Sample records for ailerons

  1. Aileron roll hysteresis effects on entry of space shuttle orbiter (United States)

    Powell, R. W.


    Six-degree-of-freedom simulations of the space shuttle orbiter entry with control hysteresis were conducted on the NASA Langley Research Center interactive simulator known as the automatic reentry flight dynamics simulator. These simulations revealed that the vehicle can tolerate control hysteresis producing a + or - 50 percent change in the nominal aileron roll characteristics and an offset in the nominal characteristics equivalent to a + or - 5 deg aileron deflection with little increase in the reaction control system's fuel consumption.

  2. Evaluation of aileron actuator reliability with censored data

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    Li Huaiyuan


    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.

  3. Mathematical modeling of a V-stack piezoelectric aileron actuation

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    Ioan URSU


    Full Text Available The article presents a mathematical modeling of aileron actuation that uses piezo V-shaped stacks. The aim of the actuation is the increasing of flutter speed in the context of a control law, in order to widen the flight envelope. In this way the main advantage of such a piezo actuator, the bandwidth is exploited. The mathematical model is obtained based on free body diagrams, and the numerical simulations allow a preliminary sizing of the actuator.

  4. Aircraft wing structural detail design (wing, aileron, flaps, and subsystems) (United States)

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


    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.

  5. Reflection plane tests of a wind turbine blade tip section with ailerons (United States)

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


    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.

  6. Flight Investigation of the Effectiveness of an Automatic Aileron Trim Control Device for Personal Airplanes (United States)

    Phillips, William H; Kuehnel, Helmut A; Whitten, James B


    A flight investigation to determine the effectiveness of an automatic aileron trim control device installed in a personal airplane to augment the apparent spiral stability has been conducted. The device utilizes a rate-gyro sensing element in order to switch an on-off type of control that operates the ailerons at a fixed rate through control centering springs. An analytical study using phase-plane and analog-computer methods has been carried out to determine a desirable method of operation for the automatic trim control.

  7. Experimental characterization of an adaptive aileron: lab tests and FE correlation (United States)

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


    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.

  8. A Method for Aileron Actuator Fault Diagnosis Based on PCA and PGC-SVM

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    Wei-Li Qin


    Full Text Available Aileron actuators are pivotal components for aircraft flight control system. Thus, the fault diagnosis of aileron actuators is vital in the enhancement of the reliability and fault tolerant capability. This paper presents an aileron actuator fault diagnosis approach combining principal component analysis (PCA, grid search (GS, 10-fold cross validation (CV, and one-versus-one support vector machine (SVM. This method is referred to as PGC-SVM and utilizes the direct drive valve input, force motor current, and displacement feedback signal to realize fault detection and location. First, several common faults of aileron actuators, which include force motor coil break, sensor coil break, cylinder leakage, and amplifier gain reduction, are extracted from the fault quadrantal diagram; the corresponding fault mechanisms are analyzed. Second, the data feature extraction is performed with dimension reduction using PCA. Finally, the GS and CV algorithms are employed to train a one-versus-one SVM for fault classification, thus obtaining the optimal model parameters and assuring the generalization of the trained SVM, respectively. To verify the effectiveness of the proposed approach, four types of faults are introduced into the simulation model established by AMESim and Simulink. The results demonstrate its desirable diagnostic performance which outperforms that of the traditional SVM by comparison.

  9. An Aileron Flutter Experiment and Analysis Using Semi-Span Model for the Small Supersonic Experimental Aircraft (United States)

    Saitoh, Kenichi; Tamayama, Masato; Kikuchi, Takao; Machida, Shigeru; Nakamichi, Jiro

    This paper reports a wind-tunnel experiment and analysis that have been conducted under the National Experimental Airplane for Supersonic Transports (NEXST-1) project of JAXA. In order to perform the flight experiment, the design of the vehicle was examined from the stand point of aeroelasticity. The aileron buzz as well as flutter was of much concern for its aileron system on the main wing. Therefore, both wind-tunnel test and analysis were carried out by using a semi-span model with fuselage. Although the buzz was not observed in the test, damping responses of the aileron rotation mode were obtained. Critical damping was observed in supersonic flow, that meant a buzz could occur in ``region C'' of Lambourne's classification. Linear unsteady aerodynamic analysis is applicable to this type of buzz and the characteristics of the buzz of the model is discussed.

  10. Failure Analysis of T-38 Aircraft Burst Hydraulic Aileron Return Line (United States)

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


    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. Investigation of the Three-Dimensional Hinge Moment Characteristics Generated by the ONERA-M6 Wing with an Aileron

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    G. Q. Zhang


    Full Text Available The hinge moment characteristics for ONERA-M6 wing with aileron configuration have been investigated numerically based on the different gaps and deflecting angles. The results show that the effects on the wing made by the deflecting aileron are notable. Comparing with the nonaileron case, the chordwise pressure coefficient distribution for the wing with aileron has shown the totally different trends. The small gap can force the air flow through and form the extremely strong spraying flow. It can directly destroy the previously formed leading edge vortex (LEV. Due to the presence of the positive deflecting angle, the trailing edge vortex (TEV will begin to generate at the trailing edge of the aileron. The induced secondary LEV will be mixed with the developing TEVs and form the stronger TEVs at the downstream position. Comparing with the subsonic flow, the curve for the supersonic flow has shown a good linear. The corresponding hinge moments are also extremely sensitive to the changing angle of attack, and the slope of curves is also bigger than that of the subsonic flow. The bigger gap and deflecting angle can result in the curve of hinge moment bending upward at high angle of attack. The corresponding pressure cloud and streamlines have also been obtained computationally and analyzed in detail.

  12. Collection of Balanced-Aileron Test Data (United States)


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  13. 带副翼三维机翼绕流的Euler方程解%Euler Solution of a 3D Wing with Aileron

    Institute of Scientific and Technical Information of China (English)

    李津; 陈泽明; 朱自强; 李海明


    提出了一种求解带副翼偏转三维机翼绕流的Euler方程计算方法.采用对接分区网格与分区求解算法的结合,有效地求解绕此外形的复杂流动.提出了一种满足通量守恒的内边界耦合条件.数值方法中选用Van Leer分裂格式离散无粘通量项,并构造了一种Limiter函数以保证TVD性质.数值算例表明本方法是求解带操纵面偏转的机翼绕流的有效方法.

  14. Euler Solution of a 3D Wing-Body Configuration with Aileron%带副翼的翼身组合体绕流的Euler方程解

    Institute of Scientific and Technical Information of China (English)

    李津; 朱自强; 吴宗成; 陈泽民


    提出了一种求解带副翼偏转的翼身组合体绕流的Euler 方程计算方法.将对接分区网格与分区求解算法相结合,有效地求解了绕此外形的复杂流动.提出了一种满足通量守恒的内边界耦合条件.数值方法中选用Van Leer 分裂格式离散无粘通量项,并构造了一种限制器(Limiter)函数以保证TVD性质.数值算例表明本文方法是求解带操纵面偏转的翼身组合体绕流的有效方法.


    Institute of Scientific and Technical Information of China (English)

    陈泽民; 李津; 朱自强; 吴宗成


    将对接分区网格与分区求解算法结合,有效地求解了带副翼偏转的翼身组合体绕流的N-S方程.数值方法中选用Van Leer分裂格式离散无粘通量项,采用中心差分格式来离散粘性通量项.分区交界面采用了一种满足通量守恒的内边界耦合条件.数值算例表明该方法是求解带操纵面偏转的翼身组合体绕流的有效方法.

  16. Commande en position et vitesse sans capteur mécanique de moteurs synchrones à aimants permanents à pôles lisses : Application à un actionneur électromécanique pour aileron


    Zgorski, Aloïs


    The issue of sensorless control of permanent magnet synchronous machine (PMSM) has been well studied in the field of automation and electrical engineering. In the following work, we were interested in the peculiar problem of low-speed sensorless control of surface PMSM. The objective is to o er a new method for the position control of an industrial electro-mechanical actuator, considering an aeronautical context (a flap actuator). First, modeling and observability analyses of surface (non-sal...

  17. 75 FR 12468 - Airworthiness Directives; Quartz Mountain Aerospace, Inc. Model 11E Airplanes (United States)


    ... With the next 10 Follow Quartz aileron pushrod bearings. hours time-in- Mountain Aerospace service (TIS... aileron Within 50 hours TIS Follow Quartz pushrod bearings. after the cleaning Mountain Aerospace and... Aerospace, Inc. Model 11E Airplanes AGENCY: Federal Aviation Administration (FAA), Department...

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

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    Mohamed Sadok GUEZGUEZ


    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.

  19. Dynamic Model and Analysis of Asymmetric Telescopic Wing for Morphing Aircraft

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    Chen Lili


    Full Text Available Morphing aircraft has been the research hot topics of new concept aircrafts in aerospace engineering. Telescopic wing is an important morphing technology for morphing aircraft. This paper describes the dynamic equations and kinematic equations based on theorem of momentum and theorem of moment of momentum, which are available for all morphing aircrafts. Meanwhile,as simplified , dynamic equations for rectangular telescopic wing are presented. In order to avoid the complexity using aileron to generate rolling moment , an new idea that asymmetry of wings can generate roll moment is introduced. Finally, roll performance comparison of asymmetric wing and aileron deflection shows that asymmetric telescopic wing can provide the required roll control moment as aileron, and in some cases, telescopic wing has the superior roll performance.

  20. Extraction from flight data of lateral aerodynamic coefficients for F-8 aircraft with supercritical wing (United States)

    Williams, J. L.; Suit, W. T.


    A parameter-extraction algorithm was used to determine the lateral aerodynamic derivatives from flight data for the F-8 aircraft with supercritical wing. The flight data used were the recorded responses to aileron or rudder pulses for Mach numbers of 0.80, 0.90, and 0.98. Results of this study showed that a set of derivatives were determined which yielded a calculated aircraft response almost identical with the response measured in flight. Derivatives extracted from motion resulting from rudder inputs were somewhat different from those resulting from aileron inputs. It was found that the derivatives obtained from the rudder-input data were highly correlated in some instances. Those from the aileron input had very low correlations and appeared to be the more reliable.

  1. Flying Qualities Flight Testing of Digital Flight Control Systems. Flight Test Techniques Series - Volume 21 (les Essais en vol des performances des systemes de ommande de vol numeriques) (United States)


    et leur sécurité. Note de traduction : l’auteur insiste lourdement dans le 2ème paragraphe sur la préparation des essais et l’analyse loop is readily apparent. As an example, a roll-rate feedback to the aileron is a good indication that the roll mode required augmentation. The...impacts of roll-rate feedback to the aileron on this mode are well understood from the classical design theory. Newer or so-called modern design

  2. Experimental Investigation into the Control and Load Alleviation Capabilities of Articulated Winglets

    Directory of Open Access Journals (Sweden)

    A. Gatto


    load. Results presented in this paper do provide good evidence of the concept's ability to adequately perform both tasks, although for the current chosen wing/winglet configuration, roll control authority was unable to achieve, per unit of control surface deflection, the same level of performance set by modern aileron-based roll control methodologies.

  3. 75 FR 7936 - Airworthiness Directives; SCHEIBE-Flugzeugbau GmbH Model SF 25C Gliders (United States)


    ... the area located at the lower side of the wing), disconnect the aileron from the wings, disconnect the... reference of this service information under 5 U.S.C. 552(a) and 1 CFR part 51. (2) For service information... reference for this AD at the National Archives and Records Administration (NARA). For information on...

  4. Wind turbine control systems: Dynamic model development using system identification and the fast structural dynamics code

    Energy Technology Data Exchange (ETDEWEB)

    Stuart, J.G.; Wright, A.D.; Butterfield, C.P.


    Mitigating the effects of damaging wind turbine loads and responses extends the lifetime of the turbine and, consequently, reduces the associated Cost of Energy (COE). Active control of aerodynamic devices is one option for achieving wind turbine load mitigation. Generally speaking, control system design and analysis requires a reasonable dynamic model of {open_quotes}plant,{close_quotes} (i.e., the system being controlled). This paper extends the wind turbine aileron control research, previously conducted at the National Wind Technology Center (NWTC), by presenting a more detailed development of the wind turbine dynamic model. In prior research, active aileron control designs were implemented in an existing wind turbine structural dynamics code, FAST (Fatigue, Aerodynamics, Structures, and Turbulence). In this paper, the FAST code is used, in conjunction with system identification, to generate a wind turbine dynamic model for use in active aileron control system design. The FAST code is described and an overview of the system identification technique is presented. An aileron control case study is used to demonstrate this modeling technique. The results of the case study are then used to propose ideas for generalizing this technique for creating dynamic models for other wind turbine control applications.

  5. Extreme Agility Micro Aerial Vehicle - Control of Hovering Maneuvers for a Mini-Aerial Vehicle with an Onboard Autopilot System (United States)


    montré que les commandes d’accélérateur ont compensé les effets des ailerons par une fonction d’anticipation, réduit l’erreur de tenue d’altitude à...expérimentales ont montré que les commandes d’accélérateur qui ont été compensées pour les effets des ailerons par une fonction d’anticipation, a réduit l’erreur...with a predictive error method found in the pro- cess identification tool from the Matlab system identification toolbox. A sampling time of 0.05 s was

  6. Active Control Synthesis for Flexible Vehicles. Volume I. KONPACT theoretical Description (United States)


    scx ^;^6 aileron de,leotion to reduce b^— K = aileron feedback gains This equation generates additional constraints on...8217 Ml’ E«9• 519 * 509* 59* * 52» • S«1’ 511’ 501’ S9f I S2t" I W I 512’ SCZ ’ 591’ 521’ S00’ S10’ 500’« M..i;«r"«:’U;;«;r-«;’^;;;;r...J9t M E2I ’T ;p" M SUM :9li• 526- SPP ■ S9i- ??/• EB9• 519- SC91 E05. SP«?1 chij • SCl- £9f S2t • sez- S»2-

  7. Flight investigation of the effect of control centering springs on the apparent spiral stability of a personal-owner airplane (United States)

    Campbell, John P; Hunter, Paul A; Hewes, Donald E; Whitten, James B


    Report presents the results of a flight investigation conducted on a typical high-wing personal-owner airplane to determine the effect of control centering springs on apparent spiral stability. Apparent spiral stability is the term used to describe the spiraling tendencies of an airplane in uncontrolled flight as affected both by the true spiral stability of the perfectly trimmed airplane and by out-of-trim control settings. Centering springs were used in both the aileron and rudder control systems to provide both a positive centering action and a means of trimming the airplane. The springs were preloaded so that when they were moved through neutral they produced a nonlinear force gradient sufficient to overcome the friction in the control surface at the proper setting for trim. The ailerons and rudder control surfaces did not have trim tabs that could be adjusted in flight.

  8. Effect of winglets on performance and handling qualities of general aviation aircraft (United States)

    Van Dam, C. P.; Holmes, B. J.; Pitts, C.


    Recent flight and wind tunnel evaluations of winglets mounted on general aviation airplanes have shown improvements in cruise fuel efficiency, and climbing and turning performance. Some of these analyses have also uncovered various effects of winglets on airplane handling qualities. Retrofitting an airplane with winglets can result in reduced cross wind take-off and landing capabilities. Also, winglets can have a detrimental effect on the lateral directional response characteristics of aircraft which have a moderate to high level of adverse yaw due to aileron. Introduction of an aileron-rudder-interconnect, and reduction of the effective dihedral by canting-in of the winglets, or addition of a lower winglet can eliminate these flying quality problems.

  9. Peak-Seeking Optimization of Spanwise Lift Distribution for Wings in Formation Flight (United States)

    Hanson, Curtis E.; Ryan, Jack


    A method is presented for the in-flight optimization of the lift distribution across the wing for minimum drag of an aircraft in formation flight. The usual elliptical distribution that is optimal for a given wing with a given span is no longer optimal for the trailing wing in a formation due to the asymmetric nature of the encountered flow field. Control surfaces along the trailing edge of the wing can be configured to obtain a non-elliptical profile that is more optimal in terms of minimum combined induced and profile drag. Due to the difficult-to-predict nature of formation flight aerodynamics, a Newton-Raphson peak-seeking controller is used to identify in real time the best aileron and flap deployment scheme for minimum total drag. Simulation results show that the peak-seeking controller correctly identifies an optimal trim configuration that provides additional drag savings above those achieved with conventional anti-symmetric aileron trim.

  10. Enhancement of roll maneuverability using post-reversal design (United States)

    Li, Wei-En

    This dissertation consists of three main parts. The first part is to discuss aileron reversal problem for a typical section with linear aerodynamic and structural analysis. The result gives some insight and ideas for this aeroelastic problem. Although the aileron in its post-reversal state will work the opposite of its design, this type of phenomenon as a design root should not be ruled out on these grounds alone, as current active flight-control systems can compensate for this. Moreover, one can get considerably more (negative) lift for positive flap angle in this unusual regime than positive lift for positive flap angle in the more conventional setting. This may have important implications for development of highly maneuverable aircraft. The second part is to involve the nonlinear aerodynamic and structural analyses into the aileron reversal problem. Two models, a uniform cantilevered lifting surface and a rolling aircraft with rectangular wings, are investigated here. Both models have trailing-edge control surfaces attached to the main wings. A configuration that reverses at a relatively low dynamic pressure and flies with the enhanced controls at a higher level of effectiveness is demonstrated. To evaluate how reliable for the data from XFOIL, the data for the wing-aileron system from advanced CFD codes and experiment are used to compare with that from XFOIL. To enhance rolling maneuverability for an aircraft, the third part is to search for the optimal configuration during the post-reversal regime from a design point of view. Aspect ratio, hinge location, airfoil dimension, inner structure of wing section, composite skin, aeroelastic tailoring, and airfoil selection are investigated for cantilevered wing and rolling aircraft models, respectively. Based on these parametric structural designs as well as the aerodynamic characteristics of different airfoils, recommendations are given to expand AAW flight program.

  11. Innovative Control Effectors (ICE) (United States)


    including weight, maneuver performance, signa- ture, hydraulic requirements, demands on the flight control system (FCS) design, and car - rier (CV...applicable to the car - rier-based configurations. Figure 7-36 summarizes an assessment of the ICE series 101 configuration control allocation evaluation. ICE...plain leading edge flaps, all moving horizontal tails, rudder, two airbrakes under fuselage F-15C inner trailing edge plain flap, outer aileron, all

  12. A Business Overview & Summary of the SM-27S/T MACHETE RDT&E Program as Undertaken by the Military Aerospace/Tactical Air Warfare Systems Division of STAVATTI (United States)


    and winglets . The wing is a fail-safe structure with two spars. Each wing is equipped with ailerons, spoilers and double slotted Fowler flaps. The wings...high temperature graphite/ polyimide and chromium/SPECTRA metal matrix composite armor, the SM-27 has an sophisticated struc- tural composition ...and maximum thrust than the MACHETE. Although equipped with an UPCO ejection seat, the mold-less composite construction technique employed throughout

  13. Effects of Boundary Layer Flow Control Using Plasma Actuator Discharges (United States)


    have run tests in this area to demonstrate plasma actuators as ailerons and winglets , as well as to reduce separation on low pressure turbine (LPT...ray component of the SEM computed the elemental composition percentages of the plate. For aluminum oxide, a 3-to-2 ratio of oxygen to aluminum was...desired. However, the electron microscopy revealed that manganese was present in the composition , due to impurities in the aluminum. Figure 13

  14. Application of winglets and/or wing tip extensions with active load control on the Boeing 747 (United States)

    Allison, R. L.; Perkin, B. R.; Schoenman, R. L.


    The application of wing tip modifications and active control technology to the Boeing 747 airplane for the purpose of improving fuel efficiency is considered. Wing tip extensions, wing tip winglets, and the use of the outboard ailerons for active wing load alleviation are described. Modest performance improvements are indicated. A costs versus benefits approach is taken to decide which, if any, of the concepts warrant further development and flight test leading to possible incorporation into production airplanes.

  15. Dynamic Model and Analysis of Asymmetric Telescopic Wing for Morphing Aircraft



    Morphing aircraft has been the research hot topics of new concept aircrafts in aerospace engineering. Telescopic wing is an important morphing technology for morphing aircraft. This paper describes the dynamic equations and kinematic equations based on theorem of momentum and theorem of moment of momentum, which are available for all morphing aircrafts. Meanwhile,as simplified , dynamic equations for rectangular telescopic wing are presented. In order to avoid the complexity using aileron to ...

  16. Experimental testing of spanwise morphing trailing edge concept (United States)

    Pankonien, Alexander; Inman, Daniel J.


    Aircraft wings with smooth, hinge-less morphing ailerons exhibit increased chordwise aerodynamic efficiency over conventional hinged ailerons. Ideally, the wing would also use these morphing ailerons to smoothly vary its airfoil shape between spanwise stations to optimize the lift distribution and further increase aerodynamic efficiency. However, the mechanical complexity or added weight of achieving such a design has traditionally exceeded the potential aerodynamic gains. By expanding upon the previously developed cascading bimorph concept, this work uses embedded Macro-Fiber Composites and a flexure box mechanism, created using multi-material 3D printing, to achieve the Spanwise Morphing Trailing Edge (SMTE) concept. The morphing actuators are spaced spanwise along the wing with an elastomer spanning the gaps between them, which allows for optimization of the spanwise lift distribution while maintaining the continuity and efficiency of the morphing trailing edge. The concept is implemented in a representative section of a UAV wing with a 305 mm chord. A novel honeycomb skin is created from an elastomeric material using a 3D printer. The actuation capabilities of the concept are evaluated with and without spanning material on a test stand, free of aerodynamic loads. In addition, the actuation restrictions of the spanning elastomer, necessary in adapting the morphing concept from 2D to 3D, are characterized. Initial aerodynamic results from the 1'×1' wind-tunnel also show the effects of aerodynamic loading on the actuation range of the SMTE concept for uniform morphing.

  17. Low-speed aerodynamic performance of an aspect-ratio-10 supercritical-wing transport model equipped with a full-span slat and part-span and full-span double-slotted flaps (United States)

    Morgan, H. L., Jr.


    An investigation was conducted in the Langley 4 by 7 Meter Tunnel to determine the static longitudinal and lateral directional aerodynamic characteristics of an advanced aspect ratio 10 supercritical wing transport model equipped with a full span leading edge slat as well as part span and full span trailing edge flaps. This wide body transport model was also equipped with spoiler and aileron roll control surfaces, flow through nacelles, landing gear, and movable horizontal tails. Six basic wing configurations were tested: (1) cruise (slats and flaps nested), (2) climb (slats deflected and flaps nested), (3) part span flap, (4) full span flap, (5) full span flap with low speed ailerons, and (6) full span flap with high speed ailerons. Each of the four flapped wing configurations was tested with leading edge slat and trailing edge flaps deflected to settings representative of both take off and landing conditions. Tests were conducted at free stream conditions corresponding to Reynolds number of 0.97 to 1.63 x 10 to the 6th power and corresponding Mach numbers of 0.12 to 0.20, through an angle of attack range of 4 to 24, and a sideslip angle range of -10 deg to 5 deg. The part and full span wing configurations were also tested in ground proximity.

  18. Investigation of Aerodynamic Characteristics of Vee Tail%V形尾翼的气动特性研究

    Institute of Scientific and Technical Information of China (English)

    孔繁美; 邱栋


    Experimental investigation of the aerodynamic design of the vee tail featuring low RCS has been carried out. The effects of the wing dihedral angles, and aileron positions on the aerodynamic characteristics of the complete aircraft with vee tail are studied. Aircraft models with two types of vee tail and two conventional tails have been tested and the longitudinal/lateral performances compared. The results show that off-shipping moment curves are caused nonlinear by the wing with dihedral, and that the same order of off-shipping efficiency of aileron as that of its rolling efficiency is caused by the deflection of inboard aileron. Therefore, it is an important criterion that vee tail should avoid the effect of the asymmetric downwash in aerodynamic design.%为了研究具有良好隐身特性的V形尾翼的气动设计准则,通过风洞实验,探讨了机翼不同上反角和副翼位置对V形尾翼的飞机全机气动特性的影响.比较了两种V形尾翼与常规尾翼的纵、横向气动特性.研究结果表明,机翼上反引起全机偏航力矩曲线呈非线性;靠近翼根的副翼偏转引起副翼偏航效率与其滚转效率同样量级.因此,设法使尾翼避开不对称的下洗流场的影响是V形尾翼设计的一项重要准则.

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

    CERN Document Server

    Annaz, Fawaz Yahya


    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

  20. Structural testing for static failure, flutter and other scary things (United States)

    Ricketts, R. H.


    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.

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

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


    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.

  2. Smart Materials and Structures-Smart Wing. Volumes 1, 2, 3 and 4 (United States)


    have a spacing of approximately 1 in. Because ULTEM  thermoplastic was used for the close-out on both the flap and aileron (due to the need to...trailing edge flap with ULTEM  close-out strip and attachment holes with inserts installed. The dark lines in the are the SMA wires IV-D-213 19” 8...Trailing Edge Flap Attachment Hole with Insert Installed Fiber Optic Position Sensor SMA Wires ULTEM Close-out Strip Figure D9. SMA Trailing Edge

  3. A study of the use of experimental stability derivatives in the calculation of the lateral disturbed motions of a swept-wing airplane and comparison with flight results (United States)

    Bird, John D; Jaquet, Byron M


    An investigation was made to determine the accuracy with which the lateral flight motions of a swept-wing airplane could be predicted from experimental stability derivatives, determined in the 6-foot-diameter rolling-flow test section and 6 by 6-foot curved-flow test section of the Langley stability tunnel. In addition, determination of the significance of including the nonlinear aerodynamic effects of sideslip in the calculations of the motions was desired. All experimental aerodynamic data necessary for prediction of the lateral flight motions are presented along with a number of comparisons between flight and calculated motions caused by rudder and aileron disturbances.

  4. An approach to the synthesis of separate surface automatic flight control systems. (United States)

    Roskam, J.; Henry, S.


    A method is presented for the analysis of separate surface automatic flight control systems. The feasibility of such systems is demonstrated by the analysis of an example system, a separate surface wing-leveler for a Cessna 172. This example system employs a separate surface aileron with 15% of the basic airplane roll control power. A 90% reduction in bank-angle gust response can be obtained when compared with the basic airplane. The system does not feed back to the pilot's wheel. When failed (even hardover) the pilot retains more than adequate control of the airplane.

  5. Program Manager: Journal of the Defense Systems Management College. Volume 17, Number 2, March-April 1988, (United States)


    the crops distribution. Today, free- maket retailers In- it grows, and a share of profits from the clude both proprietors and collectives. sale of Its...Second, the recent changes in China’s political and social at- M ’ "" ’ ,t’n ’ mosphere, when combined with its develop- ment and modernization goals, have...publishes the 800 parts, such as ailerons, stabilizers, hat- newpapers. It is the sole source of news for ches, and water tanks. These parts are all media

  6. Investigation of the cross-ship comparison monitoring method of failure detection in the HIMAT RPRV. [digital control techniques using airborne microprocessors (United States)

    Wolf, J. A.


    The Highly maneuverable aircraft technology (HIMAT) remotely piloted research vehicle (RPRV) uses cross-ship comparison monitoring of the actuator RAM positions to detect a failure in the aileron, canard, and elevator control surface servosystems. Some possible sources of nuisance trips for this failure detection technique are analyzed. A FORTRAN model of the simplex servosystems and the failure detection technique were utilized to provide a convenient means of changing parameters and introducing system noise. The sensitivity of the technique to differences between servosystems and operating conditions was determined. The cross-ship comparison monitoring method presently appears to be marginal in its capability to detect an actual failure and to withstand nuisance trips.

  7. Optimal Topology of Aircraft Rib and Spar Structures under Aeroelastic Loads (United States)

    Stanford, Bret K.; Dunning, Peter D.


    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.

  8. Experimental study of UTM-LST generic half model transport aircraft (United States)

    Ujang, M. I.; Mat, S.; Perumal, K.; Mohd. Nasir, M. N.


    This paper presents the experimental results from the investigation carried out at the UTM Low Speed wind tunnel facility (UTM-LST) on a half model generic transport aircraft at several configurations of primary control surfaces (flap, aileron and elevator). The objective is to measure the aerodynamic forces and moments due to the configuration changes. The study is carried out at two different speeds of 26.1 m/s and 43.1 m/s at corresponding Reynolds number of 1 × 106 and 2 × 106, respectively. Angle of attack of the model is varied between -2o to 20o. For the flaps, the deflection applied is 0o, 5o and 10o. Meanwhile, for aileron and elevator, the deflection applied is between -10o and 10o. The results show the differences in aerodynamic characteristics of the aircraft at different control surfaces configurations. The results obtained indicate that a laminar separation bubble developed on the surface of the wing at lower angles of attack and show that the separation process is delayed when the Reynolds number is increased.

  9. Fuzzy Logic Decoupled Lateral Control for General Aviation Airplanes (United States)

    Duerksen, Noel


    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.

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

    Directory of Open Access Journals (Sweden)

    Andreea KOREANSCHI


    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.

  11. Simulator study of the effectiveness of an automatic control system designed to improve the high-angle-of-attack characteristics of a fighter airplane (United States)

    Gilbert, W. P.; Nguyen, L. T.; Vangunst, R. W.


    A piloted, fixed-base simulation was conducted to study the effectiveness of some automatic control system features designed to improve the stability and control characteristics of fighter airplanes at high angles of attack. These features include an angle-of-attack limiter, a normal-acceleration limiter, an aileron-rudder interconnect, and a stability-axis yaw damper. The study was based on a current lightweight fighter prototype. The aerodynamic data used in the simulation were measured on a 0.15-scale model at low Reynolds number and low subsonic Mach number. The simulation was conducted on the Langley differential maneuvering simulator, and the evaluation involved representative combat maneuvering. Results of the investigation show the fully augmented airplane to be quite stable and maneuverable throughout the operational angle-of-attack range. The angle-of-attack/normal-acceleration limiting feature of the pitch control system is found to be a necessity to avoid angle-of-attack excursions at high angles of attack. The aileron-rudder interconnect system is shown to be very effective in making the airplane departure resistant while the stability-axis yaw damper provided improved high-angle-of-attack roll performance with a minimum of sideslip excursions.

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

    Directory of Open Access Journals (Sweden)

    Fendy Santoso


    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.

  13. Stability investigation of an airfoil section with active flap control

    DEFF Research Database (Denmark)

    Bergami, Leonardo; Gaunaa, Mac


    This work presents a method to determine flutter and divergence instability limits for a two-dimensional (2-D) airfoil section fitted with an actively controlled trailing edge flap. This flap consists of a deformable trailing edge, which deformation is governed by control algorithms based...... for fatigue load alleviation. The structural model of the 2-D airfoil section contains three degrees of freedom: heave translation, pitch rotation and flap deflection. A potential flow model provides the aerodynamic forces and their distribution. The unsteady aerodynamics are described using an indicial...... function approximation. Stability of the full aeroservoelastic system is determined through eigenvalue analysis by state-space formulation of the indicial approximation. Validation is carried out against an implementation of the recursive method by Theodorsen and Garrick for flexure-torsion-aileron flutter...

  14. Initial Acceleration Suppression via Gust Alleviation Controller Using Prior Gust Information (United States)

    Sato, Masayuki; Yokoyama, Nobuhiro

    This paper first shows the performance limit of gust alleviation controllers without prior gust information. Since control devices, such as elevator, aileron, and thrust, cannot be driven at the same time as their commands are given, the initial acceleration due to wind gust cannot be alleviated with any controllers which do not detect the wind gust a priori. However, if wind gust information is obtained a priori, then the initial acceleration can be suppressed with some controller, which is the second topic of this paper. Designing such controllers is formulated as a design problem of model predictive controllers, which is formulated as a convex Quadratic Programming (QP) problem and is easily solved numerically. Several numerical examples demonstrate the effectiveness of model predictive controllers to suppress the initial acceleration due to wind gust.

  15. Composite NDE using full-field pulse-echo ultrasonic propagation imaging system (United States)

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


    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.

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

    Directory of Open Access Journals (Sweden)

    Mitsuo Uno


    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.

  17. Results of investigations on a 0.0405 scale model PRR version of the NR-SSV orbiter in the North American Aeronautical Laboratory low speed wind tunnel (United States)

    Kingsland, R. B.; Vaughn, J. E.; Singellton, R.


    Experimental aerodynamic investigations were conducted in a low speed wind tunnel on a scale model space shuttle vehicle (SSV) orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional aerodynamic characteristics of the space shuttle orbiter. Emphasis was placed on model component, wing-glove, and wing-body fairing effects, as well as elevon, aileron, and rudder control effectiveness. Angles of attack from - 5 deg to + 30 deg and angles of sideslip of - 5 deg, 0 deg, and + 5 deg were tested. Static pressures were recorded on base, fuselage, and wing surfaces. Tufts and talc-kerosene flow visualization techniques were also utilized. The aerodynamic force balance results are presented in plotted and tabular form.

  18. Experimental Results of Winglets on First, Second, and Third Generation Jet Transports. [to reduce drag coefficient (United States)

    Flechner, S. G.; Jacobs, P. F.


    Results of wind tunnel investigations of four jet transport configurations representing both narrow and wide-body configurations and also a future advanced aerodynamic configuration are presented including performance and wing root bending moment data. The effects of winglets on the aerodynamic characteristics throughout the flight envelope were studied. The results indicate that winglets improved the cruise lift to drag ratio between 4 and 8 percent, depending on the transport configuration. The data also indicate that ratios of relative aerodynamic gain to relative structural weight penalty for winglets are 1.5 to 2.5 times those for wing-tip extensions. Over the complete range of flight conditions, winglets produce no adverse effects on buffet onset, lateral-directional stability, and aileron control effectiveness.

  19. Experimental results of winglets on first, second, and third generation jet transports (United States)

    Flechner, S. G.; Jacobs, P. F.


    The results of wind tunnel investigations of winglets on four jet transport configurations are presented. Performance and wing root bending moment data were given. Additionally, detailed aerodynamic characteristics are presented at the design condition and also at several off design conditions for one configuration. Results of the investigations indicate that the winglets improve the cruise lift to drag ratio between 4 and 8 percent. These data also show that the ratios of relative aerodynamic gain to relative structural weight penalty for winglets are 1.5 to 2.5 times the ratios for wing tip extensions. The comprehensive investigation of the effects of winglets indicated that winglets produce no adverse effects on buffet onset, lateral directional stability, and aileron control effectiveness.

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

    Directory of Open Access Journals (Sweden)

    Fadjar Rahino Triputra


    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.

  1. 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


    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%.

  2. Simulation to Flight Test for a UAV Controls Testbed (United States)

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


    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.

  3. Active control using control allocation for UAVs with seamless morphing wing (United States)

    Wang, Zheng-jie; Sun, Yin-di; Yang, Da-qing; Guo, Shi-jun


    In this paper, a small seamless morphing wing aircraft of MTOW=51 kg is investigated. The leading edge (LE) and trailing edge (TE) control surfaces are positioned in the wing section in span wise. Based on the studying results of aeroelastic wing characteristics, the controller should be designed depending on the flight speed. Compared with a wing of rigid hinged aileron, the morphing wing produces the rolling moment by deflecting the flexible TE and LE surfaces. An iteration method of pseudo-inverse allocation and quadratic programming allocation within the constraints of actuators have be investigated to solve the nonlinear control allocation caused by the aerodynamics of the effectors. The simulation results will show that the control method based on control allocation can achieve the control target.

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

    Institute of Scientific and Technical Information of China (English)


    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.

  5. Formulation d'un modele mathematique par des techniques d'estimation de parametres a partir de donnees de vol pour l'helicoptere Bell 427 et l'avion F/A-18 servant a la recherches en aeroservoelasticite (United States)

    Nadeau-Beaulieu, Michel

    In this thesis, three mathematical models are built from flight test data for different aircraft design applications: a ground dynamics model for the Bell 427 helicopter, a prediction model for the rotor and engine parameters for the same helicopter type and a simulation model for the aeroelastic deflections of the F/A-18. In the ground dynamics application, the model structure is derived from physics where the normal force between the helicopter and the ground is modelled as a vertical spring and the frictional force is modelled with static and dynamic friction coefficients. The ground dynamics model coefficients are optimized to ensure that the model matches the landing data within the FAA (Federal Aviation Administration) tolerance bands for a level D flight simulator. In the rotor and engine application, rotors torques (main and tail), the engine torque and main rotor speed are estimated using a state-space model. The model inputs are nonlinear terms derived from the pilot control inputs and the helicopter states. The model parameters are identified using the subspace method and are further optimised with the Levenberg-Marquardt minimisation algorithm. The model built with the subspace method provides an excellent estimate of the outputs within the FAA tolerance bands. The F/A-18 aeroelastic state-space model is built from flight test. The research concerning this model is divided in two parts. Firstly, the deflection of a given structural surface on the aircraft following a differential ailerons control input is represented by a Multiple Inputs Single Outputs linear model whose inputs are the ailerons positions and the structural surfaces deflections. Secondly, a single state-space model is used to represent the deflection of the aircraft wings and trailing edge flaps following any control input. In this case the model is made non-linear by multiplying model inputs into higher order terms and using these terms as the inputs of the state-space equations. In

  6. The Nolans Bore rare-earth element-phosphorus-uranium mineral system: geology, origin and post-depositional modifications (United States)

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


    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

  7. Aeroservoelastic model based active control for large civil aircraft

    Institute of Scientific and Technical Information of China (English)


    A modeling and control approach for an advanced configured large civil aircraft with aeroservoelasticity via the LQG method and control allocation is presented.Mathematical models and implementation issues for the multi-input/multi-output(MIMO) aeroservoelastic system simulation developed for a flexible wing with multi control surfaces are described.A fuzzy logic based optimization approach is employed to solve the constrained control allocation problem via intelligently adjusting the components of output vector and find a proper vector in the attainable moment set(AMS) autonomously.The basic idea is to minimize the L2 norm of error between the desired moment and achievable moment using the designing freedom provided by redundantly allocated actuators and control surfaces.Considering the constraints of control surfaces,in order to obtain acceptable performance of aircraft such as stability and maneuverability,the fuzzy weights are updated by the learning algorithm,which makes the closed-loop system self-adaptation.Finally,an application example of flight control designing for the advanced civil aircraft is discussed as a demonstration.The studies we have performed showed that the advanced configured large civil aircraft has good performance with the proper designed control law designed via the proposed approach.The gust alleviation and flutter suppression are applied with the synergetic effects of elevator,ailerons,equivalent rudders and flaps.The results show good closed loop performance and meet the requirement of constraint of control surfaces.

  8. 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


    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.

  9. Commande de vol non lineaire d'un drone a voilure fixe par la methode du backstepping (United States)

    Finoki, Edouard

    This thesis describes the design of a non-linear controller for a UAV using the backstepping method. It is a fixed-wing UAV, the NexSTAR ARF from HobbicoRTM. The aim is to find the expressions of the aileron, the elevator, and the rudder deflection in order to command the flight path angle, the heading angle and the sideslip angle. Controlling the flight path angle allows a steady, climb or descent flight, controlling the heading cap allows to choose the heading and annul the sideslip angle allows an efficient flight. A good technical control has to ensure the stability of the system and provide optimal performances. Backstepping interlaces the choice of a Lyapunov function with the design of feedback control. This control technique works with the true non-linear model without any approximation. The procedure is to transform intermediate state variables into virtual inputs which will control other state variables. Advantages of this technique are its recursivity, its minimum control effort and its cascaded structure that allows dividing a high order system into several simpler lower order systems. To design this non-linear controller, a non-linear model of the UAV was used. Equations of motion are very accurate, aerodynamic coefficients result from interpolations between several essential variables in flight. The controller has been implemented in Matlab/Simulink and FlightGear.

  10. DARPA/AFRL/NASA Smart Wing Second Wind Tunnel Test Results (United States)

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


    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.

  11. Squid rocket science: How squid launch into air (United States)

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


    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.

  12. Fault Tolerance Analysis of L1 Adaptive Control System for Unmanned Aerial Vehicles (United States)

    Krishnamoorthy, Kiruthika

    Trajectory tracking is a critical element for the better functionality of autonomous vehicles. The main objective of this research study was to implement and analyze L1 adaptive control laws for autonomous flight under normal and upset flight conditions. The West Virginia University (WVU) Unmanned Aerial Vehicle flight simulation environment was used for this purpose. A comparison study between the L1 adaptive controller and a baseline conventional controller, which relies on position, proportional, and integral compensation, has been performed for a reduced size jet aircraft, the WVU YF-22. Special attention was given to the performance of the proposed control laws in the presence of abnormal conditions. The abnormal conditions considered are locked actuators (stabilator, aileron, and rudder) and excessive turbulence. Several levels of abnormal condition severity have been considered. The performance of the control laws was assessed over different-shape commanded trajectories. A set of comprehensive evaluation metrics was defined and used to analyze the performance of autonomous flight control laws in terms of control activity and trajectory tracking errors. The developed L1 adaptive control laws are supported by theoretical stability guarantees. The simulation results show that L1 adaptive output feedback controller achieves better trajectory tracking with lower level of control actuation as compared to the baseline linear controller under nominal and abnormal conditions.

  13. Lamb wave-based BVID imaging for a curved composite sandwich panel (United States)

    He, Jiaze; Yuan, Fuh-Gwo


    Composite sandwich structures, consisting of a low density core sandwiched between two laminated facesheets, have been widely used in various aerospace structures. A new Lamb wave-based imaging condition, which will be referred to as the inverse incident wave energy (IIWE) imaging criterion, is proposed in this paper to resolve the situations where the incident wave energy weakly penetrates into the damaged area in the upper facesheet region. Current imaging conditions by analyzing wavefield reconstructed from laser Doppler vibrometer (LDV) scanning have been proven to be adequate for imaging damage in layered composite laminates. In this research, those current imaging conditions were applied and compared in the composite foam structures for barely visible impact damage (BVID). A piezoelectric wafer was used to excite Lamb waves into the structure and a LDV was used to scan the potential damaged areas in the upper facesheet of the panel. A BVID site in a curved composite sandwich foam aileron was inspected using various wavefield analysis methods and the damage images were compared with C-scan images. A few imaging conditions that are effective for this BVID site are identified when the incident waves have difficulties penetrating into the damaged region.

  14. Modelling and simulation of flight control electromechanical actuators with special focus on model architecting, multidisciplinary effects and power flows

    Directory of Open Access Journals (Sweden)

    Jian Fu


    Full Text Available In the aerospace field, electromechanical actuators are increasingly being implemented in place of conventional hydraulic actuators. For safety-critical embedded actuation applications like flight controls, the use of electromechanical actuators introduces specific issues related to thermal balance, reflected inertia, parasitic motion due to compliance and response to failure. Unfortunately, the physical effects governing the actuator behaviour are multidisciplinary, coupled and nonlinear. Although numerous multi-domain and system-level simulation packages are now available on the market, these effects are rarely addressed as a whole because of a lack of scientific approaches for model architecting, multi-purpose incremental modelling and judicious model implementation. In this publication, virtual prototyping of electromechanical actuators is addressed using the Bond-Graph formalism. New approaches are proposed to enable incremental modelling, thermal balance analysis, response to free-run or jamming faults, impact of compliance on parasitic motion, and influence of temperature. A special focus is placed on friction and compliance of the mechanical transmission with fault injection and temperature dependence. Aileron actuation is used to highlight the proposals for control design, energy consumption and thermal analysis, power network pollution analysis and fault response.

  15. 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.


    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.

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

    Directory of Open Access Journals (Sweden)

    Jane Maria Faulstich de Paiva


    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.

  17. Aerodynamic Characteristic of the Active Compliant Trailing Edge Concept (United States)

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


    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.

  18. Sensitivity Analysis of Linear Programming and Quadratic Programming Algorithms for Control Allocation (United States)

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


    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.

  19. Design for the Control of a Rotatable Stabiliser

    CERN Document Server

    Childs, S J


    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...

  20. The lateral-directional characteristics of a 74-degree Delta wing employing gothic planform vortex flaps (United States)

    Grantz, A. C.


    The low speed lateral/directional characteristics of a generic 74 degree delta wing body configuration employing the latest generation, gothic planform vortex flaps was determined. Longitudinal effects are also presented. The data are compared with theoretical estimates from VORSTAB, an extension of the Quasi vortex lattice Method of Lan which empirically accounts for vortex breakdown effects in the calculation of longitudinal and lateral/directional aerodynamic characteristics. It is indicated that leading edge deflections of 30 and 40 degrees reduce the magnitude of the wing effective dihedral relative to the baseline for a specified angle of attack or lift coefficient. For angles of attack greater than 15 degrees, these flap deflections reduce the configuration directional stability despite improved vertical tail effectiveness. It is shown that asymmetric leading edge deflections are inferior to conventional ailerons in generating rolling moments. VORSTAB calculations provide coarse lateral/directional estimates at low to moderate angles of attack. The theory does not account for vortex flow induced, vertical tail effects.

  1. Digital adaptive control laws for the F-8 (United States)

    Hartmann, G. L.; Harvey, C. A.


    NASA is conducting a flight control research program in digital fly-by-wire technology using a modified F-8C aircraft. The first phase of this program used Apollo hardware to demonstrate the practicality of digital fly-by-wire in an actual test vehicle. For the second phase, conventional aircraft sensors and a large floating point digital computer are being utilized to test advanced control laws and redundancy concepts. As part of NASA's research in digital fly-by-wire technology, Honeywell developed digital adaptive flight control laws for flight test in the F-8C. Adaptation of the control laws was to be based on information sensed from conventional aircraft sensors excluding air data. The control laws were constrained to use only existing elevator, rudder, and ailerons as control effectors, each powered by existing actuators. Three adaptive control laws were successfully designed using maximum likelihood estimation, a Liapunov stable model tracker and a self-excited limit cycle concept. The maximum likelihood estimation design was selected as the most promising because of its capability to identify more than surface effectiveness parameters. The adaptive concepts, the control laws and comparisons of predicted performance are described.

  2. Low-speed aerodynamic performance of a high-aspect-ratio supercritical-wing transport model equipped with full-span slat and part-span double-slotted flaps (United States)

    Morgan, H. L., Jr.; Paulson, J. W., Jr.


    An investigation was conducted in the Langley V/STOL tunnel to determine the static longitudinal and lateral-directional aerodynamic characteristics of an advanced high-aspect-ratio supercritical-wing transport model equipped with a full-span leading-edge slat and part-span double-slotted trailing-edge flaps. This wide-body transport model was also equipped with spoiler and aileron control surfaces, flow-through nacelles, landing gear, movable horizontal tails, and interchangeable wing tips with aspect ratios of 10 and 12. The model was tested with leading-edge slat and trailing-edge flap combinations representative of cruise, climb, takeoff, and landing wing configurations. The tests were conducted at free-stream conditions corresponding to Reynolds numbers (based on mean geometric chord) of 0.97 to 1.63 x 10 to the 6th power and corresponding Mach numbers of 0.12 to 0.20, through an angle-of-attack range of -2 deg to 24 deg and a sideslip-angle range of -10 deg to 5 deg.

  3. 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


    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

  4. Research on Safe Flight of the Outboard Wing Rupturing Aircraft based on CFD Simulation%基于CFD仿真的外翼断裂飞机安全飞行研究

    Institute of Scientific and Technical Information of China (English)

    姚武文; 蔡开龙


    The flight of the aircraft with outboard wing rupturing is a special problem about pneumatics and flight con-trol.The analysis model about the special pneumatics problem is set up by the CFD emulation method,and the control method about outboard wing rupturing aircraft aviating safely is brought forward.By the simulation computing,the conclu-sion is gained that the opposite damage most limit of the outboard wing based on the aileron balance method is 14.2%,and the opposite damage most limit of the outboard wing based on the side flight balance method is 24.5%.It provides the refer-ence and evaluation method for the aircraft with outboard wing rupturing flying.%针对外翼断裂飞机安全飞行问题,采用CFD 仿真方法,建立了针对该特殊气动问题的分析模型,提出了外翼断裂飞机安全飞行控制方法,并通过仿真计算得到,基于副翼平衡法的外翼相对损伤极限为14.2%,基于侧飞平衡法的外翼相对损伤极限为24.5%,为战时外翼断裂飞机带伤飞行提供了参考和评估方法。

  5. Multi-Axis Identifiability Using Single-Surface Parameter Estimation Maneuvers on the X-48B Blended Wing Body (United States)

    Ratnayake, Nalin A.; Koshimoto, Ed T.; Taylor, Brian R.


    The problem of parameter estimation on hybrid-wing-body type aircraft is complicated by the fact that many design candidates for such aircraft involve a large number of aero- dynamic control effectors that act in coplanar motion. This fact adds to the complexity already present in the parameter estimation problem for any aircraft with a closed-loop control system. Decorrelation of system inputs must be performed in order to ascertain individual surface derivatives with any sort of mathematical confidence. Non-standard control surface configurations, such as clamshell surfaces and drag-rudder modes, further complicate the modeling task. In this paper, asymmetric, single-surface maneuvers are used to excite multiple axes of aircraft motion simultaneously. Time history reconstructions of the moment coefficients computed by the solved regression models are then compared to each other in order to assess relative model accuracy. The reduced flight-test time required for inner surface parameter estimation using multi-axis methods was found to come at the cost of slightly reduced accuracy and statistical confidence for linear regression methods. Since the multi-axis maneuvers captured parameter estimates similar to both longitudinal and lateral-directional maneuvers combined, the number of test points required for the inner, aileron-like surfaces could in theory have been reduced by 50%. While trends were similar, however, individual parameters as estimated by a multi-axis model were typically different by an average absolute difference of roughly 15-20%, with decreased statistical significance, than those estimated by a single-axis model. The multi-axis model exhibited an increase in overall fit error of roughly 1-5% for the linear regression estimates with respect to the single-axis model, when applied to flight data designed for each, respectively.

  6. Rotary Balance Wind Tunnel Testing for the FASER Flight Research Aircraft (United States)

    Denham, Casey; Owens, D. Bruce


    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.

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


    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

  8. Design, realization and structural testing of a compliant adaptable wing (United States)

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


    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.

  9. Review on advanced composite materials boring mechanism and tools (United States)

    Shi, Runping; Wang, Chengyong


    With the rapid development of aviation and aerospace manufacturing technology, advanced composite materials represented by carbon fibre reinforced plastics (CFRP) and super hybrid composites (fibre/metal plates) are more and more widely applied. The fibres are mainly carbon fibre, boron fibre, Aramid fiber and Sic fibre. The matrixes are resin matrix, metal matrix and ceramic matrix. Advanced composite materials have higher specific strength and higher specific modulus than glass fibre reinforced resin composites of the 1st generation. They are widely used in aviation and aerospace industry due to their high specific strength, high specific modulus, excellent ductility, anticorrosion, heat-insulation, sound-insulation, shock absorption and high&low temperature resistance. They are used for radomes, inlets, airfoils(fuel tank included), flap, aileron, vertical tail, horizontal tail, air brake, skin, baseboards and tails, etc. Its hardness is up to 62~65HRC. The holes are greatly affected by the fibre laminates direction of carbon fibre reinforced composite material due to its anisotropy when drilling in unidirectional laminates. There are burrs, splits at the exit because of stress concentration. Besides there is delamination and the hole is prone to be smaller. Burrs are caused by poor sharpness of cutting edge, delamination, tearing, splitting are caused by the great stress caused by high thrust force. Poorer sharpness of cutting edge leads to lower cutting performance and higher drilling force at the same time. The present research focuses on the interrelation between rotation speed, feed, drill's geometry, drill life, cutting mode, tools material etc. and thrust force. At the same time, holes quantity and holes making difficulty of composites have also increased. It requires high performance drills which won't bring out defects and have long tool life. It has become a trend to develop super hard material tools and tools with special geometry for drilling

  10. Fluid induced microstructures in granulites from the Reynolds Range, central Australia (United States)

    Prent, Alexander; Beinlich, Andreas; Raimondo, Tom; Putnis, Andrew


    Fluids play a major role in the evolution of the Earth's crust, driving metamorphic reactions, facilitating transport of mass and heat, and changing the physical properties of rock. Shear zones present in intraplate orogens are ideal natural laboratories to study the relationship of fluid-driven rock weakening to deformation, and thus the impact of fluid availability on the tectonic reworking of continental interiors. Here we present preliminary observations from the Aileron Shear Zone (ASZ), Reynolds Range, central Australia, a major crustal-scale thrust of the Palaeozoic Alice Springs Orogen (ASO). This study focuses on the effects of fluids on the mineralogy and mineral chemistry of deep crustal rocks collected from a transect running through the ASZ. The ASZ is thought to have been of major importance during exhumation of the ASO, and exhumes a partly retrogressed suite of felsic and metasedimentary granulite facies gneisses. Hydration reactions associated with retrogression resulted in the partial replacement of orthopyroxene and numerous myrmekite textures associated with plagioclase and mica. In undeformed samples, orthopyroxene (En56 Fer44) rims are partly replaced by a zoned sequence of biotite (Phl70 Ann30), sub-parallel rims of magnetite, biotite and K-feldspar (Or87). Deformed samples gradually show an increase in dynamic recrystallization of quartz, with fully recrystallized bands of foam texture quartz defining the foliation together with biotite. Quartz and minor biotite replacement then dominates the mineral assemblage with increasing strain. The presence of fluid-driven mineral replacement reactions in undeformed samples suggests that hydration predates shearing and exhumation, and furthermore, that strain may have been localised in areas of intense hydration and rock weakening. Retrograde reactions and myrmekite textures suggest the availability of a silica-saturated fluid. Additional mass-balance calculations will be applied to constrain the

  11. A Method of Extremum Seeking Control Based on a Time Varying Kalman Filter and its Application to Formation Flight (United States)

    Rios-Zertuche, Rodolfo

    This dissertation presents a novel extremum seeking control method which combines a time-varying Kalman filter with a Newton Raphson algorithm. The Kalman filter is used to estimate the gradient and Hessian of a performance function. The resulting estimates are used in the Newton Raphson algorithm to guide the system to a local extremum of the performance function. Convergence of the method to a local extremum is proven when the system is subject to noisy measurements. This is accomplished by showing that the output of the algorithm is a supermartingale. It is shown that the system will converge to an area around the extremum with a radius defined, in part, by the error covariance of the Kalman filter estimates. The method is applied to two examples. The first utilizes a single independent parameter performance function. The second applies the method to the problem of formation flight for drag reduction. In the first example, two implementations of the method are examined. The first uses only gradient estimates. The second uses both gradient and Hessian estimates. Both implementations show good convergence in the presence of noisy measurements. The second example is of formation flight for drag reduction. The problem is described in some detail and includes an aerodynamic development of the drag-reduction phenomenon. The problem is explored with two simulations. The first uses coefficient of induced drag as its performance function and estimates the gradient and Hessian of the performance function. It shows good convergence of the method. The second simulation first uses pitch angle and then aileron deflection as its performance function. It estimates the gradient but not the Hessian of the performance function. It also shows good convergence.

  12. Digital Morphing Wing: Active Wing Shaping Concept Using Composite Lattice-Based Cellular Structures (United States)

    Jenett, Benjamin; Calisch, Sam; Cellucci, Daniel; Cramer, Nick; Gershenfeld, Neil; Swei, Sean


    Abstract We describe an approach for the discrete and reversible assembly of tunable and actively deformable structures using modular building block parts for robotic applications. The primary technical challenge addressed by this work is the use of this method to design and fabricate low density, highly compliant robotic structures with spatially tuned stiffness. This approach offers a number of potential advantages over more conventional methods for constructing compliant robots. The discrete assembly reduces manufacturing complexity, as relatively simple parts can be batch-produced and joined to make complex structures. Global mechanical properties can be tuned based on sub-part ordering and geometry, because local stiffness and density can be independently set to a wide range of values and varied spatially. The structure's intrinsic modularity can significantly simplify analysis and simulation. Simple analytical models for the behavior of each building block type can be calibrated with empirical testing and synthesized into a highly accurate and computationally efficient model of the full compliant system. As a case study, we describe a modular and reversibly assembled wing that performs continuous span-wise twist deformation. It exhibits high performance aerodynamic characteristics, is lightweight and simple to fabricate and repair. The wing is constructed from discrete lattice elements, wherein the geometric and mechanical attributes of the building blocks determine the global mechanical properties of the wing. We describe the mechanical design and structural performance of the digital morphing wing, including their relationship to wind tunnel tests that suggest the ability to increase roll efficiency compared to a conventional rigid aileron system. We focus here on describing the approach to design, modeling, and construction as a generalizable approach for robotics that require very lightweight, tunable, and actively deformable structures. PMID:28289574

  13. Experimental Aerodynamic Characteristics of an Oblique Wing for the F-8 OWRA (United States)

    Kennelly, Robert A., Jr.; Carmichael, Ralph L.; Smith, Stephen C.; Strong, James M.; Kroo, Ilan M.


    An experimental investigation was conducted during June-July 1987 in the NASA Ames 11-Foot Transonic Wind Tunnel to study the aerodynamic performance and stability and control characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing. This effort was part of the Oblique Wing Research Aircraft (OWRA) program performed in conjunction with Rockwell International. The Ames-designed, aspect ratio 10.47, tapered wing used specially designed supercritical airfoils with 0.14 thickness/chord ratio at the root and 0.12 at the 85% span location. The wing was tested at two different mounting heights above the fuselage. Performance and longitudinal stability data were obtained at sweep angles of 0deg, 30deg, 45deg, 60deg, and 65deg at Mach numbers ranging from 0.30 to 1.40. Reynolds number varied from 3.1 x 10(exp 6)to 5.2 x 10(exp 6), based on the reference chord length. Angle of attack was varied from -5deg to 18deg. The performance of this wing is compared with that of another oblique wing, designed by Rockwell International, which was tested as part of the same development program. Lateral-directional stability data were obtained for a limited combination of sweep angles and Mach numbers. Sideslip angle was varied from -5deg to +5deg. Landing flap performance was studied, as were the effects of cruise flap deflections to achieve roll trim and tailor wing camber for various flight conditions. Roll-control authority of the flaps and ailerons was measured. A novel, deflected wing tip was evaluated for roll-control authority at high sweep angles.

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

  15. Lateral Dynamic Stability of Aerial Refueling Aircraft and Tanker Aircraft%空中加油受油机横侧动操稳特性研究

    Institute of Scientific and Technical Information of China (English)

    宋双杰; 张玉莲


    空中加油时,加油机的气动力和力矩对受油机的操纵影响较大.需对影响大小进行定量分析,得出确定结论.本文对受油机的横侧动稳定性及操纵性进行了研究,确定了作用在受油机上由加油机翼端涡场引起的气动力和力矩的大小,提出一种简化气动模型,并将导数形式的力和力矩用于线性运动方程.研究表明,受油机具有发散的振荡特性,且主要由倾斜和侧向移动组成;飞行员在飞行时可通过不断操纵副翼来控制发散模态,保持飞机稳定.%During the air refueling, the aerodynamic force and the vortex field of the refueling tanker aircraft would greatly affect the manipulation of the refueling aircraft, and it is important to study this impact. This paper studies the cross-lateral dynamic stability and controllability of the air refueling tanker aircraft. The aerodynamic force and the vortex field of the refueling tanker aircraft on the wing of the refueling aircraft are determined through a simplified aerodynamic model. These forces and moments are expressed in a derivative form to be applied in the linear equations of motion. It is shown that the aerial refueling tanker aircraft would experience a divergent oscillation, mainly composed of tilt and lateral movement. By the manipulation of ailerons to control the divergence continuous mode, the aircraft may keep stable. The results of the present paper may serve as a theoretical basis for action that can be taken for other types of manipulation and other types of refueling aircraft.

  16. Analysis of Aircraft Control Performance using a Fuzzy Rule Base Representation of the Cooper-Harper Aircraft Handling Quality Rating (United States)

    Tseng, Chris; Gupta, Pramod; Schumann, Johann


    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.

  17. Intelligent adaptive nonlinear flight control for a high performance aircraft with neural networks. (United States)

    Savran, Aydogan; Tasaltin, Ramazan; Becerikli, Yasar


    This paper describes the development of a neural network (NN) based adaptive flight control system for a high performance aircraft. The main contribution of this work is that the proposed control system is able to compensate the system uncertainties, adapt to the changes in flight conditions, and accommodate the system failures. The underlying study can be considered in two phases. The objective of the first phase is to model the dynamic behavior of a nonlinear F-16 model using NNs. Therefore a NN-based adaptive identification model is developed for three angular rates of the aircraft. An on-line training procedure is developed to adapt the changes in the system dynamics and improve the identification accuracy. In this procedure, a first-in first-out stack is used to store a certain history of the input-output data. The training is performed over the whole data in the stack at every stage. To speed up the convergence rate and enhance the accuracy for achieving the on-line learning, the Levenberg-Marquardt optimization method with a trust region approach is adapted to train the NNs. The objective of the second phase is to develop intelligent flight controllers. A NN-based adaptive PID control scheme that is composed of an emulator NN, an estimator NN, and a discrete time PID controller is developed. The emulator NN is used to calculate the system Jacobian required to train the estimator NN. The estimator NN, which is trained on-line by propagating the output error through the emulator, is used to adjust the PID gains. The NN-based adaptive PID control system is applied to control three angular rates of the nonlinear F-16 model. The body-axis pitch, roll, and yaw rates are fed back via the PID controllers to the elevator, aileron, and rudder actuators, respectively. The resulting control system has learning, adaptation, and fault-tolerant abilities. It avoids the storage and interpolation requirements for the too many controller parameters of a typical flight control

  18. 小型无人机航路规划及自主导航算法研究%Research of MUAV route planning and autonomous navigation system

    Institute of Scientific and Technical Information of China (English)

    蒋志华; 陶德桂


    为了实现无人机自主导航功能,定义了以发射点为坐标原点的地理导航坐标系,给出由WGS⁃84坐标系到导航坐标系的具体转换公式;设计了航路中的航点结构,利用GPS接收机输出数据对无人机的位置进行两点式实时推算,对无人机在直线航路段和转弯段的判断和侧偏距进行了详细讨论,避免了超越函数的计算;利用比例⁃微分导航控制律求得无人机的滚转角,通过飞控系统控制无人机副翼舵机,修正飞行航迹,达到消除侧偏距的目的。这些方法和措施提高了小型无人机的自主导航控制精度和稳定性,使小型无人机具有较好的飞行品质。%In order to realize autonomous navigation function of UAV,a geographic navigation coordinates was defined, and the conversion formula from WGS⁃84 coordinates to navigation coordinates was offered. The navigating points in air route planning were designed. At each air⁃point,a two⁃point model was used for determine the position of the UAV by using data coming from GPS receiver. The judgment whether the UAV is in line route or turning section is discussed in detail,which avoid calcula⁃tion of transcendental function. The roll angle of the UAV was acquired by using a proportional⁃differential navigation control law. The aileron actuator is controlled by flight control system to modify the flight routine and eliminate the lateral offset. With these methods and measures,the navigation precision and stability of UAV was improved,which achieved a good flying quality of UAV.

  19. 复合材料后掠翼机翼气动弹性分析%Aeroelastic Characteristics Analysis of a Composite Backward-Swept Wing

    Institute of Scientific and Technical Information of China (English)

    周宏霞; 吕锁宁


    对于复合材料后掠翼机翼,扭转发散问题一般并不突出,其操纵面的操纵效率和颤振临界动压是比较关心的两个问题。文章采用COMPASS软件,对某复合材料后掠翼飞机进行了操纵效率分析,并重点计算了该机在不同高度下颤振速度随马赫数的变化情况,详细分析了机翼振动、颤振特性随蒙皮不同铺层比变化情况。结果表明,舵面操纵效率随着马赫数的增加而降低,飞机设计要通过设计参数调整选择合适的副翼反效动压与扭转发散动压之比,使飞行范围内的操纵效率尽可能高;同时复合材料后掠机翼的弯扭耦合效应相当突出,而复合材料剪裁可以调整0°、+45°、90°铺层比例,提高结构扭转刚度,从而提高飞机颤振速度。%For a composite material backward-swept wing, the control surface efficiency and flutter speed are more worth concerning compared with torsion divergence. The aeroelastic characteristics were calculated and analyzed by the COMPASS software, including the control efficiency of a wing separately, and various flutter speeds corresponding to different subsonic mach numbers were calculated emphatically. The results indicate that the control surface efficiency decreases as Mach number increases. The design parameters must be adjusted to obtain appropriate aileron reversal dynamics pressure to torsion radiation dynamics pressure ratio which makes the control efficiency higher. At the same time the bending and torsion coupling effect of composite material swept-back wing is quite severe. The composite material clipping can rectify the proportion of the 0°, +45° and 90° in order to enhance the structure torsion stiffness and the aircraft flutter speed.

  20. 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


    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。

  1. 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


    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.

  2. 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)


    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.

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

  4. German Contribution to the X-38 CRV Demonstrator in the Field of Guidance, Navigation and Control (GNC) (United States)

    Soppa, Uwe; Görlach, Thomas; Roenneke, Axel Justus


    As a solution to meet a safety requirement to the future full scale space station infrastructure, the Crew Return/Rescue Vehicle (CRV) was supposed to supply the return capability for the complete ISS crew of 7 astronauts back to earth in case of an emergency. A prototype of such a vehicle named X-38 has been developed and built by NASA with European partnership (ESA, DLR). An series of aerial demonstrators (V13x) for tests of the subsonic TAEM phase and the parafoil descent and landing system has been flown by NASA from 1998 to 2001. A full scale unmanned space flight demonstrator (V201) has been built at JSC Houston and although the project has been stopped for budgetary reasons in 2002, it will hopefully still be flown in near future. The X-38 is a lifting body with hypersonic lift to drag ratio about 0.9. In comparison to the Space Shuttle Orbiter, this design provides less aerodynamic maneuvrability and a different actuator layout (divided body flap and winglet rudders instead as combined aileron and elevon in addition to thrust- ers for the early re-entry phase). Hence, the guidance and control concepts used onboard the shuttle orbiter had to be adapted and further developed for the application on the new vehicle. In the frame of the European share of the X-38 project and also of the German TETRA (TEchnol- ogy for future space TRAnsportation) project different GNC related contributions have been made: First, the primary flight control software for the autonomous guidance and control of the X-38 para- foil descent and landing phase has been developed, integrated and successfully flown on multiple vehicles and missions during the aerial drop test campaign conducted by NASA. Second, a real time X-38 vehicle simulator was provided to NASA which has also been used for the validation of a European re-entry guidance and control software (see below). According to the NASA verification and validation plan this simulator is supposed to be used as an independent vali

  5. Adaptive Augmenting Control Flight Characterization Experiment on an F/A-18 (United States)

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


    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.

  6. AD-1 with research pilot Richard E. Gray (United States)


    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

  7. SR-71 Pilot Rogers E. Smith (United States)


    , such as angle of attack and sideslip, which are normally obtained with small tubes and vanes 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

  8. SR-71 Pilots and Crew (Smith, Meyer, Bohn-Meyer, Ishmael) (United States)


    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 been assigned the following NASA tail numbers: NASA 844 (A model), military

  9. SR-71 Research Engineer Marta Bohn-Meyer (United States)


    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

  10. SR-71 Pilots and Crew (Smith, Meyer, Bohn-Meyer, Ishmael) (United States)


    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

  11. M2-F1 on lakebed with pilots Milt Thompson, Chuck Yeager, Don Mallick, and Bruce Peterson (United States)


    -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

  12. Dale Reed with model in front of M2-F1 (United States)


    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