flight dynamics and control of an aircraft with segmented control

AIAA-RSC2-2003-U-010

Flight Dynamics and Control of an Aircraft 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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American Institute of Aeronautics and Astronautics

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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 ailerons for roll control and as a supplement to rudder, elevator, and flap 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 aileron 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

Small, unmanned air vehicles are increasingly used as a tool for flight research. Equipment and instrumentation that once was prohibitively large and expensive is now available for these miniature aircraft. While the use of UAVs for research continues gaining acceptance, the capabilities of the individual research teams continue to expand. No longer are flight researchers concerned with the primitive aspects of operating the equipment. The performance and reliability of small models, in 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. *Undergraduate Student, [email protected], Student Member AIAA Mechanical and Aerospace Engineering Copyright © 2003 Mujahid Abdulrahim. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission.

A preliminary approach to designing a morphing vehicle is increasing the number control surfaces. This research focuses on the development and characterization of such an aircraft. The vehicle in question is equipped with 16 independent wing control surfaces in place of the conventional ailerons. Although actuated in a similar fashion, the large number of surfaces allows for complex trailing-edge shapes which could contribute aerodynamic, structural, and control advantages.

Figure 1: NASA vision for a “morphing” aircraft

The need for such an aircraft is clear. Most types of airplane, both civil and military, operate in a wide variety of conditions. Some of these have conflicting requirements on aircraft design, where an efficient


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configuration in one instance may perform poorly in others. The rigid, non-deformable structures of these airplanes preclude any adaptation to changing conditions. Alternatively, an aircraft equipped with active wing shaping would continuously respond to a dynamic environment by deforming or deflecting parts of the airframe. The material presented in this paper provides an initial look at the issues related to developing and testing an airplane that might ultimately lead to a morphing vehicle. It is in no way a comprehensive study of the subject. Rather, it is merely the beginning of a series of design and testing.

AIRCRAFT SYSTEM In developing an airborne controls testbed, the requirement for simplicity and cost-effectiveness outweighed any performance objective. The aircraft 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.

The airplane used for this research, shown in Figure 2, 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 landing gear, rudder control, and surface finish. In the study of the effect of multiple actuators on aircraft control, the specific performance characteristics

of the flight testbed are not of interest. Rather, it is the change in performance afforded by unique actuation of the surfaces that warrants study. As such, the choice of airplane is irrelevant, provided that a minimum level of performance is available to reflect the effectiveness of the surfaces. In this regard, the MiG-27 was ideal, having basic aerobatic capability. Furthermore, the inherent stability and simple operation of the aircraft made it well suited for use as a controls testbed. SPECIFICATIONS The aircraft used in this research is largely similar to hobby remote control aircraft. The building techniques and hardware used throughout the airframe are derived exclusively from R/C modeling. The airframe is composed entirely of injection-molded Styrofoam. This facilitates assembly and allows the structure to be easily modified to incorporate actuators and instrumentation.

Figure 4: Front and side views of the MiG-27

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

Figure 3: Top view of the MiG-27, note 16 servos


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Table 1: MiG-27 Specifications

Dimensions Length 6 ft Wingspan 5.5 ft Wing Area 800 in2

Wing Loading 14.4 – 23 oz/in2

Controls Aileron -45º to +45º Elevator -20º to +30º Rudder -40º to +40º

INSTRUMENTATION The instrumentation system measures the aircraft states required for flight dynamics characterization2. Included are control, attitude, and rate sensors that describe control inputs and the resulting aircraft response. In addition to the sensors, devices are used to interpret signals and record sensor outputs for post-flight analysis. Table 2 below summarizes the measurements in terms of aircraft states.

Table 2: Measured aircraft states Linear acceleration - ax, ay, az Angular rates (roll, pitch, yaw) - p, q, r Euler flight angles - , ,Φ Θ Ψ

Control surface deflection - aδ , eδ , rδ The aircraft instrumentation system consists of two primary components: orientation sensing and data acquisition. A MicroStrain 3DM-G sensor is used for measurement of attitude and orientation. It is equipped with 3 gyros, 3 accelerometers, and 3 magnetometers. The output of these 9 sensors are internally correlated and low-pass filtered to improve signal to noise ratio. The 3DM-G sensor outputs data at 100Hz using a serial interface. A data acquisition system based on an Atmel microcontroller unit has been developed specifically for recording output from this sensor. The data is recorded in non-volatile memory and is downloaded upon landing. In addition to the 3DM-G orientation sensor and associated DAS, the MiG-27 is also equipped with a micro data acquisition system (µDAS). Developed by NASA Langley Research Center, the µDAS is much like traditional data loggers in that it converts analog signals from sensors to digital data. Sampling rates ranging from 50 to 500Hz are available on the 30 input channels of the µDAS. Each channel uses a 12-bit A-D converter. Since the voltage input range is set, sensor outputs are amplified accordingly to produce suitable resolution.

Unlike the 3-DAS, which interfaces exclusively with one sensor, the µDAS can be interfaced with a variety of sensors that output analog voltage. Most importantly, it is used to measure actuator position, which is directly related to both pilot input and control surface deflection. This process takes advantage of the position feedback potentiometer inside the control actuators. The voltage of the center pot lead, which is directly proportional to actuator position, is read for each of the primary control surface servos. For the wing servos, the voltage is measured for only one servo. The position of the remaining surfaces can then be determined with knowledge of the control algorithm. An alternate sensor system was used to generate the flight data presented in this paper. The 3DM-G orientation sensor was replaced with 3-axis rate and acceleration sensors interfaced directly to the µDAS. The output of these sensors is satisfactory for flight testing. The additional DAS required to use the 3DM-G was not completed in time for publication. ACTUATORS The basis of this research is to develop a flying testbed for control of deformable surfaces3. Part of this development involved designing and selecting hardware and software needed for such control aspirations. Elevator and rudder surfaces on the test aircraft are unmodified. However, the standard ailerons are replaced by an array of surfaces, each independently controlled. The system is used to investigate the effect of a control array on the controllability of an aircraft. The choice of using 16 actuators has no basis aside from being a convenient number to begin such an investigation. In the standard configuration, the MiG-27 uses three servos for guidance and control. These allow the pilot to command elevator, aileron, and engine throttle. For aileron control, a single servo differentially actuates two control surfaces, one on each side of the wing. Additional servos installed for this research actuate rudder and nose-gear. In transitioning from a single roll actuator to 16, the selection of servo becomes increasingly important. The servo used for conventional actuation is prohibitively large for a wing-mounted array. However, technology for small actuators has improved considerably in recent years. Miniature versions of standard-sized servos have higher strength to mass ratios, making them suitable for use in large numbers. One such servo is the Hitec HS-55, compared below (Table 3) to the standard Futaba S-148 on the conventional MiG-27.


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Table 3: Servo specifications Futaba S-148 Standard Servo Number of servos used 1 Current draw 196mA Weight 1.6oz Torque 44.3oz-in Deflection Speed 0.23s @ 60º Torque/actuating area 0.527oz/in Hitec HS-55 Micro Servo Number of servos used 16 Current draw (total) 150(2400)mA Weight (total) 0.28(4.48)oz Torque 15.3oz-in Deflection Speed 0.17s @ 60º Torque/actuating area 2.9oz/in

Although the aileron actuators are typically located inside the fuselage, the large number of servos presents logistical challenges to such a method. To simplify the 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 longitudinal roll moment of inertia. In order to recess the servo into the upper wing surface, a Dremel tool was used to mill a cavity in the exact dimensions of the casing. Each servo was then mounted in the cavity with double-sided tape and epoxy. Standard R/C control linkages and horns are used to connect the servo actuator with the control surface. Each servo wire lead was extended according to its respective position along the span. This minimizes excess wire clutter inside the fuselage, where all 16 servo leads are plugged into the left and right side servo controllers. CONTROL SYSTEM As used by the Army, the MiG-27 roll control system consists of two half-span ailerons hinged along the trailing edge of the wing. A single servo is normally used to actuate both surfaces. For the purposes of this research, these ailerons are replaced by 16 independently actuated control surfaces. Each of these are equally sized and spaced evenly along the trailing edge. A servo-actuator imbedded in the wing surface is connected to each control surface. While the total control surface area remains the same, the surfaces allow for increased configurability. i.e. the trailing

edge can be actively deformed to produce favorable aircraft responses. One of these deformations may be a mode that eliminates undesirable pitch and yaw coupling associated with conventional aileron actuation. Others modes may be more complex, acting as dual-mode flight controls that can command roll or yaw actuation without a rudder or perhaps without a vertical stabilizer.

One of the challenges of increased authority control lies in the development of the control algorithm and associated hardware. Existing hobby-type equipment does not allow for the control or configurability required to operate 16 wing servos and elevator, rudder, and throttle. Thus, both the hardware and software needed to control the servos were developed specifically for this research. The control hardware is based on a 16Mhz Atmel ATmega16 microcontroller. This interfaces with other devices through input/output ports. Thus, it acts as both a sensor and a controller. The ability to use software code to modify the control algorithm is necessary to implement the novel modes of actuation. Because the servos are not controlled directly by the pilot, the MCU reads and interprets pilot commands then controls each servo appropriately. The algorithms and mode shapes can then be changed between flights to improve aircraft response. Open-loop control is used to actuate the wing servos. Since no sensor feedback is used, the motion is based strictly on pilot input. The pilot, standing on the ground

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2

3 4

5 Figure 5: Schematic of control hardware

1- Transmitter, 2- Airborne receiver, 3-Microcontroller, 4- Servo controller, 5- Control surface actuator


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observing the airplane, commands the input to the various controls using a hand-held radio controlled (R/C) transmitter. The transmitted signals are interpreted by the airborne receiver, which is mounted inside the aircraft. The receiver in turn sends signals to the individual servos which move to the commanded position. The large number of servos precludes the use of the R/C electronics for direct servo control. The 22 servos on-board the aircraft exceed the capabilities of typical hobby transmitters and receivers. Control of the wing servos and other aircraft systems is achieved by means of a microcontroller unit (MCU). The MCU interprets pilot command signals output from the receiver. By measuring the pulse width of this signal, the MCU digitizes commands and is able to computationally adjust gain and offset for each wing actuator. In this manner, a single command can be used for control of the 16 trailing edge segments to predetermined mode shapes. The complete control system is described in greater detail below. The two command channels input to the microcontroller enable two distinct flight modes. One mode is an emulation of conventional aileron actuation, where all servos actuate simultaneously. This mode is used for takeoff, standard maneuvering, and all other non-testing aspects of the flight. The second control mode actuates a single servo, selectable by a rotating knob on the R/C transmitter. This mode is used for examination of aircraft response to a control doublet 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 or neutral command. Values of 0 and 255 are minimum and maximum command position respectively. For direct servo control, this integer becomes the commanded position. The actuation range of each servo +/- 45 degrees, with a resolution of 90/255 = 0.35 degrees. An additional control knob allows selection of individual or global servo control. The code performs a check of this position before finally commanding the

servo(s). The microcontroller generates commands in the form of numerical positions. These commands are interpreted by Scott Edwards Mini servo controllers which generate the pulse width modulation signals needed to communicate with the servos. OPEN-LOOP FLIGHT TESTING Open-loop flight testing is performed with the MiG-27 aircraft using procedures developed by NASA Dryden Flight Research Center (Figure 6 and 7). These procedures outline recommended instrumentation and maneuvers needed to characterize airplane dynamics. Specifically of interest in this case is the control pulse maneuver, which is used to determine control effectiveness and stability and control derivatives. A control pulse is a maneuver where rapid control input is used to upset the aircraft from a stable trim condition into oscillatory motion2. This input is typically in the form of a doublet, where the control surface is quickly actuated in both directions. The response of the aircraft to such inputs is used to identify frequency and damping of the oscillatory motion, in addition to the parameter identification. The latter case requires data analysis programs that determine coefficients, or parameters, of the known aircraft equations of motion. One such program is pEst (parameter estimation), also developed by NASA DFRC. This will be used extensively in the analysis of MiG-27 flight data.

Figure 6: Remotely-piloted flight test of the MiG

Control pulses are performed on all three controls (aileron, elevator, and rudder). The flight test are performed with conventional one-piece ailerons and with 16-segment ailerons. The former case is used to determine the unmodified behavior of the aircraft. The longitudinal and lateral-directional stability and control derivatives obtained from this testing and subsequent


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analysis establish a baseline for comparison with the modified aircraft. For comparative purposes, the 16-segment ailerons are configured to deflect uniformly during these flight tests (Figure 8-a).

Figure 7: MiG flying a roll control doublet maneuver The flight procedure is identical for all control pulses. Prior to take-off, the control system is checked to ensure proper actuation direction and magnitude. Once the operation of the control and instrumentation systems is verified, the MiG-27 is piloted remotely by a ground operator through the entire flight. A grass runway is used to accelerate the aircraft along the ground until suitable flying speed is reached. At this point, it is rotated and begins a climb to a maneuvering 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. With the airplane at a constant airspeed and altitude (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, the maximum bank angle achieved during the maneuver does not exceed 40º. Relatively small motions such as

these are within the linear dynamics regime. Maneuvers larger than this require increasingly complex algorithms to characterize.

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 result is an estimation of Clp and Clda, roll convergence and aileron effectiveness respectively. Identification of these parameters is a useful measure of roll performance for different aileron configurations. Thus, this procedure is repeated for each shape. An additional measure of maximum roll rate is also provided for comparison. This value is determined by absolute measure of flight data. The various aileron configurations used in such testing are shown below in Figure 8.

8-a) Uniform aileron deflection using 16 servos

(identical to conventional ailerons)

8-b) Outboard aileron segment deflection


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8-c) Progressive aileron deflection

8-d) Neutral aileron position

Figure 8: Various deflection modes implemented in

flight tests Figures 9 and 10 show examples of data from flight tests using conventional and segmented ailerons. Both graphs depict a very good correlation between command input and aircraft response. This is somewhat of a reassurance that the airplane dynamics are in fact within the range of the linear aerodynamic model. Figure 10 depicts an unexpected result of the segmented wing. There was initially some concern that the complexity of the control system would result in time delays in actuation. Based on pilot feedback and flight data, this is not a factor. Rather, the controllability of the segmented aileron was better than that of the conventional aileron. For a given command input, the corresponding roll rate achieved by the segmented wing is greater than the conventional aileron roll rate. This is likely attributed to the increased torque-surface area ratio of the actuator-control surface system.

Figure 10: Control pulse to 16-servo wing (uniform deflection)

The flight data presented is in the form of aileron deflection and roll rate versus time. Although figures may be scaled differently, each grid line corresponds to 10 degrees deflection change and 20 degree per second roll rate change. Table 4 below shows the results of model identification routines described earlier. The values represent the average of three to four maneuvers for each aileron configuration. The A-matrix term, Clp, represents the roll stability of the airplane. A more negative value indicates an increased resistance to rolling. The B-matrix term, Clda, represents control effectiveness or responsiveness. A more negative value indicates increased roll response to aileron deflection. The last two table rows, Outboard and Progressive, are mode shapes that are described in the following section. Note that all the A-matrix values obtained from flight 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


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NOVEL CONTROL METHODS The model developed from the data analysis of open-loop flight data is used to optimize control deflection of the 16 wing surfaces. Control doublets performed using individual segment deflection are used to fully characterize the control effectiveness of each surface. In other words, for a given flight condition, a certain control deflection produces a known aircraft response. This control effectiveness is described by a component of the B-matrix of the aircraft stability and control equations. For the case of the 16-servo MiG, the B-matrix is assumed to be the summation of the individual contributions of the surfaces. Knowing these contributions affords the opportunity to command a very specific response. This could be as simple as accurately following a flight path or performing a complex maneuver in the non-linear aerodynamic range. Other applications are minimally coupled maneuvers, drag-efficient control deflection, or 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 rate. This is in agreement with tabulated values for Clda.

Figure 11: Control pulse to outboard 4 segments

CONTROL OF LIFT DISTRIBUTION One application envisioned for this aircraft is active control of wing lift distribution. The aileron segments may be deflected optimally for a certain objective. In level cruise, for instance, the segments could be deflected to modify section lift properties such that an elliptical lift distribution is achieved. This distribution is classically regarded as the most efficient, having the highest theoretical lift to drag ratio. In this manner, aerodynamic properties are a function of both airframe and control configuration. For dynamic maneuvers, control deflection may be optimized to improve response or decrease unwanted coupling and/or side effects. Decreasing drag and eliminating adverse yaw associated with aileron actuation are two such examples. Progressive deflection of the aileron segments (Figures 8-c, 12) is used to minimize drag incurred during roll maneuvers. This mode commands maximum deflection to the outboard segment, and progressively smaller deflection to inboard segments.

Figure 12: Progressive aileron deflection

Figure 13: Control pulse to progressive ailerons


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Data from a progressive aileron control pulse (figure 12) shows a similarity to uniform deflection of the segments. An average segment deflection of 20º, for instance, results in 40º/sec roll rate. Whereas a uniform 12º deflection of all segments results in 25º/sec roll rate. In order to compare the relative amounts of drag produced by various aileron deflections, simplified aerodynamic formulae are used in conjunction to wind tunnel data5,6. Section lift and drag forces are calculated for each wing segment. Using data from an airfoil closely resembling that of the MiG’s (figure 14), the following equations are used to calculate drag and rolling moment.

(Drag = 8

1

, *( * ^2) / 2* ,s

Cd s rho V Area s=

∑ )

(Roll = 8

1

, *( * ^ 2) / 2* , * ,s

Cl s rho V Area s y s=

∑ )

Where, Cl,s = section lift coefficient

Cd,s = section drag coefficient rho = sea level density V = cruise airspeed, 77 ft/s Area, s = section wing area y, s = spanwise distance from centerline

Figure 14: Wind tunnel measurements of lift and drag

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

The computations indicated that uniform deflection of the aileron segments resulted in 33% greater drag than progressively deflected ailerons for comparable roll moments. It is shown that the inboard segments contribute relatively little to rolling moment, yet cause considerable amounts of drag. The bulk of the roll moment is generated by the outboard control segments. Thus, a theoretical maximum efficiency roll actuation is one where outboard surfaces are deflected to large angles, while inboard surfaces are deflected to smaller angles. This weighting scheme applies increased control authority to most-effective surfaces. YAW CONTROL While the previous section described a method to reduce drag, another method is being researched to exploit the drag by-product of aileron deflection. The controller is configured to have independent control of the 16 aileron segments. A yaw command causes adjacent wing segments to deflect in opposite directions on a single wing side (Figure 15). The surfaces are trimmed such that the roll and pitch moments induced by this actuation are negligible. However, as is evident by the airfoil data from Figure 13, the drag from this deflection is considerable. Thus, with greater drag on one wing, the airplane will incur a yawing moment. 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 vertical stabilizer 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


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CONCLUSION

An airborne testbed has been developed to evaluate the flight dynamics highly-segmented control surface on a UAV’s wing. Certain technological developments, namely in micro instrumentation and control hardware, have enabled such research. These devices allow the aircraft control system to be reconfigured for different actuation modes. Implementation of this control system on the test aircraft is achieved without adversely affecting flying characteristics. Novel control methods of the highly-segmented wings are used to investigate non-conventional modes of actuation. Flight tests show improved performance over the unmodified aircraft. Additional testing aimed at better understanding the segmented wing characteristics are ongoing.

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

[3] M. Abdulrahim and R. Lind, “Investigating Segmented Trailing-Edge Surfaces for Full Authority Control of a UAV,” submitted to AIAA Guidance, Navigation, and Control Conference, August 2003. [4] M. R. Waszak, L.N. Jenkins and P. Ifju, “Stability and Control Properties of an Aeroelastic Fixed Wing Micro Air Vehicle,” AIAA Atmospheric Flight Mechanics Conference, August 2001 [5] F.E. Weick and J.A. Shortal, “The Effect of Multiple Fixed Slots and a Trailing-Edge Flap on the Lift and Drag of a Clark Y Airfoil,” NACA Report No. 427, 1934 [6] C.J. Wenzinger, “Wind-tunnel investigation of ordinary and split flaps on airfoils of different profile,” NACA Report No. 554, 1937

flight dynamics and control of an aircraft with segmented control

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