the "spin" at high angle.

  7. Comparison of concurrent strain gage- and pressure transducer-measured flight loads on a lifting reentry vehicle and correlation with wind tunnel predictions

    Science.gov (United States)

    Tang, M. H.; Sefic, W. J.; Sheldon, R. G.

    1978-01-01

    Concurrent strain gage and pressure transducer measured flight loads on a lifting reentry vehicle are compared and correlated with wind tunnel-predicted loads. Subsonic, transonic, and supersonic aerodynamic loads are presented for the left fin and control surfaces of the X-24B lifting reentry vehicle. Typical left fin pressure distributions are shown. The effects of variations in angle of attack, angle of sideslip, and Mach number on the left fin loads and rudder hinge moments are presented in coefficient form. Also presented are the effects of variations in angle of attack and Mach number on the upper flap, lower flap, and aileron hinge-moment coefficients. The effects of variations in lower flap hinge moments due to changes in lower flap deflection and Mach number are presented in terms of coefficient slopes.

  8. Stability and Control Characteristics of a Complete Airplane Model Having a Wing with Quarter-chord Line Swept Back 40 Degrees, Aspect Ratio 2.50, and Taper Ratio 0.42

    Science.gov (United States)

    Schulderfrei, Marvin; Comisarow, Paul; Goodson, Kenneth W

    1951-01-01

    An investigation has been made of a complete airplane model having a wing with the quarter-chord line swept back 40 degrees, aspect ratio 2.50, and taper ratio 0.42 to determine its low-speed stability and control characteristics. The longitudinal stability investigation included stabilizer and tail-off tests with different wing dihedral angles (Gamma = 0 degrees and Gamma = -10 degrees) over an angle-of-attack range for the cruising and landing configurations and tests. with a high horizontal-tail location (Gamma = -10 degrees) for the cruising configuration. Tests were made of the wing alone and to determine the effect of wing end plates in pitch. Lateral stability characteristics were determined for the airplane with different geometric wing dihedrals, with end plates, and with several dorsal modifications. Tests were made with ailerons and spoilers to determine control characteristics.

  9. Variable Structure Control of a Hand-Launched Glider

    Science.gov (United States)

    Anderson, Mark R.; Waszak, Martin R.

    2005-01-01

    Variable structure control system design methods are applied to the problem of aircraft spin recovery. A variable structure control law typically has two phases of operation. The reaching mode phase uses a nonlinear relay control strategy to drive the system trajectory to a pre-defined switching surface within the motion state space. The sliding mode phase involves motion along the surface as the system moves toward an equilibrium or critical point. Analysis results presented in this paper reveal that the conventional method for spin recovery can be interpreted as a variable structure controller with a switching surface defined at zero yaw rate. Application of Lyapunov stability methods show that deflecting the ailerons in the direction of the spin helps to insure that this switching surface is stable. Flight test results, obtained using an instrumented hand-launched glider, are used to verify stability of the reaching mode dynamics.

  10. Theoretical stability and control characteristics of wings with various amounts of taper and twist

    Science.gov (United States)

    Pearson, H. A.; Jones, R. T.

    1976-01-01

    Stability derivatives have been computed for twisted wings of different planforms that include variations in both the wing taper and the aspect ratio. Taper ratios of 1.0, 0.50, and 0.25 are considered for each of three aspect ratios: 6, 10, and 16. The specific derivatives for which results are given are the rolling moment and the yawing moment derivatives with respect to rolling velocity, yawing velocity, and angle of sideslip. In addition to the stability derivatives, results are included for determining the theoretical rolling moment due to aileron deflection and a series of influence lines is given by which the loading across the span may be determined for any angle-of-attack distribution that may occur on the wing planforms considered.

  11. Nonlinear Dynamic Modeling of a Fixed-Wing Unmanned Aerial Vehicle: a Case Study of Wulung

    Directory of Open Access Journals (Sweden)

    Fadjar Rahino Triputra

    2015-07-01

    Full Text Available Developing a nonlinear adaptive control system for a fixed-wing unmanned aerial vehicle (UAV requires a mathematical representation of the system dynamics analytically as a set of differential equations in the form of a strict-feedback systems. This paper presents a method for modeling a nonlinear flight dynamics of the fixed-wing UAV of BPPT Wulung in any conditions of the flight altitude and airspeed for the first step into designing a nonlinear adaptive controller. The model was formed into 10-DOF differential equations in the form of strict-feedback systems which separates the terms of elevator, aileron, rudder and throttle from the model. The model simulation results show the behavior of the flight dynamics of the Wulung UAV and also prove the compliance with the actual flight test results.

  12. Autonomous Flying Controls Testbed

    Science.gov (United States)

    Motter, Mark A.

    2005-01-01

    The Flying Controls Testbed (FLiC) is a relatively small and inexpensive unmanned aerial vehicle developed specifically to test highly experimental flight control approaches. The most recent version of the FLiC is configured with 16 independent aileron segments, supports the implementation of C-coded experimental controllers, and is capable of fully autonomous flight from takeoff roll to landing, including flight test maneuvers. The test vehicle is basically a modified Army target drone, AN/FQM-117B, developed as part of a collaboration between the Aviation Applied Technology Directorate (AATD) at Fort Eustis,Virginia and NASA Langley Research Center. Several vehicles have been constructed and collectively have flown over 600 successful test flights.

  13. Gust response modeling and alleviation scheme design for an elastic aircraft

    Institute of Scientific and Technical Information of China (English)

    2010-01-01

    Time-domain approaches are presented for analysis of the dynamic response of aeroservoelastic systems to atmospheric gust excitations. The continuous and discrete gust inputs are defined in the time domain. The time-domain approach to continuous gust response uses a state-space formulation that requires the frequency-dependent aerodynamic coefficients to be approximated with the rational function of a Laplace variable. A hybrid method which combines the Fourier transform and time-domain approaches is used to calculate discrete gust response. The purpose of this approach is to obtain a time-domain state-space model without using rational function approximation of the gust columns. Three control schemes are designed for gust alleviation on an elastic aircraft, and three control surfaces are used: aileron, elevator and spoiler. The signals from the rate of pitch angle gyroscope or angle of attack sensor are sent to the elevator while the signals from accelerometers at the wing tip and center of gravity of the aircraft are sent to the aileron and spoiler, respectively. All the control laws are based on classical control theory. The results show that acceleration at the center of gravity of the aircraft and bending-moment at the wing-root section are mainly excited by rigid modes of the aircraft and the accelerations at the wing-tip are mainly excited by elastic modes of the aircraft. All the three control schemes can be used to alleviate the wing-root moments and the accelerations. The gust response can be alleviated using control scheme 3, in which the spoiler is used as a control surface, but the effects are not as good as those of control schemes 1 and 2.

  14. 飞翼布局无人机舵面分配算法研究%Effectors allocation of aerial refueling UAV

    Institute of Scientific and Technical Information of China (English)

    楼静梅; 张科

    2012-01-01

    A method for effectors' privilege dynamic allocation is proposed for effectors dilapidation during UAV aerial refueling. Based on the flying wing UAV given by the reference, the performances of the classic effectors allocation and the dynamic effectors allocation are analyzed for the cases of normal and the ailerons dilapidation. Effectors work on the rails, and the pitch control moment and the roll control moment meet the restraints, the attainable ranges of the pitch moment and the roll moment produced by the effectors dynamic allocation are greater than those produced by the classic effectors allocation. Thus, the biggish disturbances in the pitch channel or the roll channel can be controlled. The dynamic effectors allocation arithmetic has strong fault-tolerant capacity with elevators or ailerons dilapidation.%针对某飞翼式无人机飞行过程中出现舵面破损情况,提出了一种具有容错能力的舵面权限动态分配算法.针对舵面无故障和副翼两侧完全破损故障,分析比较了典型舵面分配算法和动态分配算法的性能.分析结果表明:舵面无故障时,在满足约束条件的情况下,由动态分配算法实现的舵面分配其滚转和俯仰操纵力矩可达范围大于由典型舵面分配算法实现的,有利于克服较大的俯仰或滚转干扰,而在升降舵或者副翼出现破损故障时具有更强的容错能力.

  15. The Nolans Bore rare-earth element-phosphorus-uranium mineral system: geology, origin and post-depositional modifications

    Science.gov (United States)

    Huston, David L.; Maas, Roland; Cross, Andrew; Hussey, Kelvin J.; Mernagh, Terrence P.; Fraser, Geoff; Champion, David C.

    2016-08-01

    Nolans Bore is a rare-earth element (REE)-U-P fluorapatite vein deposit hosted mostly by the ~1805 Ma Boothby Orthogneiss in the Aileron Province, Northern Territory, Australia. The fluorapatite veins are complex, with two stages: (1) massive to granular fluorapatite with inclusions of REE silicates, phosphates and (fluoro)carbonates, and (2) calcite-allanite with accessory REE-bearing phosphate and (fluoro)carbonate minerals that vein and brecciate the earlier stage. The veins are locally accompanied by narrow skarn-like (garnet-diopside-amphibole) wall rock alteration zones. SHRIMP Th-Pb analyses of allanite yielded an age of 1525 ± 18 Ma, interpreted as the minimum age of mineralisation. The maximum age is provided by a ~1550 Ma SHRIMP U-Pb age for a pegmatite that predates the fluorapatite veins. Other isotopic systems yielded ages from ~1443 to ~345 Ma, implying significant post-depositional isotopic disturbance. Calculation of initial ɛNd and 87Sr/86Sr at 1525 Ma and stable isotope data are consistent with an enriched mantle or lower crust source, although post-depositional disturbance is likely. Processes leading to formation of Nolans Bore began with north-dipping subduction along the south margin of the Aileron Province at 1820-1750 Ma, producing a metasomatised, volatile-rich, lithospheric mantle wedge. About 200 million years later, near the end of the Chewings Orogeny, this reservoir and/or the lower crust sourced alkaline low-degree partial melts which passed into the mid- and upper-crust. Fluids derived from these melts, which may have included phosphatic melts, eventually deposited the Nolans Bore fluorapatite veins due to fluid-rock interaction, cooling, depressurisation and/or fluid mixing. Owing to its size and high concentration of Th (2500 ppm), in situ radiogenic heating caused significant recrystallisation and isotopic resetting. The system finally cooled below 300 °C at ~370 Ma, possibly in response to unroofing during the Alice Springs

  16. Design and verification of a smart wing for an extreme-agility micro-air-vehicle

    International Nuclear Information System (INIS)

    A special class of fixed-wing micro-air-vehicle (MAV) is currently being designed to fly and hover to provide range superiority as well as being able to hover through a flight maneuver known as prop-hanging to accomplish a variety of surveillance missions. The hover maneuver requires roll control of the wing through differential aileron deflection but a conventional system contributes significantly to the gross weight and complexity of a MAV. Therefore, it is advantageous to use smart structure approaches with active materials to design a lightweight, robust wing for the MAV. The proposed smart wing consists of an active trailing edge flap integrated with bimorph actuators with piezoceramic fibers. Actuation is enhanced by preloading the bimorph actuators with a compressive axial load. The preload is exerted on the actuators through a passive latex or electroactive polymer (EAP) skin that wraps around the airfoil. An EAP skin would further enhance the actuation by providing an electrostatic effect of the dielectric polymer to increase the deflection. Analytical modeling as well as finite element analysis show that the proposed concept could achieve the target bi-directional deflection of 30° in typical flight conditions. Several bimorph actuators were manufactured and an experimental setup was designed to measure the static and dynamic deflections. The experimental results validated the analytical technique and finite element models, which have been further used to predict the performance of the smart wing design for a MAV

  17. Optimization Design System for Composite Structures Based on Grid Technology

    Institute of Scientific and Technical Information of China (English)

    CHENG Wen-yuan; CHANG Yan; CUI De-gang; XIE Xiang-hui

    2007-01-01

    To solve the topology optimization of complicated multi-objective continuous/discrete design variables in aircraft structure design, a Parallel Pareto Genetic Algorithm (PPGA) is presented based on grid platform in this paper. In the algorithm, the commercial finite element analysis (FEA) software is integrated as the calculating tool for analyzing the objective functions and the filter of Pareto solution set based on weight information is introduced to deal with the relationships among all objectives. Grid technology is utilized in PPGA to realize the distributed computations and the user interface is developed to realize the job submission and job management locally/remotely. Taking the aero-elastic tailoring of a composite wing for optimization as an example, a set of Pareto solutions are obtained for the decision-maker. The numerical results show that the aileron reversal problem can be solved by adding the limited skin weight in this system. The algorithm can be used to solve complicated topology optimization for composite structures in engineering and the computation efficiency can be improved greatly by using the grid platform that aggregates numerous idle resources.

  18. NDE of Damage in Aircraft Flight Control Surfaces

    Science.gov (United States)

    Hsu, David K.; Barnard, Daniel J.; Dayal, Vinay

    2007-03-01

    Flight control surfaces on an aircraft, such as ailerons, flaps, spoilers and rudders, are typically adhesively bonded composite or aluminum honeycomb sandwich structures. These components can suffer from damage caused by hail stone, runway debris, or dropped tools during maintenance. On composites, low velocity impact damages can escape visual inspection, whereas on aluminum honeycomb sandwich, budding failure of the honeycomb core may or may not be accompanied by a disbond. This paper reports a study of the damage morphology in such structures and the NDE methods for detecting and characterizing them. Impact damages or overload failures in composite sandwiches with Nomex or fiberglass core tend to be a fracture or crinkle or the honeycomb cell wall located a distance below the facesheet-to-core bondline. The damage in aluminum honeycomb is usually a buckling failure, propagating from the top skin downward. The NDE methods used in this work for mapping out these damages were: air-coupled ultrasonic scan, and imaging by computer aided tap tester. Representative results obtained from the field will be shown.

  19. Airplane Upset Training Evaluation Report

    Science.gov (United States)

    Gawron, Valerie J.; Jones, Patricia M. (Technical Monitor)

    2002-01-01

    Airplane upset accidents are a leading factor in hull losses and fatalities. This study compared five types of airplane-upset training. Each group was composed of eight, non-military pilots flying in their probationary year for airlines operating in the United States. The first group, 'No aero / no upset,' was made up of pilots without any airplane upset training or aerobatic flight experience; the second group, 'Aero/no upset,' of pilots without any airplane-upset training but with aerobatic experience; the third group, 'No aero/upset,' of pilots who had received airplane-upset training in both ground school and in the simulator; the fourth group, 'Aero/upset,' received the same training as Group Three but in addition had aerobatic flight experience; and the fifth group, 'In-flight' received in-flight airplane upset training using an instrumented in-flight simulator. Recovery performance indicated that clearly training works - specifically, all 40 pilots recovered from the windshear upset. However few pilots were trained or understood the use of bank to change the direction of the lift vector to recover from nose high upsets. Further, very few thought of, or used differential thrust to recover from rudder or aileron induced roll upsets. In addition, recovery from icing-induced stalls was inadequate.

  20. Peak-Seeking Control For Reduced Fuel Consumption: Flight-Test Results For The Full-Scale Advanced Systems Testbed FA-18 Airplane

    Science.gov (United States)

    Brown, Nelson

    2013-01-01

    A peak-seeking control algorithm for real-time trim optimization for reduced fuel consumption has been developed by researchers at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center to address the goals of the NASA Environmentally Responsible Aviation project to reduce fuel burn and emissions. The peak-seeking control algorithm is based on a steepest-descent algorithm using a time-varying Kalman filter to estimate the gradient of a performance function of fuel flow versus control surface positions. In real-time operation, deflections of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of an F/A-18 airplane are used for optimization of fuel flow. Results from six research flights are presented herein. The optimization algorithm found a trim configuration that required approximately 3 percent less fuel flow than the baseline trim at the same flight condition. This presentation also focuses on the design of the flight experiment and the practical challenges of conducting the experiment.

  1. Flight Test of the F/A-18 Active Aeroelastic Wing Airplane

    Science.gov (United States)

    Voracek, David

    2007-01-01

    A viewgraph presentation of flight tests performed on the F/A active aeroelastic wing airplane is shown. The topics include: 1) F/A-18 AAW Airplane; 2) F/A-18 AAW Control Surfaces; 3) Flight Test Background; 4) Roll Control Effectiveness Regions; 5) AAW Design Test Points; 6) AAW Phase I Test Maneuvers; 7) OBES Pitch Doublets; 8) OBES Roll Doublets; 9) AAW Aileron Flexibility; 10) Phase I - Lessons Learned; 11) Control Law Development and Verification & Validation Testing; 12) AAW Phase II RFCS Envelopes; 13) AAW 1-g Phase II Flight Test; 14) Region I - Subsonic 1-g Rolls; 15) Region I - Subsonic 1-g 360 Roll; 16) Region II - Supersonic 1-g Rolls; 17) Region II - Supersonic 1-g 360 Roll; 18) Region III - Subsonic 1-g Rolls; 19) Roll Axis HOS/LOS Comparison Region II - Supersonic (open-loop); 20) Roll Axis HOS/LOS Comparison Region II - Supersonic (closed-loop); 21) AAW Phase II Elevated-g Flight Test; 22) Region I - Subsonic 4-g RPO; and 23) Phase II - Lessons Learned

  2. DARPA/AFRL/NASA Smart Wing Second Wind Tunnel Test Results

    Science.gov (United States)

    Scherer, L. B.; Martin, C. A.; West, M.; Florance, J. P.; Wieseman, C. D.; Burner, A. W.; Fleming, G. A.

    2001-01-01

    To quantify the benefits of smart materials and structures adaptive wing technology, Northrop Grumman Corp. (NGC) built and tested two 16% scale wind tunnel models (a conventional and a "smart" model) of a fighter/attack aircraft under the DARPA/AFRL/NASA Smart Materials and Structures Development - Smart Wing Phase 1. Performance gains quantified included increased pitching moment (C(sub M)), increased rolling moment (C(subl)) and improved pressure distribution. The benefits were obtained for hingeless, contoured trailing edge control surfaces with embedded shape memory alloy (SMA) wires and spanwise wing twist effected by SMA torque tube mechanisms, compared to conventional hinged control surfaces. This paper presents an overview of the results from the second wind tunnel test performed at the NASA Langley Research Center s (LaRC) 16ft Transonic Dynamic Tunnel (TDT) in June 1998. Successful results obtained were: 1) 5 degrees of spanwise twist and 8-12% increase in rolling moment utilizing a single SMA torque tube, 2) 12 degrees of deflection, and 10% increase in rolling moment due to hingeless, contoured aileron, and 3) demonstration of optical techniques for measuring spanwise twist and deflected shape.

  3. Design and verification of a smart wing for an extreme-agility micro-air-vehicle

    Science.gov (United States)

    Wickramasinghe, Viresh; Chen, Yong; Martinez, Marcias; Wong, Franklin; Kernaghan, Robert

    2011-12-01

    A special class of fixed-wing micro-air-vehicle (MAV) is currently being designed to fly and hover to provide range superiority as well as being able to hover through a flight maneuver known as prop-hanging to accomplish a variety of surveillance missions. The hover maneuver requires roll control of the wing through differential aileron deflection but a conventional system contributes significantly to the gross weight and complexity of a MAV. Therefore, it is advantageous to use smart structure approaches with active materials to design a lightweight, robust wing for the MAV. The proposed smart wing consists of an active trailing edge flap integrated with bimorph actuators with piezoceramic fibers. Actuation is enhanced by preloading the bimorph actuators with a compressive axial load. The preload is exerted on the actuators through a passive latex or electroactive polymer (EAP) skin that wraps around the airfoil. An EAP skin would further enhance the actuation by providing an electrostatic effect of the dielectric polymer to increase the deflection. Analytical modeling as well as finite element analysis show that the proposed concept could achieve the target bi-directional deflection of 30° in typical flight conditions. Several bimorph actuators were manufactured and an experimental setup was designed to measure the static and dynamic deflections. The experimental results validated the analytical technique and finite element models, which have been further used to predict the performance of the smart wing design for a MAV.

  4. Flaperon Modification Effect on Jet-Flap Interaction Noise Reduction for Chevron Nozzles

    Science.gov (United States)

    Thomas, Russell H.; Mengle, Vinod G.; Stoker, Robert W.; Brusniak, Leon; Elkoby, Ronen

    2007-01-01

    Jet-flap interaction (JFI) noise can become an important component of far field noise when a flap is immersed in the engine propulsive stream or is in its entrained region, as in approach conditions for under-the-wing engine configurations. We experimentally study the effect of modifying the flaperon, which is a high speed aileron between the inboard and outboard flaps, at both approach and take-off conditions using scaled models in a free jet. The flaperon modifications were of two types: sawtooth trailing edge and mini vortex generators (vg s). Parametric variations of these two concepts were tested with a round coaxial nozzle and an advanced chevron nozzle, with azimuthally varying fan chevrons, using both far field microphone arrays and phased microphone arrays for source diagnostics purposes. In general, the phased array results corroborated the far field results in the upstream quadrant pointing to JFI near the flaperon trailing edge as the origin of the far field noise changes. Specific sawtooth trailing edges in conjunction with the round nozzle gave marginal reduction in JFI noise at approach, and parallel co-rotating mini-vg s were somewhat more beneficial over a wider range of angles, but both concepts were noisier at take-off conditions. These two concepts had generally an adverse JFI effect when used in conjunction with the advanced chevron nozzle at both approach and take-off conditions.

  5. Performance, Stability, and Control Investigation at Mach Numbers from 0.4 to 0.9 of a Model of the "Swallow" with Outer Wing Panels Swept 25 degree with and without Power Simulation

    Science.gov (United States)

    Runckel, Jack F.; Schmeer, James W.; Cassetti, Marlowe D.

    1960-01-01

    An investigation of the performance, stability, and control characteristics of a variable-sweep arrow-wing model (the "Swallow") with the outer wing panels swept 25 deg has been conducted in the Langley 16-foot transonic tunnel. The wing was uncambered and untwisted and had RAE 102 airfoil sections with a thickness-to-chord ratio of 0.14 normal to the leading edge. Four outboard engines located above and below the wing provided propulsive thrust, and, by deflecting in the pitch direction and rotating in the lateral plane, also produced control forces. A pair of swept lateral fins and a single vertical fin were mounted on each engine nacelle to provide aerodynamic stability and control. Jets-off data were obtained with flow-through nacelles, stimulating the effects of inlet flow; jet thrust and hot-jet interference effects were obtained with faired-nose nacelles housing hydrogen peroxide gas generators. Six-component force and moment data were obtained through a Mach number range of 0.40 to 0.90 at angles of attack and angles of sideslip from 0 deg to 15 deg. Longitudinal, directional, and lateral control were obtained by deflecting the nacelle-fin combinations as elevators, rudders, and ailerons at several fixed angles for each control.

  6. Design for the Control of a Rotatable Stabiliser

    CERN Document Server

    Childs, S J

    2014-01-01

    This research sets out a design for the control of a rotatable stabiliser which, it is proposed, might augment, or fully replace, the conventional control mechanisms for pitch and yaw in aircraft. The anticipated advantages of such a device are around 25% less drag, for a capability which ranges between equivalent and greater than twofold that of the conventional tail. One, anticipated handicap of such a device is the potential for it to stall, from its tips, inward, if rotated too fast. For succinctness, a mapping betweeen states of the device and the position of a two-axis controller (e.g. a joystick) is formulated. The function of the joystick traditionally assigned to the control of ailerons is replaced by that traditionally associated with the rudder pedals. Its function is otherwise conventional. From this topology it follows that small and continuous adjustments of the controls should cause the stabiliser to rotate in a direction opposite to that of the joystick (when viewed from aft) and the deflectio...

  7. Squid rocket science: How squid launch into air

    Science.gov (United States)

    O'Dor, Ron; Stewart, Julia; Gilly, William; Payne, John; Borges, Teresa Cerveira; Thys, Tierney

    2013-10-01

    Squid not only swim, they can also fly like rockets, accelerating through the air by forcefully expelling water out of their mantles. Using available lab and field data from four squid species, Sthenoteuthis pteropus, Dosidicus gigas, Illex illecebrosus and Loligo opalescens, including sixteen remarkable photographs of flying S. pteropus off the coast of Brazil, we compared the cost of transport in both water and air and discussed methods of maximizing power output through funnel and mantle constriction. Additionally we found that fin flaps develop at approximately the same size range as flight behaviors in these squids, consistent with previous hypotheses that flaps could function as ailerons whilst aloft. S. pteropus acceleration in air (265 body lengths [BL]/s2; 24.5m/s2) was found to exceed that in water (79BL/s2) three-fold based on estimated mantle length from still photos. Velocities in air (37BL/s; 3.4m/s) exceed those in water (11BL/s) almost four-fold. Given the obvious advantages of this extreme mode of transport, squid flight may in fact be more common than previously thought and potentially employed to reduce migration cost in addition to predation avoidance. Clearly squid flight, the role of fin flaps and funnel, and the energetic benefits are worthy of extended investigation.

  8. Sonic anemometry measurements to determine airflow patterns in multi-tunnel greenhouse

    Energy Technology Data Exchange (ETDEWEB)

    Lopez, A.; Valera, D. L.; Molina-aiz, F. D.; Pena, A.

    2012-11-01

    The present work describes a methodology for studying natural ventilation in Mediterranean greenhouses using sonic anemometry. The experimental work took place in the three-span greenhouse located at the agricultural research farm belonging to the University of Almeria. This methodology has allowed us to obtain patterns of natural ventilation of the experimental greenhouse under the most common wind regimes for this region. It has also enabled us to describe how the wind and thermal effects interact in the natural ventilation of the greenhouse, as well as to detect deficiencies in the ventilation of the greenhouse, caused by the barrier effect of the adjacent greenhouse (imply a mean reduction in air velocity close to the greenhouse when facing windward of 98% for u, 63% for u, and more importantly 88% for ux, the component of air velocity that is perpendicular to the side vent). Their knowledge allows us to improve the current control algorithms that manage the movement of the vents. In this work we make a series of proposals that could substantially improve the natural ventilation of the experimental greenhouse. For instance, install vents equipped with ailerons which guide the air inside, or with vents in which the screen is not placed directly over the side surface of the greenhouse. A different proposal is to prolong the opening of the side vents down to the soil, thus fomenting the entrance of air at crop level. (Author) 34 refs.

  9. Comparison of tensile strength of different carbon fabric reinforced epoxy composites

    Directory of Open Access Journals (Sweden)

    Jane Maria Faulstich de Paiva

    2006-03-01

    Full Text Available Carbon fabric/epoxy composites are materials used in aeronautical industry to manufacture several components as flaps, aileron, landing-gear doors and others. To evaluate these materials become important to know their mechanical properties, for example, the tensile strength. Tensile tests are usually performed in aeronautical industry to determinate tensile property data for material specifications, quality assurance and structural analysis. For this work, it was manufactured four different laminate families (F155/PW, F155/HS, F584/PW and F584/HS using pre-impregnated materials (prepregs based on F155TM and F584TM epoxy resins reinforced with carbon fiber fabric styles Plain Weave (PW and Eight Harness Satin (8HS. The matrix F155TM code is an epoxy resin type DGEBA (diglycidil ether of bisphenol A that contains a curing agent and the F584TM code is a modified epoxy resin type. The laminates were obtained by handing lay-up process following an appropriate curing cycle in autoclave. The samples were evaluated by tensile tests according to the ASTM D3039. The F584/PW laminates presented the highest values of tensile strength. However, the highest modulus results were determined for the 8HS composite laminates. The correlation of these results emphasizes the importance of the adequate combination of the polymeric matrix and the reinforcement arrangement in the structural composite manufacture. The microscopic analyses of the tested specimens show valid failure modes for composites used in aeronautical industry.

  10. Sensitivity Analysis of Linear Programming and Quadratic Programming Algorithms for Control Allocation

    Science.gov (United States)

    Frost, Susan A.; Bodson, Marc; Acosta, Diana M.

    2009-01-01

    The Next Generation (NextGen) transport aircraft configurations being investigated as part of the NASA Aeronautics Subsonic Fixed Wing Project have more control surfaces, or control effectors, than existing transport aircraft configurations. Conventional flight control is achieved through two symmetric elevators, two antisymmetric ailerons, and a rudder. The five effectors, reduced to three command variables, produce moments along the three main axes of the aircraft and enable the pilot to control the attitude and flight path of the aircraft. The NextGen aircraft will have additional redundant control effectors to control the three moments, creating a situation where the aircraft is over-actuated and where a simple relationship does not exist anymore between the required effector deflections and the desired moments. NextGen flight controllers will incorporate control allocation algorithms to determine the optimal effector commands and attain the desired moments, taking into account the effector limits. Approaches to solving the problem using linear programming and quadratic programming algorithms have been proposed and tested. It is of great interest to understand their relative advantages and disadvantages and how design parameters may affect their properties. In this paper, we investigate the sensitivity of the effector commands with respect to the desired moments and show on some examples that the solutions provided using the l2 norm of quadratic programming are less sensitive than those using the l1 norm of linear programming.

  11. Reverse Engineering Crosswind Limits - A New Flight Test Technique?

    Science.gov (United States)

    Asher, Troy A.; Willliams, Timothy L.; Strovers, Brian K.

    2013-01-01

    During modification of a Gulfstream III test bed aircraft for an experimental flap project, all roll spoiler hardware had to be removed to accommodate the test article. In addition to evaluating the effects on performance and flying qualities resulting from the modification, the test team had to determine crosswind limits for an airplane previously certified with roll spoilers. Predictions for the modified aircraft indicated the maximum amount of steady state sideslip available during the approach and landing phase would be limited by aileron authority rather than by rudder. Operating out of a location that tends to be very windy, an arbitrary and conservative wind limit would have either been overly restrictive or potentially unsafe if chosen poorly. When determining a crosswind limit, how much reserve roll authority was necessary? Would the aircraft, as configured, have suitable handling qualities for long-term use as a flying test bed? To answer these questions, the test team combined two typical flight test techniques into a new maneuver called the sideslip-to-bank maneuver, and was able to gather flying qualities data, evaluate aircraft response and measure trends for various crosswind scenarios. This paper will describe the research conducted, the maneuver, flight conditions, predictions, and results from this in-flight evaluation of crosswind capability.

  12. Active wing design with integrated flight control using piezoelectric macro fiber composites

    International Nuclear Information System (INIS)

    Piezoelectric macro fiber composites (MFCs) have been implemented as actuators into an active composite wing. The goal of the project was the design of a wing for an unmanned aerial vehicle (UAV) with a thin profile and integrated roll control with piezoelectric elements. The design and its optimization were based on a fully coupled structural fluid dynamics model that implemented constraints from available materials and manufacturing. A scaled prototype wing was manufactured. The design model was validated with static and preliminary dynamic tests of the prototype wing. The qualitative agreement between the numerical model and experiments was good. Dynamic tests were also performed on a sandwich wing of the same size with conventional aileron control for comparison. Even though the roll moment generated by the active wing was lower, it proved sufficient for the intended roll control of the UAV. The active wing with piezoelectric flight control constitutes one of the first examples where such a design has been optimized and the numerical model has been validated in experiments

  13. Aerodynamic Characteristic of the Active Compliant Trailing Edge Concept

    Science.gov (United States)

    Nie, Rui; Qiu, Jinhao; Ji, Hongli; Li, Dawei

    2016-06-01

    This paper introduces a novel Morphing Wing structure known as the Active Compliant Trailing Edge (ACTE). ACTE structures are designed using the concept of “distributed compliance” and wing skins of ACTE are fabricated from high-strength fiberglass composites laminates. Through the relative sliding between upper and lower wing skins which are connected by a linear guide pairs, the wing is able to achieve a large continuous deformation. In order to present an investigation about aerodynamics and noise characteristics of ACTE, a series of 2D airfoil analyses are established. The aerodynamic characteristics between ACTE and conventional deflection airfoil are analyzed and compared, and the impacts of different ACTE structure design parameters on aerodynamic characteristics are discussed. The airfoils mentioned above include two types (NACA0012 and NACA64A005.92). The computing results demonstrate that: compared with the conventional plane flap airfoil, the morphing wing using ACTE structures has the capability to improve aerodynamic characteristic and flow separation characteristic. In order to study the noise level of ACTE, flow field analysis using LES model is done to provide noise source data, and then the FW-H method is used to get the far field noise levels. The simulation results show that: compared with the conventional flap/aileron airfoil, the ACTE configuration is better to suppress the flow separation and lower the overall sound pressure level.

  14. Wind Tunnel Test of an RPV with Shape-Change Control Effector and Sensor Arrays

    Science.gov (United States)

    Raney, David L.; Cabell, Randolph H.; Sloan, Adam R.; Barnwell, William G.; Lion, S. Todd; Hautamaki, Bret A.

    2004-01-01

    A variety of novel control effector concepts have recently emerged that may enable new approaches to flight control. In particular, the potential exists to shift the composition of the typical aircraft control effector suite from a small number of high authority, specialized devices (rudder, aileron, elevator, flaps), toward larger numbers of smaller, less specialized, distributed device arrays. The concept envisions effector and sensor networks composed of relatively small high-bandwidth devices able to simultaneously perform a variety of control functions using feedback from disparate data sources. To investigate this concept, a remotely piloted flight vehicle has been equipped with an array of 24 trailing edge shape-change effectors and associated pressure measurements. The vehicle, called the Multifunctional Effector and Sensor Array (MESA) testbed, was recently tested in NASA Langley's 12-ft Low Speed wind tunnel to characterize its stability properties, control authorities, and distributed pressure sensitivities for use in a dynamic simulation prior to flight testing. Another objective was to implement and evaluate a scheme for actively controlling the spanwise pressure distribution using the shape-change array. This report describes the MESA testbed, design of the pressure distribution controller, and results of the wind tunnel test.

  15. UAV Flight Control System Based on an Intelligent BEL Algorithm

    Directory of Open Access Journals (Sweden)

    Huangzhong Pu

    2013-02-01

    Full Text Available A novel intelligent control strategy based on a brain emotional learning (BEL algorithm is investigated in the application of the attitude control of a small unmanned aerial vehicle (UAV in this study. The BEL model imitates the emotional learning process in the amygdala‐ orbitofrontal (A‐O system of mammalian brains. Here it is used to develop the flight control system of the UAV. The control laws of elevator, aileron and rudder manipulators adopt the forms of traditional flight control laws, and three BEL models are used in above three control loops, to on‐ line regulate the control gains of each controller. Obviously, a BEL intelligent control system is self‐learning and self‐adaptive, which is important for UAVs when flight conditions change, while traditional flight control systems remain unchanged after design. In simulation, the UAV is on a flat flight and suddenly a wind disturbs it making it depart from the equilibrium state. In order to make the UAV recover to the original equilibrium state, the BEL intelligent control system is adopted. The simulation results illustrate that the BEL‐based intelligent flight control system has characteristics of better adaptability and stronger robustness, when compared with the traditional flight control system.

  16. 可重构旋翼无人飞行器的动力学建模与分析%Dynamics Modeling and Analysis of a Reconfigurable Rotorcraft Unmanned Aerial Vehicle

    Institute of Scientific and Technical Information of China (English)

    阮晓钢; 侯旭阳; 龚道雄

    2013-01-01

    提出了一种具有可重构能力的旋翼无人飞行器(RUAV),其执行机构主要由内置在涵道中的主旋翼、环绕主旋翼的4个辅旋翼以及涵道末端的2个副翼组成,其中辅旋翼与主旋翼、副翼的部分功能重合以使系统具备重构控制能力.运用牛顿-欧拉方法建立了旋翼无人飞行器的6自由度(6DOF)动力学模型.基于此模型,首先分析了在悬停状态附近系统发生不同故障时的可控性,然后基于控制可重构度的概念分析了在发生不同程度故障时系统的容错能力,在此基础上构建了飞行器的多模型重构控制器,最后通过仿真实验分别对系统的动态响应特性和重构控制效果进行了分析.结果显示,旋翼无人飞行器具有较好的动态响应特性,且对一定范围内的故障具有较好的鲁棒性.本文提出的模型及相关分析为旋翼无人飞行器的容错设计和控制提供了一定的理论依据.%A kind of reconfigurable rotorcraft unmanned aerial vehicle (RUAV) is presented. Its actuators consist of the major rotor inside the duct, four auxiliary rotors surrounding the major rotor and two ailerons at the end of the duct, and the auxiliary rotors have some similar functions as that of the major rotor and ailerons in order to achieve reconfigurable control of the system. The Newton-Euler method is adopted to build 6-DOF (degree of freedom) dynamic model of the RUAV. Based on the model, the controllability of the system in different fault cases is analyzed near the hover state. Then, the fault-tolerance performance of the system with different fault degrees is analyzed based on the notation of control reconfigurability, and the analysis helps to build the multi-model reconfigurable controller. At last, the dynamic response characteristics and reconfigurable control performance of the system are analyzed by simulation, respectively. The result shows that the RUAV has good dynamic response characteristics and

  17. Rotary Balance Wind Tunnel Testing for the FASER Flight Research Aircraft

    Science.gov (United States)

    Denham, Casey; Owens, D. Bruce

    2016-01-01

    Flight dynamics research was conducted to collect and analyze rotary balance wind tunnel test data in order to improve the aerodynamic simulation and modeling of a low-cost small unmanned aircraft called FASER (Free-flying Aircraft for Sub-scale Experimental Research). The impetus for using FASER was to provide risk and cost reduction for flight testing of more expensive aircraft and assist in the improvement of wind tunnel and flight test techniques, and control laws. The FASER research aircraft has the benefit of allowing wind tunnel and flight tests to be conducted on the same model, improving correlation between wind tunnel, flight, and simulation data. Prior wind tunnel tests include a static force and moment test, including power effects, and a roll and yaw damping forced oscillation test. Rotary balance testing allows for the calculation of aircraft rotary derivatives and the prediction of steady-state spins. The rotary balance wind tunnel test was conducted in the NASA Langley Research Center (LaRC) 20-Foot Vertical Spin Tunnel (VST). Rotary balance testing includes runs for a set of given angular rotation rates at a range of angles of attack and sideslip angles in order to fully characterize the aircraft rotary dynamics. Tests were performed at angles of attack from 0 to 50 degrees, sideslip angles of -5 to 10 degrees, and non-dimensional spin rates from -0.5 to 0.5. The effects of pro-spin elevator and rudder deflection and pro- and anti-spin elevator, rudder, and aileron deflection were examined. The data are presented to illustrate the functional dependence of the forces and moments on angle of attack, sideslip angle, and angular rate for the rotary contributions to the forces and moments. Further investigation is necessary to fully characterize the control effectors. The data were also used with a steady state spin prediction tool that did not predict an equilibrium spin mode.

  18. Design, realization and structural testing of a compliant adaptable wing

    Science.gov (United States)

    Molinari, G.; Quack, M.; Arrieta, A. F.; Morari, M.; Ermanni, P.

    2015-10-01

    This paper presents the design, optimization, realization and testing of a novel wing morphing concept, based on distributed compliance structures, and actuated by piezoelectric elements. The adaptive wing features ribs with a selectively compliant inner structure, numerically optimized to achieve aerodynamically efficient shape changes while simultaneously withstanding aeroelastic loads. The static and dynamic aeroelastic behavior of the wing, and the effect of activating the actuators, is assessed by means of coupled 3D aerodynamic and structural simulations. To demonstrate the capabilities of the proposed morphing concept and optimization procedure, the wings of a model airplane are designed and manufactured according to the presented approach. The goal is to replace conventional ailerons, thus to achieve controllability in roll purely by morphing. The mechanical properties of the manufactured components are characterized experimentally, and used to create a refined and correlated finite element model. The overall stiffness, strength, and actuation capabilities are experimentally tested and successfully compared with the numerical prediction. To counteract the nonlinear hysteretic behavior of the piezoelectric actuators, a closed-loop controller is implemented, and its capability of accurately achieving the desired shape adaptation is evaluated experimentally. Using the correlated finite element model, the aeroelastic behavior of the manufactured wing is simulated, showing that the morphing concept can provide sufficient roll authority to allow controllability of the flight. The additional degrees of freedom offered by morphing can be also used to vary the plane lift coefficient, similarly to conventional flaps. The efficiency improvements offered by this technique are evaluated numerically, and compared to the performance of a rigid wing.

  19. Gust Alleviation of a Large Aircraft with a Passive Twist Wingtip

    Directory of Open Access Journals (Sweden)

    Shijun Guo

    2015-04-01

    Full Text Available This paper presents an investigation into the gust response and wing structure load alleviation of a 200-seater aircraft by employing a passive twist wingtip (PTWT. The research was divided into three stages. The first stage was the design and analysis of the baseline aircraft, including aerodynamic analysis, structural design using the finite element (FE method and flutter analysis to meet the design requirements. Dynamic response analysis of the aircraft to discrete (one-cosin gust was also performed in a range of gust radiances specified in the airworthiness standards. In the second stage, a PTWT of a length of 1.13 m was designed with the key parameters determined based on design constraints and, in particular, the aeroelastic stability and gust response. Subsequent gust response analysis was performed to evaluate the effectiveness of the PTWT for gust alleviation. The results show that the PTWT produced a significant reduction of gust-induced wingtip deflection by 21% and the bending moment at the wing root by 14% in the most critical flight case. In the third stage, effort was made to study the interaction and influence of the PTWT on the symmetric and unsymmetrical manoeuvring of the aircraft when ailerons were in operation. The results show the that PTWT influence with a reduction of the aircraft normal velocity and heave motion by 1.7% and 3%, respectively, is negligible. However, the PTWT influence on the aircraft roll moment with a 20.5% reduction is significant. A locking system is therefore required in such a manoeuvring condition. The investigation has shown that the PTWT is an effective means for gust alleviation and, therefore, has potential for large aircraft application.

  20. Performance, Stability, and Control Investigation at Mach Numbers from 0.60 to 1.05 of a Model of the "Swallow" with Outer Wing Panels Swept 75 degree with and without Power Simulations

    Science.gov (United States)

    Schmeer, James W.; Cassetti, Marlowe D.

    1960-01-01

    An investigation of the performance, stability, and control characteristics of a variable-sweep arrow-wing model with the outer wing panels swept 75 deg. has been conducted in the Langley 16-foot transonic tunnel. Four outboard engines located above and below the wing provided propulsive thrust, and, by deflecting in the pitch direction and rotating in the lateral plane, also produced control forces. The engine nacelles incorporated swept lateral and vertical fins for aerodynamic stability and control. Jet-off data were obtained with flow-through nacelles, simulating inlet flow; jet thrust and hot-jet interference effects were obtained with faired-nose nacelles housing hydrogen peroxide gas generators. Six-component force and moment data were obtained at Mach numbers from 0.60 to 1.05 through a range of angles of attack and angles of side-slip. Control characteristics were obtained by deflecting the nacelle-fin combinations as elevators, rudders, and ailerons at several fixed angles for each control. The results indicate that the basic wing-body configuration becomes neutrally stable or unstable at a lift coefficient of 0.15; addition of nacelles with fins delayed instability to a lift coefficient of 0.30. Addition of nacelles to the wing-body configuration increased minimum drag from 0.0058 to 0.0100 at a Mach number of 0.60 and from 0.0080 to 0.0190 at a Mach number of 1.05 with corresponding reductions in maximum lift-drag ratio of 12 percent and 33 percent, respectively. The nacelle-fin combinations were ineffective as longitudinal controls but were adequate as directional and lateral controls. The model with nacelles and fins was directionally and laterally stable; the stability generally increased with increasing lift. Jet interference effects on stability and control characteristics were small but the adverse effects on drag were greater than would be expected for isolated nacelles.

  1. Neural Flight Control System

    Science.gov (United States)

    Gundy-Burlet, Karen

    2003-01-01

    The Neural Flight Control System (NFCS) was developed to address the need for control systems that can be produced and tested at lower cost, easily adapted to prototype vehicles and for flight systems that can accommodate damaged control surfaces or changes to aircraft stability and control characteristics resulting from failures or accidents. NFCS utilizes on a neural network-based flight control algorithm which automatically compensates for a broad spectrum of unanticipated damage or failures of an aircraft in flight. Pilot stick and rudder pedal inputs are fed into a reference model which produces pitch, roll and yaw rate commands. The reference model frequencies and gains can be set to provide handling quality characteristics suitable for the aircraft of interest. The rate commands are used in conjunction with estimates of the aircraft s stability and control (S&C) derivatives by a simplified Dynamic Inverse controller to produce virtual elevator, aileron and rudder commands. These virtual surface deflection commands are optimally distributed across the aircraft s available control surfaces using linear programming theory. Sensor data is compared with the reference model rate commands to produce an error signal. A Proportional/Integral (PI) error controller "winds up" on the error signal and adds an augmented command to the reference model output with the effect of zeroing the error signal. In order to provide more consistent handling qualities for the pilot, neural networks learn the behavior of the error controller and add in the augmented command before the integrator winds up. In the case of damage sufficient to affect the handling qualities of the aircraft, an Adaptive Critic is utilized to reduce the reference model frequencies and gains to stay within a flyable envelope of the aircraft.

  2. Differential Canard deflection for generation of yawing moment on the X-31 with and without the vertical tail. M.S. Thesis - George Washington Univ.

    Science.gov (United States)

    Whiting, Matthew Robert

    1996-01-01

    The feasibility of augmenting the available yaw control power on the X-31 through differential deflection of the canard surfaces was studied as well as the possibility of using differential canard control to stabilize the X-31 with its vertical tail removed. Wind-tunnel tests and the results of departure criteria and linear analysis showed the destabilizing effect of the reduction of the vertical tail on the X-31. Wind-tunnel testing also showed that differential canard deflection was capable of generating yawing moments of roughly the same magnitude as the thrust vectoring vanes currently in place on the X-31 in the post-stall regime. Analysis showed that the X-31 has sufficient aileron roll control power that with the addition of differential canard as a yaw controller, the wind-axis roll accelerations will remain limited by yaw control authority. It was demonstrated, however, that pitch authority may actually limit the maximum roll rate which can be sustained. A drop model flight test demonstrated that coordinated, wind axis rolls could be performed with roll rates as high as 50 deg/sec (full scale equivalent) at 50 deg angle of attack. Another drop model test was conducted to assess the effect of vertical tail reduction, and an analysis of using differential canard deflection to stabilize the tailless X-31 was performed. The results of six-degree-of-freedom, non-linear simulation tests were correlated with the drop model flights. Simulation studies then showed that the tailless X-31 could be controlled at angles of attack at or above 20 deg using differential canard as the only yaw controller.

  3. Using Remotely Piloted Aircraft System to Study the Evolution of the Boundary Layer Related to Fog Events

    Science.gov (United States)

    Roberts, G. C.; Cayez, G.; Ronflé-Nadaud, C.; Albrand, M.; Dralet, J. P.; Momboisse, G.; Nicoll, K.; Seity, Y.; Bronz, M.; Hattenberger, G.; Gorraz, M.; Bustico, A.

    2014-12-01

    Over the past decade, the scientific community has embraced the use of RPAS (remotely piloted aircraft system) as a tool to improve observations of the Earth's surface and atmospheric phenomena. The use of small RPAS (Remotely Piloted Aircraft System) in atmospheric research has increased because of their relative low-cost, compact size and ease of operation. Small RPAS are especially adapted for observing the atmospheric boundary layer processes at high vertical and temporal resolution. To this end, CNRM, ENAC, and ENM have developed the VOLTIGE (Vecteurs d'Observation de La Troposphere pour l'Investigation et la Gestion de l'Environnement) program to study the life cycle of fog with multiple, small RPAS. The instrumented RPAS flights have successfully observed the evolution of the boundary layer and dissipation of fog events. In addition, vertical profiles from the RPAS have been compared with Météo France forecast models, and the results suggest that forecast models may be improved using high resolution and frequent in-situ measurements. Within the VOLTIGE project, a flying-wing RPAS with four control surfaces was developed to separate elevator and aileron controls in order to reduce the pitch angle envelope and improve turbulence and albedo measurements. The result leads to a small RPAS with the capability of flying up to two hours with 150 grams of payload, while keeping the hand-launch capability as a constraint for regular atmospheric research missions. High frequency data logging has been integrated into the main autopilot in order to synchronize navigation and payload measurements, as well as allowing an efficient sensor-based navigation. The VOLTIGE program also encourages direct participation of students on the advancement of novel observing systems for atmospheric sciences, and provides a step towards deploying small RPAS in an operational network. VOLTIGE is funded by the Agence Nationale de Recherche (ANR-Blanc 2012) and supported by Aerospace

  4. Aircraft Abnormal Conditions Detection, Identification, and Evaluation Using Innate and Adaptive Immune Systems Interaction

    Science.gov (United States)

    Al Azzawi, Dia

    simulator. The abnormal conditions considered in this work include locked actuators (stabilator, aileron, rudder, and throttle), structural damage of the wing, horizontal tail, and vertical tail, malfunctioning sensors, and reduced engine effectiveness. The results of applying the proposed approach to this wide range of abnormal conditions show its high capability in detecting the abnormal conditions with zero false alarms and very high detection rates, correctly identifying the failed subsystem and evaluating the type and severity of the failure. The results also reveal that the post-failure flight envelope can be reasonably predicted within this framework.

  5. Optimal Aircraft Control Upset Recovery With and Without Component Failures

    Science.gov (United States)

    Sparks, Dean W.; Moerder, Daniel D.

    2002-01-01

    This paper treats the problem of recovering sustainable nondescending (safe) flight in a transport aircraft after one or more of its control effectors fail. Such recovery can be a challenging goal for many transport aircraft currently in the operational fleet for two reasons. First, they have very little redundancy in their means of generating control forces and moments. These aircraft have, as primary control surfaces, a single rudder and pairwise elevators and aileron/spoiler units that provide yaw, pitch, and roll moments with sufficient bandwidth to be used in stabilizing and maneuvering the airframe. Beyond this, throttling the engines can provide additional moments, but on a much slower time scale. Other aerodynamic surfaces, such as leading and trailing edge flaps, are only intended to be placed in a position and left, and are, hence, very slow-moving. Because of this, loss of a primary control surface strongly degrades the controllability of the vehicle, particularly when the failed effector becomes stuck in a non-neutral position where it exerts a disturbance moment that must be countered by the remaining operating effectors. The second challenge in recovering safe flight is that these vehicles are not agile, nor can they tolerate large accelerations. This is of special importance when, at the outset of the recovery maneuver, the aircraft is flying toward the ground, as is frequently the case when there are major control hardware failures. Recovery of safe flight is examined in this paper in the context of trajectory optimization. For a particular transport aircraft, and a failure scenario inspired by an historical air disaster, recovery scenarios are calculated with and without control surface failures, to bring the aircraft to safe flight from the adverse flight condition that it had assumed, apparently as a result of contact with a vortex from a larger aircraft's wake. An effort has been made to represent relevant airframe dynamics, acceleration limits

  6. Analysis of Aircraft Control Performance using a Fuzzy Rule Base Representation of the Cooper-Harper Aircraft Handling Quality Rating

    Science.gov (United States)

    Tseng, Chris; Gupta, Pramod; Schumann, Johann

    2006-01-01

    control; the tracking error is a good measurement for performance needed in the rating scheme. Finally, the change of the control amount or the output of a confidence tool, which has been developed by the authors, can be used as an indication of pilot compensation. We use a number of known aircraft flight scenarios with known pilot ratings to calibrate our fuzzy membership functions. These include normal flight conditions and situations in which partial or complete failure of tail, aileron, engine, or throttle occurs.

  7. Flight testing of a remotely piloted vehicle for aircraft parameter estimation purposes

    Science.gov (United States)

    Seanor, Brad A.

    2002-01-01

    The contribution of this research effort was to show that a reliable RPV could be built, tested, and successfully used for flight testing and parameter estimation purposes, in an academic setting. This was a fundamental step towards the creation of an automated Unmanned Aerial Vehicle (UAV). This research project was divided into four phases. Phase one involved the construction, development, and initial flight of a Remotely Piloted Vehicle (RPV), the West Virginia University (WVU) Boeing 777 (B777) aircraft. This phase included the creation of an onboard instrumentation system to provide aircraft flight data. The objective of the second phase was to estimate the longitudinal and lateral-directional stability and control derivatives from actual flight data for the B777 model. This involved performing and recording flight test maneuvers used for analysis of the longitudinal and lateral-directional estimates. Flight maneuvers included control surface doublets produced by the elevator, aileron, and rudder controls. A parameter estimation program known as pEst, developed at NASA Dryden Flight Research Center (DFRC), was used to compute the off-line estimates of parameters from collected flight data. This estimation software uses the Maximum Likelihood (ML) method with a Newton-Raphson (NR) minimization algorithm. The mathematical model used a traditional static and dynamic derivative buildup. Phase three focused on comparing a linear model obtained from the phase two ML estimates, with linear models obtained from a (i) Batch Least Squares Technique (BLS) and (ii) a technique from the Matlab system identification toolbox. Historically, aircraft parameter estimation has been performed off-line using recorded flight data from specifically designed maneuvers. In recent years, several on-line parameter identification techniques have been evaluated for real-time on-line applications. Along this research line, a novel contribution of this work was to compare the off

  8. Contribution of stable isotopes and age dating tools to the understanding of pesticide transfer into surface and ground-waters in Martinique (French West Indies)

    Science.gov (United States)

    Gourcy, Laurence; Arnaud, Luc; Baran, Nicole; Petelet-Giraud, Emmanuelle

    2013-04-01

    In Martinique, chlordecone, a synthetic chlorinated organic compound has mainly been used as an insecticide for banana farming up to 1993. The intrinsic characteristic of this contaminant makes it still quite abundant in soil, surface and groundwater. Since 2004 and the implementation of the Water Framework Directive the concentration of chlordecone in groundwater has been monitored regularly (two to four times / year) at different points of the island by the ODE (Office de l'Eau). Previous study (Gourcy et al. 2009, Arnaud et al. 2012) showed that variations of pesticides concentrations in groundwater are temporally strong and not always easy to correlate to climate, geological or hydrogeological context. The objective of the present study was to explore new investigation ways to identify, in a specific site and for high sampling frequency possible pathways of chlordecone into surface and ground-waters. A major sampling campaign was carried out in December 2011 including 12 surface and groundwater points located in Chalvet and Chez Lélène wells watersheds. Besides, monthly or weekly samples were taken at these two groundwater monitoring wells and the Falaise river up to August 2012. Major dissolved ions, δ18O, δ2H, chlordecone concentrations were determined for all samples. CFC-11, CFC-12, CFC-113 and SF6 analyses were performed for groundwater for apparent age estimation. Punctual or cumulative rainfalls were sampled at Chalvet (30 m NGM) and Aileron (800 m NGM) for stable isotopes determination. The isotope data are indicating a deuterium excess higher for surface water, groundwater and rainfall collected at high altitude vs. samples corresponding to lowest altitudes. This data can therefore be used to estimate the average altitude of recharge area of groundwater. This altitude of recharge, between 30 and 350m corresponds to the altitude of banana growing ; it is therefore in accordance with the presence of chlordecone in soils. This information is also

  9. The influence of time dependent flight and maneuver velocities and elastic or viscoelastic flexibilities on aerodynamic and stability derivatives

    Energy Technology Data Exchange (ETDEWEB)

    Cochrane, Alexander P. [Aerospace Engineering Department, University of Glasgow, University Avenue, Glasgow, Lanarkshire (United Kingdom); Merrett, Craig G. [Mechanical and Aerospace Engineering Department, Carleton Univ., 1125 Col. By Dr., Ottawa, ON (Canada); Hilton, Harry H. [Aerospace Engineering Department in the College of Engineering and Private Sector Program Division at the National Center for Supercomputing Applications, University of Illinois at Urbana-Champaign, 104 South Wright Street, Urbana, IL 61801 (United States)

    2014-12-10

    which control reversal takes place (V{sub REV}{sup E}). Since elastic formulations constitute viscoelastic initial conditions, viscoelastic reversal may occur at speeds V{sub REV<}{sup ≧}V{sub REV}{sup E}, but furthermore does so in time at 0 < t{sub REV} ≤ ∞. This paper reports on analytical analyses and simulations of the effects of flexibility and time dependent material properties (viscoelasticity) on aerodynamic derivatives and on lateral, longitudinal, directional and spin stability derivatives. Cases of both constant and variable flight and maneuver velocities are considered. Analytical results for maneuvers involving constant and time dependent rolling velocities are analyzed, discussed and evaluated. The relationships between rolling velocity p and aileron angular displacement β as well as control effectiveness are analyzed and discussed in detail for elastic and viscoelastic wings. Such analyses establish the roll effectiveness derivatives (∂[p(t)])/(V{sub ∞}∂β(t)) . Similar studies involving other stability and aerodynamic derivatives are also undertaken. The influence of the twin effects of viscoelastic and elastic materials and of variable flight, rolling, pitching and yawing velocities on longitudinal, lateral and directional are also investigated. Variable flight velocities, encountered during maneuvers, render the usually linear problem at constant velocities into a nonlinear one.

  10. 小型无人倾转旋翼机气动与操纵特性试验研究%Testing study on aerodynamics and control characteristics of a small unmanned tilt rotor

    Institute of Scientific and Technical Information of China (English)

    郭剑东; 宋彦国

    2015-01-01

    It is very difficult to determine the aerodynamics and control characteristics theo-retically for tiltrotor aircraft because of multi-flight modes,complexity of aerodynamic interac-tions,and redundancy of control surfaces.Especially for the tilting flight mode,the layout of the aircraft is transformed between the helicopter mode and the fixed-wing airplane mode with the na-celle driven rotor system tilting.In order to investigate the aerodynamics and control characteris-tics,the full-span and full-envelop flight modes of a small unmanned tilt rotor are tested in wind-tunnel prior to flight.The un-powered test is mainly determining the flight characteristics with different attack angles,nacelle angles and forward speeds.The powered test is focused on the aerodynamic interactions among rotor,wing and flaperon wing,with and without wings,as well as the efficacy manipulation of the collective aileron and elevator.According to the experimental data,the full-envelop flight control characteristics for the tiltrotor is deduced,improves aircraft designing,and provides a priori knowledge for successful flight tests.%由于倾转旋翼机飞行模式多,各部件气动干扰复杂且操纵面冗余,特别是倾转过渡模式,短舱带动旋翼系统倾转,结构布局发生改变,从理论上确定气动与操纵特性难度大。为了研究倾转旋翼机的气动与操纵特性,对某小型无人倾转旋翼机展开全尺寸、全模式吹风试验,其中不带动力试验主要用于研究倾转旋翼机在不同迎角、短舱倾角、前飞速度等飞行状态下的气动特性;带动力试验主要用于研究倾转旋翼机不同飞行模式带机翼与不带机翼时,旋翼/机翼/襟副翼相互干扰作用,以及总距、副翼、升降舵的操纵功效。根据试验数据推导出小型无人倾转旋翼机全包线飞行的操纵特性方法,对进一步完善倾转旋翼机设计以及试飞试验的成功提供了参考。

  11. Experiment on static aeroelastic in high speed wind tunnel%静气动弹性模型高速风洞试验研究

    Institute of Scientific and Technical Information of China (English)

    杨贤文; 余立; 吕彬彬; 郭洪涛; 杨振华; 寇西平

    2015-01-01

    对采用复合材料玻璃纤维、碳纤维加工的静气动弹性模型进行了高速风洞试验研究,测试了模型的柔度矩阵、气动力、表面压力、弯/扭应变信号及弯/扭变形,为静气动弹性模型刚度试验、弯/扭应变信号测量、模型变形视频测量(VMD)及风洞总压控制等静气动弹性风洞试验能力的提高积累了经验,为飞行器静气动弹性研究提供了良好的试验平台。研究表明:静气动弹性模型较刚性模型升力线斜率及襟副翼效率下降、气动焦点前移;静气动弹性模型与刚性模型表面压力差异明显;在小迎角范围内,静气动弹性机翼模型弯/扭应变信号随迎角增加基本呈线性变化;在正迎角时,大展弦比后掠机翼静气动弹性模型的剖面扭转变形使有效迎角减小,剖面越靠近翼尖弯/扭变形越大。%Experimental investigation on static aeroelastic was carried out in high speed wind tunnel using glass fiber models and carbon fiber models respectively.Flexibility matrix,aerody-namic force,pressure,bending/torsion strain signals and bending/torsional deformations of models are measured.Experience is accumulated for developing high speed wind tunnel static aeroelastic experimental techniques such as stiffness test,bending/torsion strain signals gauging, videogrammetric model deformation and wind tunnel total pressure control.The investigation shows:comparing with rigid model,lift curve slope and aileron/flap efficiency of static aeroelas-tic model decrease,aerodynamic center of static aerodynamic model shifts forward,pressure dis-tributions of static aeroelastic model change significantly;static aeroelastic wing model bending/torsion strain signals,which can be used to analyze aerodynamic loads,are nearly linear variation with attack angle at small attack angles,for lift coefficient and pitching moment coefficient of static aeroelastic wing model change linearly with attack

  12. New mechanism for upset of electronics.

    Energy Technology Data Exchange (ETDEWEB)

    Loubriel, Guillermo Manuel; Molina, Luis Leroy; Salazar, Robert Austin; Patterson, Paull Edward; Bacon, Larry Donald

    2004-03-01

    measured results. This study shows that models based on SPICE, although they exhibit chaotic behavior, do not properly reproduce circuit behavior without modifying diode parameters. This report describes the models and considerations used to model circuit behavior in the nonlinear range of operation. Further, it describes how a modified SPICE diode model improves the simulation results. We also studied the nonlinear behavior of a phased-locked-loop. Phased-locked loops are fundamental building block to many major systems (aileron, seeker heads, etc). We showed that an injected RF signal could drive the phased-locked-loop into chaos. During these chaotic episodes, the frequency of the phased-locked-loop takes excursion outside its normal range of operation. In addition to these excursions, the phased-locked-loop and the system it is controlling requires some time to get back into normal operation. The phased-locked-loop only needs to be upset enough long enough to keep it off balance.

  13. Energy conserving numerical methods for the computation of complex vortical flows

    Science.gov (United States)

    Allaneau, Yves

    One of the original goals of this thesis was to develop numerical tools to help with the design of micro air vehicles. Micro Air Vehicles (MAVs) are small flying devices of only a few inches in wing span. Some people consider that as their size becomes smaller and smaller, it would be increasingly more difficult to keep all the classical control surfaces such as the rudders, the ailerons and the usual propellers. Over the years, scientists took inspiration from nature. Birds, by flapping and deforming their wings, are capable of accurate attitude control and are able to generate propulsion. However, the biomimicry design has its own limitations and it is difficult to place a hummingbird in a wind tunnel to study precisely the motion of its wings. Our approach was to use numerical methods to tackle this challenging problem. In order to precisely evaluate the lift and drag generated by the wings, one needs to be able to capture with high fidelity the extremely complex vortical flow produced in the wake. This requires a numerical method that is stable yet not too dissipative, so that the vortices do not get diffused in an unphysical way. We solved this problem by developing a new Discontinuous Galerkin scheme that, in addition to conserving mass, momentum and total energy locally, also preserves kinetic energy globally. This property greatly improves the stability of the simulations, especially in the special case p=0 when the approximation polynomials are taken to be piecewise constant (we recover a finite volume scheme). In addition to needing an adequate numerical scheme, a high fidelity solution requires many degrees of freedom in the computations to represent the flow field. The size of the smallest eddies in the flow is given by the Kolmogoroff scale. Capturing these eddies requires a mesh counting in the order of Re³ cells, where Re is the Reynolds number of the flow. We show that under-resolving the system, to a certain extent, is acceptable. However our

  14. Adaptive Augmenting Control Flight Characterization Experiment on an F/A-18

    Science.gov (United States)

    VanZwieten, Tannen S.; Orr, Jeb S.; Wall, John H.; Gilligan, Eric T.

    2014-01-01

    dynamics such as slosh and bending modes, as well as atmospheric disturbances, are being produced by the airframe via modification of bending filters and the use of secondary control surfaces, including leading and trailing edge flaps, symmetric ailerons, and symmetric rudders. The platform also has the ability to inject signals in flight to simulate structural mode resonances or other challenging dynamics. This platform also offers more test maneuvers and longer maneuver times than a single rocket or missile test, which provides ample opportunity to fully and repeatedly exercise all aspects of the algorithm. Prior to testing on an F/A-18, AAC was the only component of the SLS autopilot design that had not been flight tested. The testing described in this paper raises the Technology Readiness Level (TRL) early in the SLS Program and is able to demonstrate its capabilities and robustness in a flight environment.

  15. AD-1 with research pilot Richard E. Gray

    Science.gov (United States)

    1982-01-01

    Standing in front of the AD-1 Oblique Wing research aircraft is research pilot Richard E. Gray. Richard E. Gray joined National Aeronautics and Space Administration's Johnson Space Center, Houston, Texas, in November 1978, as an aerospace research pilot. In November 1981, Dick joined the NASA's Ames-Dryden Flight Research Facility, Edwards, California, as a research pilot. Dick was a former Co-op at the NASA Flight Research Center (a previous name of the Ames-Dryden Flight Research Facility), serving as an Operations Engineer. At Ames-Dryden, Dick was a pilot for the F-14 Aileron Rudder Interconnect Program, AD-1 Oblique Wing Research Aircraft, F-8 Digital Fly-By-Wire and Pilot Induced Oscillations investigations. He also flew the F-104, T-37, and the F-15. On November 8, 1982, Gray was fatally injured in a T-37 jet aircraft while making a pilot proficiency flight. Dick graduated with a Bachelors degree in Aeronautical Engineering from San Jose State University in 1969. He joined the U.S. Navy in July 1969, becoming a Naval Aviator in January 1971, when he was assigned to F-4 Phantoms at Naval Air Station (NAS) Miramar, California. In 1972, he flew 48 combat missions in Vietnam in F-4s with VF-111 aboard the USS Coral Sea. After making a second cruise in 1973, Dick was assigned to Air Test and Evaluation Squadron Four (VX-4) at NAS Point Mugu, California, as a project pilot on various operational test and evaluation programs. In November 1978, Dick retired from the Navy and joined NASA's Johnson Space Center. At JSC Gray served as chief project pilot on the WB-57F high-altitude research projects and as the prime television chase pilot in a T-38 for the landing portion of the Space Shuttle orbital flight tests. Dick had over 3,000 hours in more than 30 types of aircraft, an airline transport rating, and 252 carrier arrested landings. He was a member of the Society of Experimental Test Pilots serving on the Board of Directors as Southwest Section Technical Adviser in

  16. AD-1 with research pilot Richard E. Gray

    Science.gov (United States)

    1982-01-01

    Standing in front of the AD-1 Oblique Wing research aircraft is research pilot Richard E. Gray. Richard E. Gray joined National Aeronautics and Space Administration's Johnson Space Center, Houston, Texas, in November 1978, as an aerospace research pilot. In November 1981, Dick joined the NASA's Ames-Dryden Flight Research Facility, Edwards, California, as a research pilot. Dick was a former Co-op at the NASA Flight Research Center (a previous name of the Ames-Dryden Flight Research Facility), serving as an Operations Engineer. At Ames-Dryden, Dick was a pilot for the F-14 Aileron Rudder Interconnect Program, AD-1 Oblique Wing Research Aircraft, F-8 Digital Fly-By-Wire and Pilot Induced Oscillations investigations. He also flew the F-104, T-37, and the F-15. On November 8, 1982, Gray was fatally injured in a T-37 jet aircraft while making a pilot proficiency flight. Dick graduated with a Bachelors degree in Aeronautical Engineering from San Jose State University in 1969. He joined the U.S. Navy in July 1969, becoming a Naval Aviator in January 1971, when he was assigned to F-4 Phantoms at Naval Air Station (NAS) Miramar, California. In 1972, he flew 48 combat missions in Vietnam in F-4s with VF-111 aboard the USS Coral Sea. After making a second cruise in 1973, Dick was assigned to Air Test and Evaluation Squadron Four (VX-4) at NAS Point Mugu, California, as a project pilot on various operational test and evaluation programs. In November 1978, Dick retired from the Navy and joined NASA's Johnson Space Center. At JSC Gray served as chief project pilot on the WB-57F high-altitude research projects and as the prime television chase pilot in a T-38 for the landing portion of the Space Shuttle orbital flight tests. Dick had over 3,000 hours in more than 30 types of aircraft, an airline transport rating, and 252 carrier arrested landings. He was a member of the Society of Experimental Test Pilots serving on the Board of Directors as Southwest Section Technical Adviser in

  17. SR-71 Research Engineer Marta Bohn-Meyer

    Science.gov (United States)

    1992-01-01

    extending into the airstream. One of Dryden's SR-71s was used for the Linear Aerospike Rocket Engine, or LASRE Experiment. Another earlier project consisted of a series of flights using the SR-71 as a science camera platform for NASA's Jet Propulsion Laboratory in Pasadena, California. An upward-looking ultraviolet video camera placed in the SR-71's nosebay studied a variety of celestial objects in wavelengths that are blocked to ground-based astronomers. Earlier in its history, Dryden had a decade of past experience at sustained speeds above Mach 3. Two YF-12A aircraft and an SR-71 designated as a YF-12C were flown at the center between December 1969 and November 1979 in a joint NASA/USAF program to learn more about the capabilities and limitations of high-speed, high-altitude flight. The YF-12As were prototypes of a planned interceptor aircraft based on a design that later evolved into the SR-71 reconnaissance aircraft. Dave Lux was the NASA SR-71 project manger for much of the decade of the 1990s, followed by Steve Schmidt. Developed for the USAF as reconnaissance aircraft more than 30 years ago, SR-71s are still the world's fastest and highest-flying production aircraft. The aircraft can fly at speeds of more than 2,200 miles per hour (Mach 3+, or more than three times the speed of sound) and at altitudes of over 85,000 feet. The Lockheed Skunk Works (now Lockheed Martin) built the original SR-71 aircraft. Each aircraft is 107.4 feet long, has a wingspan of 55.6 feet, and is 18.5 feet high (from the ground to the top of the rudders, when parked). Gross takeoff weight is about 140,000 pounds, including a possible fuel weight of 80,280 pounds. The airframes are built almost entirely of titanium and titanium alloys to withstand heat generated by sustained Mach 3 flight. Aerodynamic control surfaces consist of all-moving vertical tail surfaces, ailerons on the outer wings, and elevators on the trailing edges between the engine exhaust nozzles. The two SR-71s at Dryden have

  18. SR-71 Pilots and Crew (Smith, Meyer, Bohn-Meyer, Ishmael)

    Science.gov (United States)

    1991-01-01

    LASRE Experiment. Another earlier project consisted of a series of flights using the SR-71 as a science camera platform for NASA's Jet Propulsion Laboratory in Pasadena, California. An upward-looking ultraviolet video camera placed in the SR-71's nosebay studied a variety of celestial objects in wavelengths that are blocked to ground-based astronomers. Earlier in its history, Dryden had a decade of past experience at sustained speeds above Mach 3. Two YF-12A aircraft and an SR-71 designated as a YF-12C were flown at the center between December 1969 and November 1979 in a joint NASA/USAF program to learn more about the capabilities and limitations of high-speed, high-altitude flight. The YF-12As were prototypes of a planned interceptor aircraft based on a design that later evolved into the SR-71 reconnaissance aircraft. Dave Lux was the NASA SR-71 project manger for much of the decade of the 1990s, followed by Steve Schmidt. Developed for the USAF as reconnaissance aircraft more than 30 years ago, SR-71s are still the world's fastest and highest-flying production aircraft. The aircraft can fly at speeds of more than 2,200 miles per hour (Mach 3+, or more than three times the speed of sound) and at altitudes of over 85,000 feet. The Lockheed Skunk Works (now Lockheed Martin) built the original SR-71 aircraft. Each aircraft is 107.4 feet long, has a wingspan of 55.6 feet, and is 18.5 feet high (from the ground to the top of the rudders, when parked). Gross takeoff weight is about 140,000 pounds, including a possible fuel weight of 80,280 pounds. The airframes are built almost entirely of titanium and titanium alloys to withstand heat generated by sustained Mach 3 flight. Aerodynamic control surfaces consist of all-moving vertical tail surfaces, ailerons on the outer wings, and elevators on the trailing edges between the engine exhaust nozzles. The two SR-71s at Dryden have been assigned the following NASA tail numbers: NASA 844 (A model), military serial 61-7980 and NASA

  19. Dale Reed with model in front of M2-F1

    Science.gov (United States)

    1967-01-01

    Briegleb Glider Company. The budget was $30,000. NASA craftsmen and engineers built the tubular steel interior frame. Its mahogany plywood shell was hand-made by Gus Briegleb and company. Ernie Lowder, a NASA craftsman who had worked on the Howard Hughes 'Spruce Goose,' was assigned to help Briegleb. The prototype of a 21st Century spacecraft required the fabrication of hundreds of small wooden parts meticulously nailed and glued together. It was a product of craftsmanship that was nearly obsolete in the 1940s. Final assembly of the remaining components (including aluminum tail surfaces, push rod controls, and landing gear from a Cessna 150) was done back at the NASA facility. In the meantime, other NASA engineers devised a special M2-F1 flight simulator, and a hot rod shop near Long Beach souped-up a Pontiac convertible to be used as the lifting body ground-tow vehicle. The M2-F1 did not have ailerons. Instead, it had elevons which were attached to each of the two rudders. A large flap on the trailing edge of the body acted as an elevator. This unconventional arrangement prompted the engineers to rethink the flight control system as well. They eventually devised two schemes. One system was fairly traditional. It used rudder pedal inputs to move the rudders for yaw control, and stick inputs to provide differential deflections of the elevons for roll. The other system used stick inputs to control the rudders for yaw, while rudder pedal deflections moved the elevons for roll. Milt Thompson tried both systems in the simulator and surprised the design team when he said he preferred system number two. He reasoned that although sideslip delayed roll (which was a result of dihedral effect), the roll rate was twice as high using the rudders instead of the elevons. He said he would rather have the higher roll rates available to him if needed, while the slip could be overcome with proper piloting technique. This was the system that Thompson practiced on the simulator, and he used it

  20. M2-F1 on lakebed with pilots Milt Thompson, Chuck Yeager, Don Mallick, and Bruce Peterson

    Science.gov (United States)

    1963-01-01

    -made by Gus Briegleb and company. Ernie Lowder, a NASA craftsman who had worked on the Howard Hughes 'Spruce Goose,' was assigned to help Briegleb. The prototype of a 21st Century spacecraft required the fabrication of hundreds of small wooden parts meticulously nailed and glued together. It was a product of craftsmanship that was nearly obsolete in the 1940s. Final assembly of the remaining components (including aluminum tail surfaces, push rod controls, and landing gear from a Cessna 150) was done back at the NASA facility. In the meantime, other NASA engineers devised a special M2-F1 flight simulator, and a hot rod shop near Long Beach souped-up a Pontiac convertible to be used as the lifting body ground-tow vehicle. The M2-F1 did not have ailerons. Instead, it had elevons which were attached to each of the two rudders. A large flap on the trailing edge of the body acted as an elevator. This unconventional arrangement prompted the engineers to rethink the flight control system as well. They eventually devised two schemes. One system was fairly traditional. It used rudder pedal inputs to move the rudders for yaw control, and stick inputs to provide differential deflections of the elevons for roll. The other system used stick inputs to control the rudders for yaw, while rudder pedal deflections moved the elevons for roll. Milt Thompson tried both systems in the simulator and surprised the design team when he said he preferred system number two. He reasoned that although sideslip delayed roll (which was a result of dihedral effect), the roll rate was twice as high using the rudders instead of the elevons. He said he would rather have the higher roll rates available to him if needed, while the slip could be overcome with proper piloting technique. This was the system that Thompson practiced on the simulator, and he used it during the initial auto tows. Auto tows were done using a 1000 foot rope fastened to the NASA Pontiac. Rogers Dry Lake provided miles of unobstructed