AIAA-RSC2-2003-U-010

Flight Dynamics and Control of an With Segmented Control Surfaces

Mujahid Abdulrahim Undergraduate University of Florida Gainesville, FL

AIAA 54th Southeastern Regional Student Conference March 27-28, 2003 Kill Devil Hills, NC

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FLIGHT DYNAMICS AND CONTROL OF AN AIRCRAFT WITH SEGMENTED CONTROL SURFACES

Mujahid Abdulrahim* University of Florida, Gainesville, Florida

Abstract

Flight researchers are increasingly turning towards small, unmanned aircraft for achieving mission objectives. These aircraft are simple to operate and offer numerous advantages over larger manned vehicles. In addition to being light, inexpensive, and readily available, they are also more versatile in that they can be used for flight experiments that are either too risky or uncertain for a manned flight test program. One application of unmanned vehicles is in the area of increased control authority research. This paper presents the preliminary stages of one such application, where an existing UAV is modified with 16 independent wing control surfaces. These surfaces are used in place of conventional for roll control and as a supplement to , , and controls. Instrumentation and sensors on-board the aircraft allow complete characterization of the flight dynamics. A traditional control system is replaced with a microcontroller that commands each segment independently. Various modes of actuation can be implemented to improve roll, pitch, and yaw response, minimize induced drag, and provide numerous levels of redundancy. The results indicate that the segmented control surfaces can be configured for a superior level of control.

INTRODUCTION A preliminary approach to designing a morphing Small, unmanned air vehicles are increasingly used as a vehicle is increasing the number control surfaces. This tool for flight research. Equipment and instrumentation research focuses on the development and that once was prohibitively large and expensive is now characterization of such an aircraft. The vehicle in available for these miniature aircraft. While the use of question is equipped with 16 independent wing control UAVs for research continues gaining acceptance, the surfaces in place of the conventional ailerons. capabilities of the individual research teams continue to Although actuated in a similar fashion, the large expand. No longer are flight researchers concerned number of surfaces allows for complex trailing-edge with the primitive aspects of operating the equipment. shapes which could contribute aerodynamic, structural, The performance and reliability of small models, in and control advantages. addition to their considerably lower cost and simplified operation, create an environment where high-risk, high- payoff experiments can be conducted.

One of the concepts under investigation is active wing shaping. Somewhat reminiscent of the 1903 Wright Brother’s wing warping scheme, active wing shaping strategies employ the wing as an entire control surface. Through various methods, the wing is shaped, deflected, or deformed to respond to changing conditions or impart changes on the aircraft’s flight path1. The shaping produces much more complex modes of actuation (Figure 1) than can be achieved with conventional control surfaces.

Figure 1: NASA vision for a “morphing” aircraft *Undergraduate Student, [email protected], Student Member AIAA Mechanical and Aerospace Engineering The need for such an aircraft is clear. Most types of airplane, both civil and military, operate in a wide Copyright © 2003 Mujahid Abdulrahim. Published by variety of conditions. Some of these have conflicting the American Institute of Aeronautics and Astronautics, requirements on aircraft design, where an efficient Inc. with permission.

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configuration in one instance may perform poorly in of the flight testbed are not of interest. Rather, it is the others. The rigid, non-deformable structures of these change in performance afforded by unique actuation of airplanes preclude any adaptation to changing the surfaces that warrants study. As such, the choice of conditions. Alternatively, an aircraft equipped with airplane is irrelevant, provided that a minimum level of active wing shaping would continuously respond to a performance is available to reflect the effectiveness of dynamic environment by deforming or deflecting parts the surfaces. In this regard, the MiG-27 was ideal, of the . having basic aerobatic capability. Furthermore, the inherent stability and simple operation of the aircraft The material presented in this paper provides an initial made it well suited for use as a controls testbed. look at the issues related to developing and testing an airplane that might ultimately lead to a morphing SPECIFICATIONS vehicle. It is in no way a comprehensive study of the subject. Rather, it is merely the beginning of a series of The aircraft used in this research is largely similar to design and testing. hobby remote control aircraft. The building techniques and hardware used throughout the airframe are derived AIRCRAFT SYSTEM exclusively from R/C modeling. The airframe is composed entirely of injection-molded Styrofoam. In developing an airborne controls testbed, the This facilitates assembly and allows the structure to be requirement for simplicity and cost-effectiveness easily modified to incorporate actuators and outweighed any performance objective. The aircraft instrumentation. used must be easily modified to incorporate actuators and instrumentation. It also must be large enough to sustain the weight of such payload without affecting flight performance.

Figure 2: FQM-117B “MiG-27” aircraft in flight

The airplane used for this research, shown in Figure 2, Figure 3: Top view of the MiG-27, note 16 servos is a military designation FQM-117B radio controlled miniature aerial target. Donated by Ft. Eustis Army Base, this military target drone is shaped entirely out of white Styrofoam, facilitating construction and modifications to the structure. The model is similar in shape to a Russian MiG-27 “Flogger”, referred to as MiG-27 for short (Figures 3 and 4). Although the original purpose of the aircraft was to provide target practice for Stinger missiles, simple modifications converted it to a suitable research platform. The modifications included addition of , rudder control, and surface finish.

In the study of the effect of multiple actuators on aircraft control, the specific performance characteristics Figure 4: Front and side views of the MiG-27

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Table 1: MiG-27 Specifications Unlike the 3-DAS, which interfaces exclusively with Dimensions one sensor, the µDAS can be interfaced with a variety Length 6 ft of sensors that output analog voltage. Most Wingspan 5.5 ft importantly, it is used to measure actuator position, Wing Area 800 in2 which is directly related to both pilot input and control Wing Loading 14.4 – 23 oz/in2 surface deflection. This process takes advantage of the Controls position feedback potentiometer inside the control Aileron -45º to +45º actuators. The voltage of the center pot lead, which is Elevator -20º to +30º directly proportional to actuator position, is read for Rudder -40º to +40º each of the primary control surface servos. For the wing servos, the voltage is measured for only one INSTRUMENTATION servo. The position of the remaining surfaces can then be determined with knowledge of the control algorithm. The instrumentation system measures the aircraft states required for flight dynamics characterization2. Included An alternate sensor system was used to generate the are control, attitude, and rate sensors that describe flight data presented in this paper. The 3DM-G control inputs and the resulting aircraft response. In orientation sensor was replaced with 3-axis rate and addition to the sensors, devices are used to interpret acceleration sensors interfaced directly to the µDAS. signals and record sensor outputs for post-flight The output of these sensors is satisfactory for flight analysis. Table 2 below summarizes the measurements testing. The additional DAS required to use the 3DM- in terms of aircraft states. G was not completed in time for publication.

Table 2: Measured aircraft states ACTUATORS Linear acceleration - ax, ay, az Angular rates (roll, pitch, yaw) - p, q, r The basis of this research is to develop a flying testbed for control of deformable surfaces3. Part of this Euler flight angles - FQY,, development involved designing and selecting Control surface deflection - d a ,d e ,d r hardware and software needed for such control aspirations. Elevator and rudder surfaces on the test The aircraft instrumentation system consists of two aircraft are unmodified. However, the standard ailerons primary components: orientation sensing and data are replaced by an array of surfaces, each independently acquisition. A MicroStrain 3DM-G sensor is used for controlled. The system is used to investigate the effect measurement of attitude and orientation. It is equipped of a control array on the controllability of an aircraft. with 3 gyros, 3 accelerometers, and 3 magnetometers. The choice of using 16 actuators has no basis aside The output of these 9 sensors are internally correlated from being a convenient number to begin such an and low-pass filtered to improve signal to noise ratio. investigation.

The 3DM-G sensor outputs data at 100Hz using a serial In the standard configuration, the MiG-27 uses three interface. A data acquisition system based on an Atmel servos for guidance and control. These allow the pilot microcontroller unit has been developed specifically for to command elevator, aileron, and engine . For recording output from this sensor. The data is recorded aileron control, a single servo differentially actuates in non-volatile memory and is downloaded upon two control surfaces, one on each side of the wing. landing. Additional servos installed for this research actuate rudder and nose-gear. In addition to the 3DM-G orientation sensor and associated DAS, the MiG-27 is also equipped with a In transitioning from a single roll actuator to 16, the micro data acquisition system (µDAS). Developed by selection of servo becomes increasingly important. The NASA Langley Research Center, the µDAS is much servo used for conventional actuation is prohibitively like traditional data loggers in that it converts analog large for a wing-mounted array. However, technology signals from sensors to digital data. Sampling rates for small actuators has improved considerably in recent ranging from 50 to 500Hz are available on the 30 input years. Miniature versions of standard-sized servos have channels of the µDAS. Each channel uses a 12-bit A-D higher strength to mass ratios, making them suitable for converter. Since the voltage input range is set, sensor use in large numbers. One such servo is the Hitec HS- outputs are amplified accordingly to produce suitable 55, compared below (Table 3) to the standard Futaba S- resolution. 148 on the conventional MiG-27.

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Table 3: Servo specifications edge can be actively deformed to produce favorable Futaba S-148 Standard Servo aircraft responses. One of these deformations may be a Number of servos used 1 mode that eliminates undesirable pitch and yaw Current draw 196mA coupling associated with conventional aileron actuation. Weight 1.6oz Others modes may be more complex, acting as dual- Torque 44.3oz-in mode flight controls that can command roll or yaw Deflection Speed 0.23s @ 60º actuation without a rudder or perhaps without a vertical Torque/actuating area 0.527oz/in .

Hitec HS-55 Micro Servo 1 Number of servos used 16 Current draw (total) 150(2400)mA Weight (total) 0.28(4.48)oz 2 Torque 15.3oz-in Deflection Speed 0.17s @ 60º Torque/actuating area 2.9oz/in

Although the aileron actuators are typically located inside the , the large number of servos presents logistical challenges to such a method. To simplify the 3 4 control linkages, each individual servo actuator is imbedded in the foam wing surface directly before the corresponding surface. Given the comparatively small surface area of each aileron segment, only small servos are needed for actuation. The miniscule size and weight of the servos allows them to be imbedded fully in the wing surface, with negligible effect on the 5 longitudinal roll moment of inertia.

Figure 5: Schematic of control hardware In order to recess the servo into the upper wing surface, 1- Transmitter, 2- Airborne receiver, a Dremel tool was used to mill a cavity in the exact 3-Microcontroller, 4- Servo controller, dimensions of the casing. Each servo was then 5- Control surface actuator mounted in the cavity with double-sided tape and epoxy. Standard R/C control linkages and horns are One of the challenges of increased authority control lies used to connect the servo actuator with the control in the development of the control algorithm and surface. associated hardware. Existing hobby-type equipment does not allow for the control or configurability Each servo wire lead was extended according to its required to operate 16 wing servos and elevator, rudder, respective position along the span. This minimizes and throttle. Thus, both the hardware and software excess wire clutter inside the fuselage, where all 16 needed to control the servos were developed servo leads are plugged into the left and right side servo specifically for this research. controllers. The control hardware is based on a 16Mhz Atmel CONTROL SYSTEM ATmega16 microcontroller. This interfaces with other devices through input/output ports. Thus, it acts as both As used by the Army, the MiG-27 roll control system a sensor and a controller. The ability to use software consists of two half-span ailerons hinged along the code to modify the control algorithm is necessary to of the wing. A single servo is normally implement the novel modes of actuation. Because the used to actuate both surfaces. For the purposes of this servos are not controlled directly by the pilot, the MCU research, these ailerons are replaced by 16 reads and interprets pilot commands then controls each independently actuated control surfaces. Each of these servo appropriately. The algorithms and mode shapes are equally sized and spaced evenly along the trailing can then be changed between flights to improve aircraft edge. A servo-actuator imbedded in the wing surface is response. connected to each control surface. While the total Open-loop control is used to actuate the wing servos. control surface area remains the same, the surfaces Since no sensor feedback is used, the motion is based allow for increased configurability. i.e. the trailing strictly on pilot input. The pilot, standing on the ground

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observing the airplane, commands the input to the servo(s). The microcontroller generates commands in various controls using a hand-held radio controlled the form of numerical positions. These commands are (R/C) transmitter. The transmitted signals are interpreted by Scott Edwards Mini servo controllers interpreted by the airborne receiver, which is mounted which generate the pulse width modulation signals inside the aircraft. The receiver in turn sends signals to needed to communicate with the servos. the individual servos which move to the commanded position. The large number of servos precludes the use OPEN-LOOP FLIGHT TESTING of the R/C electronics for direct servo control. The 22 servos on-board the aircraft exceed the capabilities of Open-loop flight testing is performed with the MiG-27 typical hobby transmitters and receivers. aircraft using procedures developed by NASA Dryden Flight Research Center (Figure 6 and 7). These Control of the wing servos and other is procedures outline recommended instrumentation and achieved by means of a microcontroller unit (MCU). maneuvers needed to characterize airplane dynamics. The MCU interprets pilot command signals output from Specifically of interest in this case is the control pulse the receiver. By measuring the pulse width of this maneuver, which is used to determine control signal, the MCU digitizes commands and is able to effectiveness and stability and control derivatives. computationally adjust gain and offset for each wing actuator. In this manner, a single command can be used A control pulse is a maneuver where rapid control for control of the 16 trailing edge segments to input is used to upset the aircraft from a stable trim predetermined mode shapes. The complete control condition into oscillatory motion2. This input is system is described in greater detail below. typically in the form of a doublet, where the control surface is quickly actuated in both directions. The The two command channels input to the response of the aircraft to such inputs is used to identify microcontroller enable two distinct flight modes. One frequency and damping of the oscillatory motion, in mode is an emulation of conventional aileron actuation, addition to the parameter identification. The latter case where all servos actuate simultaneously. This mode is requires data analysis programs that determine used for takeoff, standard maneuvering, and all other coefficients, or parameters, of the known aircraft non-testing aspects of the flight. The second control equations of motion. One such program is pEst mode actuates a single servo, selectable by a rotating (parameter estimation), also developed by NASA knob on the R/C transmitter. This mode is used for DFRC. This will be used extensively in the analysis of examination of aircraft response to a control doublet MiG-27 flight data. performed on the single servo. In this way, the surfaces can be characterized individually using standard flight test maneuvers and non-linear parameter estimation software.

The microcontroller performs three functions: Reading the command signal, computing the control algorithm, and commanding the 16 servo positions.

The signal from the R/C receiver to the microcontroller is in the form of pulse width modulation (PMW). The command is sent by varying the duty cycle, changing the high pulse from 1ms to 2ms. Normally, the R/C servo position is directly proportional to the pulse width. In this instance, where the command is input into an MCU, the signal is timed and converter into an integer from 0 to 255. A value of 127 indicates center Figure 6: Remotely-piloted flight test of the MiG or neutral command. Values of 0 and 255 are minimum and maximum command position respectively. For Control pulses are performed on all three controls direct servo control, this integer becomes the (aileron, elevator, and rudder). The flight test are commanded position. The actuation range of each performed with conventional one-piece ailerons and servo +/- 45 degrees, with a resolution of 90/255 = 0.35 with 16-segment ailerons. The former case is used to degrees. An additional control knob allows selection of determine the unmodified behavior of the aircraft. The individual or global servo control. The code performs a longitudinal and lateral-directional stability and control check of this position before finally commanding the derivatives obtained from this testing and subsequent

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analysis establish a baseline for comparison with the these are within the linear dynamics regime. modified aircraft. For comparative purposes, the 16- Maneuvers larger than this require increasingly segment ailerons are configured to deflect uniformly complex algorithms to characterize. during these flight tests (Figure 8-a). MODEL IDENTIFICATION

The data collected during flight test maneuvers is used to identify the flight dynamics of the MiG. Although the complete aerodynamic model cannot be developed without extensive testing and data analysis, relevant coefficients of the A, B, C, and D stability and control matricies can be determined4. In particular, the roll dynamics terms Clp and Clda are easily computed using the roll command and roll rate signals1. A 3rd order Butterworth low-pass filter is applied in Matlab to remove noise from the data and aid in convergence of subsequent analyses. An autoregressive, moving- average (ARMA) process is then used to represent the dynamics. Using standard regression analysis, the coefficients to the ARMA process are determined. The Figure 7: MiG flying a roll control doublet maneuver result is an estimation of Clp and Clda, roll convergence and aileron effectiveness respectively. The flight procedure is identical for all control pulses. Prior to take-off, the control system is checked to Identification of these parameters is a useful measure of ensure proper actuation direction and magnitude. Once roll performance for different aileron configurations. the operation of the control and instrumentation Thus, this procedure is repeated for each shape. An systems is verified, the MiG-27 is piloted remotely by a additional measure of maximum roll rate is also ground operator through the entire flight. A grass provided for comparison. This value is determined by runway is used to accelerate the aircraft along the absolute measure of flight data. The various aileron ground until suitable flying speed is reached. At this configurations used in such testing are shown below in point, it is rotated and begins a climb to a maneuvering Figure 8. altitude of 400 feet. During this portion of flight, control input is a combination of aileron, elevator, and rudder command. Although the instrumentation records the entire flight, the only portion of interest is the time immediately surrounding the pulse maneuver.

Once at sufficient altitude, the airplane is trimmed for straight and level flight. This is checked by releasing the control stick and ensuring the airplane does not deviate from the initial flight path. Several passes may be needed to ensure adequate trim.

8-a) Uniform aileron deflection using 16 servos With the airplane at a constant airspeed and altitude (identical to conventional ailerons) (non accelerated flight), a roll is commanded in one direction for a small time, then commanded in the opposite direction, and finally back to neutral position. Once completed, the pilot immediately releases the controls and allows the airplane to oscillate and recover from the upset. Active control of the airplane is resumed several seconds after the maneuver, when the oscillations have sufficiently decreased. Pitch and yaw doublets are performed in a similar manner. Control pulse maneuvers generally involve small aircraft movements. With roll doublets, for instance, 8-b) Outboard aileron segment deflection the maximum bank angle achieved during the maneuver does not exceed 40º. Relatively small motions such as

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8-c) Progressive aileron deflection

8-d) Neutral aileron position Figure 10: Control pulse to 16-servo wing Figure 8: Various deflection modes implemented in (uniform deflection) flight tests The flight data presented is in the form of aileron Figures 9 and 10 show examples of data from flight deflection and roll rate versus time. Although figures tests using conventional and segmented ailerons. Both may be scaled differently, each grid line corresponds to graphs depict a very good correlation between 10 degrees deflection change and 20 degree per second command input and aircraft response. This is roll rate change. somewhat of a reassurance that the airplane dynamics are in fact within the range of the linear aerodynamic Table 4 below shows the results of model identification model. routines described earlier. The values represent the average of three to four maneuvers for each aileron Figure 10 depicts an unexpected result of the segmented configuration. The A-matrix term, Clp, represents the wing. There was initially some concern that the roll stability of the airplane. A more negative value complexity of the control system would result in time indicates an increased resistance to rolling. The B- delays in actuation. Based on pilot feedback and flight matrix term, Clda, represents control effectiveness or data, this is not a factor. Rather, the controllability of responsiveness. A more negative value indicates the segmented aileron was better than that of the increased roll response to aileron deflection. conventional aileron. For a given command input, the corresponding roll rate achieved by the segmented wing The last two table rows, Outboard and Progressive, are is greater than the conventional aileron roll rate. This is mode shapes that are described in the following section. likely attributed to the increased torque-surface area Note that all the A-matrix values obtained from flight ratio of the actuator-control surface system. tests of 16-servo wing are very similar. This is an

expected result, as the Clp coefficient is independent of control. The discrepancy between the conventional aileron Clp is likely due to the increased roll moment of inertia that comes from installing servos in the wing rather than inside the fuselage.

Table 4: Identified roll dynamics

Aileron A (Clp) B (Clda) C D Conventional -20.43 -17.03 1 -0.423 Uniform* -11.79 -22.60 1 -0.510 Outboard* -15.33 -11.46 1 -0.270 Progressive* -13.93 -11.82 1 -0.273

* denotes 16-servo deflection – see figure 8

Figure 9: Control pulse to conventional aileron 7 American Institute of Aeronautics and Astronautics

NOVEL CONTROL METHODS CONTROL OF LIFT DISTRIBUTION

The model developed from the data analysis of open- One application envisioned for this aircraft is active loop flight data is used to optimize control deflection of control of wing lift distribution. The aileron segments the 16 wing surfaces. Control doublets performed may be deflected optimally for a certain objective. In using individual segment deflection are used to fully level cruise, for instance, the segments could be characterize the control effectiveness of each surface. deflected to modify section lift properties such that an In other words, for a given flight condition, a certain elliptical lift distribution is achieved. This distribution control deflection produces a known aircraft response. is classically regarded as the most efficient, having the This control effectiveness is described by a component highest theoretical lift to drag ratio. In this manner, of the B-matrix of the aircraft stability and control aerodynamic properties are a function of both airframe equations. For the case of the 16-servo MiG, the B- and control configuration. For dynamic maneuvers, matrix is assumed to be the summation of the individual control deflection may be optimized to improve contributions of the surfaces. response or decrease unwanted coupling and/or side effects. Decreasing drag and eliminating adverse yaw Knowing these contributions affords the opportunity to associated with aileron actuation are two such command a very specific response. This could be as examples. Progressive deflection of the aileron simple as accurately following a flight path or segments (Figures 8-c, 12) is used to minimize drag performing a complex maneuver in the non-linear incurred during roll maneuvers. This mode commands aerodynamic range. Other applications are minimally maximum deflection to the outboard segment, and coupled maneuvers, drag-efficient control deflection, or progressively smaller deflection to inboard segments. increased control redundancy.

PARTIAL AILERON DEFLECTION

A flight test of select actuator deflection was performed to investigate the effectiveness of individual surfaces. While independent actuation of each of the 16 surfaces is desirable, a simplified flight test was flown where only the four outboard aileron segments were actuated during the roll control pulse (figure 8-b). Figure 11 below shows the flight data obtained during the maneuver. Contrasting this data to the uniform deflection, it is evident that this actuation mode is considerably less effective. A cursory examination of both data reveals that 8 aileron segments deflected twice as much as 16 segments produces similar roll Figure 12: Progressive aileron deflection rate. This is in agreement with tabulated values for Clda.

Figure 11: Control pulse to outboard 4 segments Figure 13: Control pulse to progressive ailerons

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Data from a progressive aileron control pulse (figure The computations indicated that uniform deflection of 12) shows a similarity to uniform deflection of the the aileron segments resulted in 33% greater drag than segments. An average segment deflection of 20º, for progressively deflected ailerons for comparable roll instance, results in 40º/sec roll rate. Whereas a uniform moments. It is shown that the inboard segments 12º deflection of all segments results in 25º/sec roll contribute relatively little to rolling moment, yet cause rate. considerable amounts of drag. The bulk of the roll In order to compare the relative amounts of drag moment is generated by the outboard control segments. produced by various aileron deflections, simplified Thus, a theoretical maximum efficiency roll actuation is aerodynamic formulae are used in conjunction to wind one where outboard surfaces are deflected to large tunnel data5,6. Section lift and drag forces are angles, while inboard surfaces are deflected to smaller calculated for each wing segment. Using data from an angles. This weighting scheme applies increased airfoil closely resembling that of the MiG’s (figure 14), control authority to most-effective surfaces. the following equations are used to calculate drag and rolling moment. YAW CONTROL 8 (Drag = åCd,s*(rho*V^2)/2*,Areas) While the previous section described a method to s=1 reduce drag, another method is being researched to exploit the drag by-product of aileron deflection. The 8 controller is configured to have independent control of (Roll = åCl,s*(rho*V^2)/2*Area,*,sys) the 16 aileron segments. A yaw command causes s=1 adjacent wing segments to deflect in opposite directions on a single wing side (Figure 15). The surfaces are Where, Cl,s = section lift coefficient trimmed such that the roll and pitch moments induced Cd,s = section drag coefficient by this actuation are negligible. However, as is evident rho = sea level density by the airfoil data from Figure 13, the drag from this V = cruise airspeed, 77 ft/s deflection is considerable. Thus, with greater drag on Area, s = section wing area one wing, the airplane will incur a yawing moment. y, s = spanwise distance from centerline Initially, this is being investigated as a primary control effector. The wing-induced yaw may be used in later research as a measure of active stability, replacing the entirely.

Time restrictions precluded flight testing of the wing- yaw control mode. However, preliminary calculations of expected yaw moment are similar to the moment generated by the vertical stabilizer/rudder. This mode shape will be investigated in future flight tests.

Figure 15: Segments on left wing deflected to induce yaw moment Figure 14: Wind tunnel measurements of lift and drag

for a Clark Y (courtesy of NACA Report # 554)

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CONCLUSION [3] M. Abdulrahim and R. Lind, “Investigating Segmented Trailing-Edge Surfaces for Full Authority An airborne testbed has been developed to evaluate the Control of a UAV,” submitted to AIAA Guidance, flight dynamics highly-segmented control surface on a Navigation, and Control Conference, August 2003. UAV’s wing. Certain technological developments, namely in micro instrumentation and control hardware, [4] M. R. Waszak, L.N. Jenkins and P. Ifju, “Stability have enabled such research. These devices allow the and Control Properties of an Aeroelastic Fixed Wing aircraft control system to be reconfigured for different Micro Air Vehicle,” AIAA Atmospheric Flight actuation modes. Implementation of this control system Mechanics Conference, August 2001 on the test aircraft is achieved without adversely affecting flying characteristics. Novel control methods [5] F.E. Weick and J.A. Shortal, “The Effect of of the highly-segmented wings are used to investigate Multiple Fixed Slots and a Trailing-Edge Flap on the non-conventional modes of actuation. Flight tests show Lift and Drag of a Clark Y Airfoil,” NACA Report No. improved performance over the unmodified aircraft. 427, 1934 Additional testing aimed at better understanding the segmented wing characteristics are ongoing. [6] C.J. Wenzinger, “Wind-tunnel investigation of ordinary and split flaps on airfoils of different profile,” NACA Report No. 554, 1937

ACKNOWLEDGMENTS

The author would like to acknowledge the contributions of numerous individuals to this research. Rick Lind of the University of Florida continually advised the author on all aspects of the research, from flight testing procedure to controller development. Mike French of the Ft. Eustis Army Base generously provided the healthy-sized fleet of MiG-27 models for use in this research. Mark Motter of the NASA Langley Research Center, who is working on similar research, collaborated on many issues. He shared his experience with operating similar aircraft and offered advice to improve the quality of the research. The instrumentation and sensors used to measure and record flight data were developed and provided by Marty Waszak also of NASA Langley Research Center. Jason Grzywna, Joe Pippin, and Erik Sandem provided hardware and advice for developing the MiG control system. Peter Ifju provided advice, insight, and funds throughout the testing process. Finally, Rich Maine of NASA Dryden Flight Research Center provided the parameter estimation code and assisted in our analysis of flight data.

REFERENCES

[1] H.M. Garcia, M. Abdulrahim and R. Lind, “Roll Control for a Micro Air Vehicle using Active Wing Morphing,” submitted to AIAA Guidance, Navigation, and Control Conference, August 2003.

[2] R.G. Hoey, “Control Pulse,” NASA Education Notes

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