Development of Flight Control System and Troubleshooting on Flight Test of a Tilt-Rotor Unmanned Aerial Vehicle

Paper Int’l J. of Aeronautical & Space Sci. 17(1), 120–131 (2016) DOI: http://dx.doi.org/10.5139/IJASS.2016.17.1.120 Development of Flight Control System and Troubleshooting on Flight Test of a Tilt-Rotor Unmanned Aerial Vehicle

Youngshin Kang*, Bum-Jin Park**, Am Cho***, Chang-Sun Yoo**** and Sam-Ok Koo***** Korea Aerospace Research Institute, Daejeon 34133, Republic of Korea

Min-Jea Tahk****** Korea Advanced Institute of Science and Technology, Daejeon 34141, Republic of Korea

Abstract The full results of troubleshooting process related to the flight control system of a tilt-rotor type UAV in the flight tests are described. Flight tests were conducted in helicopter, conversion, and airplane modes. The vehicle was flown using automatic functions, which include speed-hold, altitude-hold, heading-hold, guidance modes, as well as automatic take-off and landing. Many unexpected problems occurred during the envelope expansion tests which were mostly under those automatic functions. The anomalies in helicopter mode include vortex ring state (VRS), long delay in the automatic take-off, and the initial overshoot in the automatic landing. In contrast, the anomalies in conversion mode are untrimmed AOS oscillation and the calibration errors of the air data sensors. The problems of low damping in rotor speed and roll rate responses are found in airplane mode. Once all of the known problems had been solved, the vehicle in airplane mode gradually reached the maximum design speed of 440km/h at the operation altitude of 3km. This paper also presents a comprehensive detailing of the control systems of the tilt-rotor unmanned air vehicle (UAV).

Key words: troubleshooting, tilt-rotor, flight test, flight control, vertical take-off and landing (VTOL)

1. Introduction was strongly inspired by Bell’s TR-911X, the first UAV of this kind [4, 5]. The TR-911X was a 7/8 scaled demonstrator of Smart unmanned air vehicle (UAV) is a tilt-rotor type aircraft that combines the ability to perform vertical take-off and landing while maintaining the ability to meet high speed cruise performance for the Korean civil UAV missions. These missions include coast guard, disaster detection, weather forecasting, and environmental monitoring, etc. It reaches

the maximum design speed of VTAS 440km/h at an altitude of 3km [1, 2]. A snapshot of the flight test in airplane mode of the Smart UAV is shown in Fig. 1. In the initial phase of the Smart UAV program, a trade- off was conducted among many VTOL configurations, and the tilt-rotor type achieved the best results in terms of performance for the civil missions, especially at high speeds

and high altitudes [3]. The development of the Smart UAV Fig. 1. FlightFig .test1 F lightin airplane test in airplane mode of mode the Smartof the SmartUAV [2] UAV [2]

This is an Open Access article distributed under the terms of the Creative Com- * Principal Researcher, Future Aircraft Research Division, Ph. D, Corre- mons Attribution Non-Commercial License (http://creativecommons.org/licenses/by- sponding author: [email protected] nc/3.0/) which permits unrestricted non-commercial use, distribution, and reproduc- ** Senior Researcher, Future Aircraft Research Division, Ph. D. tion in any medium, provided the original work is properly cited. *** Senior Researcher, Future Aircraft Research Division, Ph. D. **** Principal Researcher, Future Aircraft Research Division, Ph. D. ***** Principal Researcher, Future Aircraft Research Division, Ph. D. ****** Professor, Department of Aerospace Engineering, Ph.D. Received: April 20, 2015 Revised: March 19, 2016 Accepted: March 20, 2016 Copyright ⓒ The Korean Society for Aeronautical & Space Sciences 120 http://ijass.org pISSN: 2093-274x eISSN: 2093-2480

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Fig. 2 Unmanned Aircraft Systems of Smart UAV

22 Youngshin Kang Development of Flight Control System and Troubleshooting on Flight Test of a Tilt-Rotor Unmanned ....

the production model, TR-918 [6]. Most of the detailed this study experienced in the helicopter mode were the long troubleshooting on the flight test results of those unmanned delay on the ground in automatic take-off and the initial and manned tilt-rotor aircraft [7, 8] were undisclosed. overshoot when the automatic landing was engaged. The The basic specifications of the Smart UAV, TR-100, are anomalies are found in conversion mode as well. They were compared with those of the TR-918 Eagle Eye in Table 1 [5, 6]. the untrimmed AOS problem, which caused slow and large The Smart UAV is configured similarly to the benchmarked oscillation, and the calibration errors of air data sensors. Eagle Eye. The remarkable difference between the Smart Most of the problems related to autonomous functions were UAV and the Eagle Eye is the T-tail, which is intended to avoid found and solved in the earlier phases, and the only issue immersion of the tail into the rotor wake under a low speed that the airplane mode had was the low damping in rotor [4], and internal design. The maximum design speed of the speed and roll rate response. TR-100 is much greater than that of the TR-918. However, the The contribution of this paper is listed as follows: maximum payload of the latter is higher. 1) This paper presents the flight control system of the Before conducting the flight test of the full-scale Smart Smart UAV, its verification via the flight test, and the trouble- UAV, the control law of the tilt-rotor UAV was verified through shooting data. They will establish a better understanding of fully autonomous flights of a 40% scaled model named TR-40 the development and flight test of unconventional tilt-rotor [9-11]. The control law was designed using the general tilt- UAV. rotor nonlinear simulation model based on NASA technical 2) This paper elaborates on the flight control system and reports [12, 13]. The basic control surface mixers of the Smart anomalies associated with it, which are found during the UAV also followed the NASA reports in terms of modifications flight test of tilt-rotor type Smart UAV. The anomalies of any based on the flight test results of the TR-40 and Smart UAV. UAV have seldom been discussed in detail in the literature. Although the control law has been proven through the The troubleshooting procedure shown in this paper may flight test of the TR-40 [11], many unexpected problems inspire readers and help them avoid anomalies of their own were still inevitable due to the differences in scale and UAV and solve the problems when they encounter inevitable engine type. The troubleshooting was conducted mostly anomalies. in helicopter mode, the initial phase of flight test [14]. In The organization of the paper is as follows: Section II this stage, the vehicle encountered vortex ring state (VRS), provides a summary of the flight control system and control divergence in heading, large oscillation of vertical speed, modes. Section III elaborates on the anomalies and solutions and various sensor problems, etc. The VRS during the abrupt in helicopter, conversion, and airplane modes of flight test descent [14] will be revisited to explain the improvement of for the Smart UAV. The comparison of the flight envelops the control law on the heave axis. The other anomalies that between the Smart UAV and Eagle Eye will be presented. The

Table 1 Specification of the Smart UAV Table 1. Specification of the Smart UAV and Eagle Eye

Parameters TR-100 TR-918 Smart UAV Eagle Eye [5,6] Design gross weight (Kg) 1000 1179.3 Payload weight (Kg) 100 90-135 Wing span (m) 4 4.32 Length (m) 5 5.56 Height(m) 1.87 1.89 Rotor diameter (m) 2.864 3.05 Disc area per rotor (m2) 6.44 7.30 Rotor speed (rpm) in helicopter/conversion 1573 (98%) 1509* (100%) in airplane 1256 (82%) 1207* (80%) Conversion Range (0° at airplane) 0-93° 0-95° Maximum level flight speed (km/h) 440 370.4 Service ceiling(m) 6,000 6,096 Range (km) 200 185.2 Endurance (hrs) > 5 > 6 Power Plant PW 206C PW207D Maximum Power 550 HP 641 HP * estimated value based on the tip speed of rotor

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conclusion is presented in Section IV. where ρ and ρ0 are the air density at the altitude and seal level, respectively. Although the radar altimeter and ADC have single 2. Control System Developments redundancy as shown in Fig 2, the reliability of the single ADC is acceptable because it has a backup of GPS/INS 2.1 Operation and Control Systems sensors. The production type Smart UAV will be equipped The Smart UAV operation system consists of a ground with an upgraded dual DFCC, which contains the embedded control station (GCS), dual data-link system, and the tilt-rotor dual ADC function. The radar altimeter is solely used for aircraft as illustrated in Fig. 2. The figure shows the primary automatic landing and only determines the flare altitude where the descent rate fall below 1m/sec for smooth touch and the tilt--rotorrotor aircraftaircraft asas illustratedillustratedcomponents inin Fig.Fig. 2.. TheThe of figure figurethe flight showsshows control thethe primaryprimary system componentscomponents (FCS) as partsofof thethe of flightflight the aircraft system. The FCS consists of the digital flight control down. The failure of the radar altimeter is not fatal, because control system (FCS) as parts of thecomputer aircraft system.(DFCC), The GPS/INS FCS consists sensor, of radar the digitalaltimeter, flight air control data the failed radar altimeter is smoothly and automatically replaced by the GPS altitude. Once the radar altimeter fails, computer (DFCC), GPS/INS sensor,computer radar altimeter (ADC),, airair connected datadata computercomputer with (ADC)(ADC) the nose connectedconnected boom sensor,withwith thethe and nosenose the actuator control unit (ACU) integrated with 13 electric the flare altitude is determined using the difference between boom sensor, and the actuator controlactuators. unit (ACU) The integrated dual FCS with architecture 13 electric actuators. is designed The dualto haveFCS the specified touch down altitude and current GPS altitude. a primary and backup structure. The backup channel is Because the height error of GPS sensor is critical, the flare architecture is designed to have a primary and backup structure. The backup channel is activated when activated when the primary channel experiences any form altitude is set to be higher than the valid altitude from the thethe primaryprimary channelchannel experiencesexperiences anyanyof formform critical ofof criticalcritical failure. failure.failure. The TheTheDFCC DFCCDFCC and andand ACU ACUACU have havehave cross-channel crosscross--channel radar altimeter. It is noteworthy that radar altimeter just data links between the primary and backup channels, which measures the distance between the touch down point and data links between the primary and backup channels, which makes the channels mutually accessible makes the channels mutually accessible when obtaining the the aircraft. Therefore, the DFCC should know the GPS when obtaining the status, which inclustatus,des the which failure includes information the andfailure sensor information data. Both ofand the sensor DFCCs altitude of the touch down point in order to determine the data. Both of the DFCCs monitor the built-in test (BIT) flare altitude in case of radar altimeter failure. monitor the built--inin testtest (BIT)(BIT) signalssignalssignals ofof allall thetheof all sensorssensors the sensors andand thethe and ACU.ACU. the ACU. The dual ACUs control the dual channel electric actuators of which the controllers are newly developed because they The GPS/INS sensor and itsits NovatelThe GPS receiverGPS/INS have sensor dual andredundancy. its Novatel The GPS/INSGPS receiver is used have not dual redundancy. The GPS/INS is used not only as a are not available from the manufacturer [15]. The dual only as a navigation sensor but alsonavigation as a backup sensor for the butestimation also as of a calibrated backup forairspeed the estimation (VCAS),), truetrue ACU controllers of the left and right rotors simultaneously manage all three actuators for collective, longitudinal, and of calibrated airspeed (VCAS), true airspeed (VTAS), and the airspeed (VTASTAS),), and the altitude in the case of ADC failure. The estimated air speed can be calculated altitude in the case of ADC failure. The estimated air speed lateral cyclic. The flaperon ACU and conversion ACU manage as Eqss.. (1)(1) and (2).(2). can be calculated as Eqs. (1) and (2). only the flaperon and the nacelle conversion actuator, respectively. The elevator actuator has a single channel estest 22 22 22 (1) VTASTAS ≈ ((u + v + w )) (1)(1) because it requires only a small volume for installation. Therefore, the elevator surface is separated into left and ρ right sides, where each consists of independent actuators estest estest Fig. 1 Flight test in airplane mode of the Smart UAV [2] VCAS = VTASTAS (2)(2)(2) ρ00 for dual redundancy. The throttle actuator is also chosen to

where ρ and ρ00 are the air density at the altitude and seal level,, respectivelyrespectively..

Although the radarradar altimeteraltimeter and ADC have single redundancy as shown in Fig 2, the reliability of thethe singlesingle ADCADC isis acceptable because it has a backup of GPS/INS sensors.. TheThe productionproduction typetype SmartSmart

UAV will be equipped with an upgraded dual DFCC, which contains the embedded dual ADC function.function. TheThe radarradar altimeteraltimeter isis solelysolely usedused forfor automaticautomatic landinglanding andand onlyonly determinesdetermines thethe flareflare altitudeitude wherewhere thethe descentdescent raterate fallfall below 1m/sec for smooth touch down. The failure of the radarradar altimeter isis notnot fatal,fatal, because thethe failedfailed radarradar altimeteraltimeter isis smoothlysmoothly andand automaticallyautomatically replacedreplaced byby thethe

GPS altitude. Once the radarradar altimeteraltimeter failsfails,, thethe flareflare altitudealtitude isis determined using the difference between the specifiedspecified touchtouch downdown altitudealtitude andand currentcurrent GPSGPS altitude.altitude. BecauseBecause thethe heightheight errorerror ofof GPSGPS sensorsensor isis critical,critical, thethe flareflare altitudealtitude isisFig.setset 2. toto Unmanned bebe higher Aircraftthan thethe Systems validvalid altitudealtitude of Smart fromfrom UAV thetheFig radarradar. 2 Unmanne alaltimetertimeterd Aircraft.. ItIt Systems of Smart UAV isis nnoteworthy thatthat radarradar altimeteraltimeter justjust measuresmeasures thethe distance between thethe touchtouch downdown pointpoint andand thethe 4 DOI: http://dx.doi.org/10.5139/IJASS.2016.17.1.1204 122

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have single redundancy because its high level of reliability designed using control moment derivatives, i.e. the elements allows the possibility for the second channel to be removed. of B matrix of the linear model, to yield gradual increase The production type will be upgraded with all dual channel according to the increase of speed [16]. This behavior is actuators to gain a higher level of reliability. similar to a general fixed wing aircraft model. The GCS provides the control function for both the The primary control modes of the Smart UAV is presented external pilot (EP) and internal pilot (IP) as listed in Table 2. in Table 2. The attitude stability and control augmentation The EP can use only the external pilot box (PBX) for manual system (SCAS) is the basic augmented controller of the Smart control while the IP can use both of the PBX and the pilot bay UAV. The longitudinal and lateral stick controls generate (PBY) inside the GCS for all the control interfaces including the pitch and roll attitude commands, respectively. The the knob control and guidance function as shown in the left directional stick generates the yaw rate command, which is side of Fig. 2. combined with the sideslip (β) feedback. The heading-hold is engaged when the directional stick command is at neutral. 2.2 Control Law Development The vertical stick generates the simple collective and throttle The most fundamental controller of the tilt-rotor aircraft command with the vertical speed augmentation. The rate is the control surface mixer because the pilot has only 4 SCAS, named as SAS stick mode, can be used for manual stick commands, i.e. longitudinal, lateral, directional, and mode in case of a failure of attitude and speed sensors. collective, to control 13 actuators, i.e. 6 for rotors, 2 for In order to provide the EP with easier hover control using nacelles, 2 for flaperons, 2 for elevators, and 1 for the power a longitudinal and lateral ground speed command, the GPS lever. Therefore, the control surface mixer is designed to stick mode can be engaged by the IP when he or she pushes allocate the control stick commands to the actuators of the the button of the touch panel of the PBY. If the EP, who has control surfaces according to the airspeed or nacelle tilt the control authority, releases the stick into neutral center angle as shown in Fig. 3. position then the aircraft keeps hovering at the current The initial control surface mixer of the Smart UAV is position. This scheme is similar to the inertial velocity mode designed similarly to that of the Bell XV-15 [12] with the of the Eagle Eye [4]. The vertical stick command generates exception of the lateral cyclic and rudder. The XV-15 does not the altitude rate command and maintains the altitude if the have the lateral cyclic control, i.e. 6th and 7th outputs of Fig. stick command is neutral. 3, and the Smart UAV does not have the rudder surface. The The altitude, speed, heading, or roll hold modes can rudder is replaced with the 2deg. differential collective out be engaged and controlled by the touch panel and knob of total 48deg., i.e. 8th and 9th outputs of Fig. 3, similar to that dials of the PBY. In helicopter mode, where the nacelle tilt of Bell’s Eagle Eye [4, 5]. The mixer for control surface can be angle is greater than 80 degrees, the altitude is controlled

Table 2 Flight control modes of Smart UAV Table 2. Flight control modes of Smart UAV

Inter- Prio- Control mode Remark face rity SAS Stick 1 Rate and thrust command (Autopilot Off) EP/ SCAS Stick IP’s 2 Attitude and vertical speed command (Autopilot On) PBX Ground speed command and position hold 4 GPS Stick (VCAS<100km/h) Return home Loss of link 2 Collision Avoid Intruder distance<1km (ADS-B based) Airspeed hold GPS speed hold for VGPS<50 km/h Baro-altitude hold GPS altitude hold for V <50 km/h 3 GPS Heading hold Prior to roll hold PBY Roll hold Disengaged VGPS<50 km/h for Auto hover Approach and Hover (VCAS<100km/h) IP Auto take off VGPS<50 km/h Auto landing VGPS<50 km/h 4 Point Turn Fade out near hover speed Preprogrammed Climb turn is available Camera Guide Fade out near hover speed Auto recovery Linked return home

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by both the engine power and collective pitch, while the stick command is not controllable under the situation of link airspeed is controlled by the pitch attitude. The altitude loss or collision avoidance. and speed controls are replaced with the pitch attitude The guidance controls follow the Line-Of-Sight (LOS) and power controls, respectively, in airplane mode [2, 4]. method [21]. The common structure of the LOS guidance law These automatic changes in control authority are scheduled is applied to each of the helicopter, conversion, and airplane according to airspeed. Therefore, the altitude and speed modes. The buttons for the guidance command are activated holds are engaged simultaneously when the IP pushes when the IP establishes the required navigation information either the altitude or speed hold button. The heading hold using the map screen on the upper side of the PBY. When generates turn coordination commands, which includes the IP pushes any button on the touch panel, the aircraft flies the roll attitude command and corresponding yaw rate autonomously and follows the given flight path using the command at a high speed of helicopter mode. The roll provided information. command generated by the heading hold mode fades out gradually at a low speed close to hover. The guidance laws consist of auto-hover, auto-take- 3. Flight Test Results and Troubleshooting off [17, 18], auto-landing [17, 18], point turn [1, 18], preprogrammed [2, 17, 18], camera guide [19], auto-recovery Most flight tests were performed under the guidance law [17], return home [17], and collision avoidance [20] as listed that was explained in Chapter 2, with the exception of the in Table 2, where a lower number indicates higher priority. first flight and the low speed turn-around flight near the Therefore, the return home and collision avoidance modes take-off and landing position of the initial phase. have priority over the other guidance modes. The SCAS stick mode and the emergent guidance modes are of the same 3.1 Helicopter Mode level of priority. This is because if the SCAS stick mode is not activated, then the return home and collision avoidance The flight test in helicopter mode of the Smart UAV modes cannot be engaged autonomously whereas SCAS consists of manual and autonomous modes by the EP and

Fig. 3. Control surface mixer of Smart UAV Fig. 3 Control surface mixer of Smart UAV

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23 Youngshin Kang Development of Flight Control System and Troubleshooting on Flight Test of a Tilt-Rotor Unmanned ....

IP, respectively. The EP performed the manual take-offs and of the aircraft followed by a significant left roll due to the landings from hover, low-speed turns. The IP performed the VRS as shown in the first plot of Fig. 4 (b). The descent rate same sorties using the autonomous function. The number of reached about -500m/min, as shown in the third plot of Fig. 4 manual sorties was minimized due to the high workload of (b), before the EP abruptly took control using the “authority” the EP, and most flights were performed by the IP. Because switch of his PBX as shown in the first plot of Fig. 4 (b). Note the autonomous take-off and landing functions were verified that turning this switch enables the EP to override the whole during the early phase of the flight test, all the subsequent control authority of the IP. The altitude rate feedback in the flight test missions were performed using the autonomous basic SCAS mode generated the maximum power command functions in order to reduce the level of risk in the test [17, to recover the descent rate, and the rotor governor decreased 18]. the collective angle in order to reduce the surge in torque of the engine as shown in the second plot of Fig. 4 (b). These automatic recovery procedures conducted by the 3.1.1 Recovery from Vortex Ring State (VRS) SCAS controller and the EP barely saved the aircraft from The most dangerous situation throughout the whole the VRS by taking over the control authority from the IP. The flight test sorties of the Smart UAV occurred during the first reason for the unexpected climb was later found to be that automatic flight trial [14]. The aircraft climbed slowly under the lower limit of the throttle command in the altitude hold the altitude and speed hold mode as shown in Fig. 4 (a) controller was set higher than the required power. The lower when the IP took over the control from the EP and increased limit was scheduled in terms of the tilt angle which began the speed while holding altitude at 140m. The climb rate to decrease from a speed of 50km/h. Therefore, the altitude was so small that nobody recognized it until the altitude hold controller could not reduce the power to descend at reached 200m. As soon as the IP noticed it, he attempted to around this speed of which the power required is much decrease the speed, trying to recover the altitude. However, lower than that of hover as shown in Fig. 4 (c). his command was too excessive and resulted in the plunge The architecture of the altitude hold mode was modified

UncommandedUncommanded Climb Climb Uncommanded Climb Escape fromEscape VRS from VRS Escape from VRS 80 80 100 100 Φ Φ Φ 80 100 c md c md c md Φ Φ 60 60 60 50 50 50 Φ [km/h] [km/h] [km/h] Autority SWAutority SW Autority SW [deg] [deg] 40 40 [deg] 40 Φ Φ

Φ 0 CAS 0 CAS 0 CAS V V V 20 20 20 -50 -50 -50 2460 24602470 24702480 248024902460 249025002470 2500 2480 2490 2500 2600 2610 2620 2630 2640∆ H H 2600 26002610 26102620HGPS 26202630 26302640∆ PLA 2640∆ PLA PLA 200 200 200 GPS GPS20 20 20 θ θ θ 0G 0G 0G 180 180 180 10 10 10 H [m] H [m] H [m] 160 160 160 0 0 0 PLA [deg] PLA [deg] PLA [deg] 140 H H 140 140 -10 -10 -10 HGPS GPS GPS 2460 24602470 24702480 248024902460 249025002470 2500 26002480 260024902610 261026202500 262026302600 263026402610 2640 2620 2630 2640 200 200 200 Hdot Hdot Hdot Hdot 400 400 Hdot 400 GPS GPS GPS HdotGPS GPS GPS 200 200 100 100 100 200 0 0 0 0 0 0 -200 -200 -200 -400 -400 -400 dH/dt [m/min] dH/dt dH/dt [m/min] dH/dt dH/dt [m/min] dH/dt dH/dt [m/min] dH/dt dH/dt [m/min] dH/dt -100 -100 -100 [m/min] dH/dt -600 -600 -600 2460 24602470 24702480 248024902460 249025002470 2500 26002480 260024902610 261026202500 262026302600 263026402610 2640 2620 2630 2640 Time [sec]Time [sec] Time [sec] Time [sec]Time [sec] Time [sec]

(a) Slow(a) ascending Slow ascending due to thedue limit to(a) the Slowerror limit ascending error(b) Escape (b)due Escapefrom to the vortex limitfrom errorringvortex state ring(b) by stateEscape SCAS by fromSCAS vortex ring state by SCAS

Power LimitPower Limit Power Limit 500 500 500

400 400 400 HP HP HPREQ REQ REQ old old HPold HP HPLIM LIM LIM 300 300 300 HP HP HP new new HPnew HP HPLIM LIM LIM

200 200 200

Flight IdleFlight Idle Flight Idle 100 100 100 0 1000 100200 200300 300400 0 400500100 500200 300 400 500 V [km/h] V [km/h] VTAS [km/h]TAS TAS

Fig. 4. Entry into and recovery from theFig vortex. 4 EntryFig ring. 4 into Entrystate and [14]into recovery and recovery fromFig. the4 fEntryrom vortex the into ringvortex and state recovery ring[1 state4] from[14] the vortex ring state [14]

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24 24 24 The most dangerous situation throughout the whole flight test sorties of the Smart UAV occurred The most dangerous situation throughout the whole flight test sorties of the Smart UAV occurred during the first automatic flight trial [14]. The aircraft climbed slowly under the altitude and speed during the first automatic flight trial [14]. The aircraft climbed slowly under the altitude and speed hold mode as shown in Fig. 4 (a) when the IP took over the control from the EP and increased the hold mode as shown in Fig. 4 (a) when the IP took over the control from the EP and increased the speed while holding altitude at 140m. The climb rate was so small that nobody recognized it until the speed while holding altitude at 140m. The climb rate was so small that nobody recognized it until the altitude reached 200m. As soon as the IP noticed it, he attempted to decrease the speed, trying to altitude reached 200m. As soon as the IP noticed it, he attempted to decrease the speed, trying to recover the altitude. However, his command was too excessive and resulted in the plunge of the recover the altitude. However, his command was too excessive and resulted in the plunge of the aircraft followed by a significant left roll due to the VRS as shown in the first plot of Fig. 4 (b). The aircraft followed by a significant left roll due to the VRS as shown in the first plot of Fig. 4 (b). The descent rate reached about -500m/min, as shown in the third plot of Fig. 4 (b), before the EP abruptly descent rate reached about -500m/min, as shown in the third plot of Fig. 4 (b), before the EP abruptly took control using the “authority” switch of his PBX as shown in the first plot of Fig. 4 (b). Note that took control using the “authority” switch of his PBX as shown in the first plot of Fig. 4 (b). Note that turning this switch enables the EP to override the whole control authority of the IP. The altitude rate turning this switch enables the EP to override the whole control authority of the IP. The altitude rate feedback in the basic SCAS mode generated the maximum power command to recover the descent feedback in the basic SCAS mode generated the maximum power command to recover the descent rate, and the rotor governor decreased the collective angle in order to reduce the surge in torque of the rate, and the rotor governor decreased the collective angle in order to reduce the surge in torque of the engine as shown in the second plot of Fig. 4 (b). engine as shown in the second plot of Fig. 4 (b). These automatic recovery procedures conducted by the SCAS controller and the EP barely saved These automatic recovery procedures conducted by the SCAS controller and the EP barely saved the aircraft from the VRS by taking over the control authority from the IP. The reason for the the aircraft from the VRS by taking over the control authority from the IP. The reason for the unexpected climb was later found to be that the lower limit of the throttle command in the altitude unexpected climb was later found to be that the lower limit of the throttle command in the altitude hold controller was set higher than the required power. The lower limit was scheduled in terms of the hold controller was set higher than the required power. The lower limit was scheduled in terms of the tilt angle which began to decrease from a speed of 50km/h. Therefore, the altitude hold controller tilt angle which began to decrease from a speed of 50km/h. Therefore, the altitude hold controller could not reduce the power to descend at around this speed of which the power required is much could not reduce the power to descend at around this speed of which the power required is much Int’l J. of Aeronautical & Space Sci. 17(1), 120–131 (2016) lower than that of hover as shown in Fig. 4 (c). lower than that of hover as shown in Fig. 4 (c). The architecture of the altitude hold mode was modified from the altitude error command with the The architecture of the altitude holdfrom mode the was altitude modified error from command the altitude with error thecommand altitude with rate the were based on the (WGS 84) ellipsoid and the geoid (relative altitude rate damper to the altitude ratedamper error command to the altitude system ofrate Eqs. error (3) andcommand (4) [16] .system of Eqs. to the mean sea level) [22], respectively. altitude rate damper to the altitude rate error command system of Eqs. (3) and (4) [16]. (3) and (4) [16]. K 3.1.3 Initial Climb in Automatic Landing dhi δ =+Kdh ⋅− δCOL=+K dh i ⋅−( hhcmd ) (3) COLK dh s ( hhcmd ) (3) During the first automatic landing trial, the aircraft s climbed up for few seconds, and started descent for touch where where where down as shown in the second plot of Fig. 6 (a). Although this problem was not critical or dangerous, it was not easy to  hcmd= Kh h( cmd − h) (4)(4) hcmd= Kh h( cmd − h) (4) solve in the initial phase of the helicopter flight test. The root cause of this problem was identified that the complicatedly In addition, the lower limit of the throttle command of the entangled sequence of the altitude hold decision logics lost altitude hold8 mode was brought down to the flight idle power 8 one frame of altitude hold mode when the IP engaged the as shown in Fig. 4(c). This means that the altitude hold mode automatic landing. The altitude hold mode should not be has the full authority throughout the flight envelopes. The disengaged for any changes of autonomous control mode instability of engine operation under low power conditions during the automatic flight. The altitude hold mode was can be prevented by applying the altitude rate limit (= -1.5 set prior to the determination of automatic take-off and m/sec) which is chosen to avoid VRS during the descent in landing. Therefore, if the IP engaged the automatic take-off helicopter mode. or landing, the altitude hold mode was disengaged within 1 frame (= 0.02sec) before the engagement. This caused an 3.1.2 Long Delay in Automatic Take-off instant shift and return of the throttle command between the At the initial phase of the automatic flight test, the aircraft vertical stick command of PBX and altitude hold command. could not instantly follow the take-off command from the IP, However, this anomaly did not affect the automatic take-off instead it delayed approximately 5 seconds as shown in the because of the time delay in the initial climb. This problem second plot of Fig. 5 (a). was fixed with the change of the sequence to determine the In order to decrease the time delay on the ground for altitude hold mode at the last decision logic. The improved the automatic take-off, an initial value of the integrator in result was shown in Fig. 6 (b). collective control of Eq. (3) is added when it was engaged by the amount of hover power. As a result, the time delay was 3.2 Conversion Mode decreased to less than 1.8 second as shown in the second plot of Fig. 5 (b). Further reduction in delay could not be The flight test in conversion mode consists of flights under achieved due to the rate limiter of the power lever actuator the speed, altitude and heading or roll hold modes for every as shown in 10th output path of Fig. 3. 10deg. tilt angles from 80deg. (= helicopter mode) to 0deg. Note that the ground altitudes in the first plots of Fig. 5 (=airplane mode). The acceleration and deceleration, climb (a) and (b) exhibit significant difference because they have and descent, left and right turns were performed at the same distinct references of GPS altitude; the former and the latter tilt angle interval. [23]

Auto Take-off Auto Take-off HGPS 90 30 HGPS 80 20 70

ALT [m] 10 ALT [m] 60 Hdot 50 GPS 0 Hdot 1550 1560 1570 1580 1590 1600 2110 2120 2130 2140 2150 2160GPS Auto.TO 2 2 Auto.TO

0 0 dH/dt [m/sec] dH/dt

dH/dt [m/sec] dH/dt Power Lever -2 Power Lever -2 2110 2120 2130 2140 2150 θ 2160 1550 1560 1570 1580 1590 θ 1600 0 15 0 15

10 10

5 5 Collective [deg] Collective [deg] 0 0 1550 1560 1570 1580 1590 1600 2110 2120 2130 2140 2150 2160 Time [sec] Time [sec]

(a) Long delay at automatic take-off (b) Shorten delay at automatic take-off

Fig. 5. Improvement of the automatic take-off Fig. 5 Improvement of the automatic take-off

DOI: http://dx.doi.org/10.5139/IJASS.2016.17.1.120Auto Landing 126 Auto Landing 90 60 H HGPS 80 GPS 40 70

ALT [m] 20 60 ALT [m] 50 0 3380 3390 3400 3410 3420 3430 Hdot3440 GPS 3240 3250 3260 3270 3280 3290 Hdot3300 2 GPS Auto.Land 2 Auto.Land

0 0 (120~131)15-069.indd 126 2016-04-04 오전 10:58:25 dH/dt [m/sec] dH/dt -2 [m/sec] dH/dt Power Lever 3380 3390 3400 3410 3420 3430 Power3440 Lever -2 3240 3250 3260 3270 3280 3290 θ 3300 15 θ 0 0 15

10 10

5 5 Collective [deg] 0 Collective [deg] 0 3380 3390 3400 3410 3420 3430 3440 3240 3250 3260 3270 3280 3290 3300 Time [sec] Time [sec]

(a) Initial climb at auto landing (b) Initial descent at auto landing

Fig. 6 Improvement of the automatic landing

25 Youngshin Kang Development of Flight Control System and Troubleshooting on Flight Test of a Tilt-Rotor Unmanned ....

3.2.1 Large Sideslip Oscillation was completed in two sorties of flights after the vertical tail During the conversion tests in the envelop expansion, the modification. slow and large angles of sideslip (AOS) oscillations occurred around a tilt angle of 75deg. The attempt to reduce the 3.2.2 Calibration Errors in True Air Speed AOS oscillation through the yaw control, which increases The true airspeed are calibrated from the commercial

the differential longitudinal cyclic pitch (i.e.δ B1), was not air data system through the flight test results as shown effective because of the hysteresis phenomenon that is Auto Take-offAuto Take-off in Fig. 8 (a)-(b).Auto Take-offAuto Take-off The errors in airspeed and barometric HGPS HGPS 90 90 H H 30 30 shown in Fig. 7 (a). The integrator component of the AOSGPS GPS altitude result mostly from the pressure measurement error 80 80 20 20 feedback in the yaw axis failed70 70 to trim out the AOS because of of pitot-static tube due to its installation configuration

ALT [m] 10 ALT [m] 10 ALT [m] 60 ALT [m] 60 Hdot Hdot the increase in δB1 (i.e. pushed50 50 heading to the left) to reduceGPS GPS 0 [24,0 25]. By comparing the HdotmeasuredHdot true airspeed (VTAS) 1550 1550 1560 1560 1570 1570 1580 1580 1590 1590 1600 1600 2110 2110 2120 2120 2130 2130 2140 2140 2150 2150 2160GPS2160GPS Auto.TOAuto.TO 2 2 2 2 Auto.TOAuto.TO negative AOS resulted incomparing excessive the heading measured overshoot true airspeed to a (VTAS)with with thethe estimated V VTASTAS fromfrom the the GPS GPS ground ground velocity, velocity, the large positive AOS, and vice0 versa.0 Therefore, the additional 0 dynamic0 pressure error was identified to be about 16.08% dH/dt [m/sec] dH/dt [m/sec] dH/dt

the [m/sec] dH/dt [m/sec] dH/dt dynamic pressure error was identified to be about 16.08% of measuredPowerPower Lever impact Lever pressure (qc) as vertical fins were attached-2 to-2 the tips of the horizontalPower tail PowerLever as Lever -2 -2 2110of2110 measured2120 2120 2130 impact2130 2140 2140 pressure2150 2150θ 2160 θ(q2160c) as shown in Fig. 8 (a). The 1550 1550 1560 1560 1570 1570 1580 1580 1590 1590θ 1600θ 1600 0 0 15 15 0 0 15 15 shown in Fig. 7 (b). shown in Fig. 8 (a). The impact pressure errorimpact (Δq) pressure was measure error d(Δq)in constantwas measured bank turn in sconstantat the bank 10 10 10 10 Although the original vertical tail is only half the size turns at the airspeeds of 110, 150, 190, and 250km/h. The 5 5 5 5

airspeeds of 110, 150, 190, and 250km/hCollective [deg] . TheCollective [deg] measurement was taken twice at speed of 150 and of the horizontal tail, theCollective [deg] windCollective [deg] tunnel test verified that it 0 0 0 measurement0 was taken twice at speed of 150 and 250km/h, 1550 1550 1560 1560 1570 1570 1580 1580 1590 1590 1600 1600 2110 2110 2120 2120 2130 2130 2140 2140 2150 2150 2160 2160 provides sufficient directional stability.Time [sec]Time Therefore,[sec] the size Time [sec]Time [sec] 250km/h, because of the node points of nacellebecause conversion of the schedul node pointse at 70deg. of nacelle and 0deg. conversionThe error schedule of at of the vertical tail remained unchanged until the flight 70deg. and 0deg. The error of impact pressure, Δq, was fitted (a) Long(a) Long delay delay at auto at automaticmatictaketake-off-off (b) Shorten(b) Shorten delay delay at automatic at automatic take take-off-off demonstration which impactrevealed pressure, its poorΔq, was performance. fitted by the least squareby the method least square as expressed method in Eq. as expressed(7). in Eq. (5). The additional vertical fins resolved theFig .Fig5AOSImprovement. 5 Improvementhysteresis of the of automaticthe automatic take take-off-off ∆ ≈ × (7) (5) issues without any side effects. The full conversion test q 0.1608 qc

As a result, the corrected VTAS is calibrated as Eq. (8) and shown in Fig. 8 (b). Auto LandingAuto Landing Auto LandingAuto Landing 90 90 60 60 H H HGPS HGPS 80 80 GPS GPS 40 cal40 70 70 VVTAS =1.1608 × TAS (8)

ALT [m] ALT [m] 20 20 60 60 ALT [m] ALT [m] 50 50 0 0 3380 33803390 33903400 34003410 34103420 34203430 3430Hdot3440Hdot3440 GPS GPS 3240 32403250 32503260 32603270 32703280 32803290 3290Hdot3300Hdot3300 2 2 GPS GPS Auto.LandAuto.Land 2 2 3.3 Airplane Mode Auto.LandAuto.Land

0 0 No critical problems in handling qualities0 occurred0 during the flight tests in airplane mode. This dH/dt [m/sec] dH/dt [m/sec] dH/dt -2 -2 [m/sec] dH/dt [m/sec] dH/dt PowerPower Lever Lever 3380 33803390 33903400 34003410 34103420 34203430 Power34303440Power Lever3440 Lever -2 -2 3240 32403250 32503260 32603270 32703280 32803290 θ32903300θ 3300 15 15 θ θ 0 0 was because the fidelity of nonlinear0 0 dynamic15 15 model was high enough to design an accurate control

10 10 10 10

law5 5of Smart UAV [13]. The only issues were5 5the tendencies of low damping in rotor speed and roll Collective [deg] Collective [deg] 0 0 Collective [deg] 0 Collective [deg] 0 3380 33803390 33903400 34003410 34103420 34203430 34303440 3440 3240 32403250 32503260 32603270 32703280 32803290 32903300 3300 rate responses. Time [sec]Time [sec] Time [sec]Time [sec]

3.3.1 Rotor(a) Initial(a) RPM Initial climb O climbscillation at auto at auto landing landing (b) Initial(b) Initial descent descent at auto at auto landing landing

Fig. 6. Improvement of the automaticA noteworthy landing tendencyFig.Fig6 ofImprovement. 6TRImprovement-100 was of the theof automaticthe oscillation automatic landing oflanding rotor speed in the airplane mode as shown

in Fig. 9 (a). TheAOSAOS Hysteresismaximum Hysteresis overshoot of the RPM in airplane mode was approximately ±2%, which 15 15

10 caused10 distinctive rotor noise. One of the root causes was that the collective angle was scheduled by

5 airspeed5 in order to reduce the amount of output in the integrator controller of the rotor governor [12].

0 0 When the rotor speed increased, the collective (θ0) also increased due to the rotor governor (θ0G), AOS [deg] AOS [deg] -5 -5 25 25 which increased the airspeed (VCAS) and the scheduled collective as well. Consequently, the rotor -10 -10 speed dropped back due to the excessive collective, and vice versa, as shown in Fig. 9 (a). Therefore, -15 -15 -4 -4 -2 -2 0 0 2 2 4 4 6 6 8 8 10 1012 12 10δ 10 [deg]δ [deg] the parameter for B1collectiveB1 trim schedule was changed from the measured airspeed to the airspeed

command of(a)the AOS(a)IP AOS hysteresiswhich hysteresis has rate limiter . The other(b) Vertical(b) cause Vertical finof extensionsfinthis extensions issue was the low value of the

Fig. 7. Vertical fins on the outsideproportionalFig.Fig7 Verticalof. 7horizontalVertical gainfins fins on intail theonthe to theoutside rotorovercome outside speedof horizontal ofAOS horizontalfeedback hysteresis tail tailto. The overcome to overcomegain wasAOS AOS scheduled hysteresis hysteresis to be reduced in airplane mode

in order to avoid the speed sensitivity which is caused by the active collective command (θ0G) of the 127 http://ijass.org 11

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26 26 comparing the measured true airspeed (VTAS) with the estimated VTAS from the GPS ground velocity, the dynamic pressure error was identified to be about 16.08% of measured impact pressure (qc) as shown in Fig. 8 (a). The impact pressure error (Δq) was measured in constant bank turns at the airspeeds of 110, 150, 190, and 250km/h. The measurement was taken twice at speed of 150 and

250km/h, because of the node points of nacelle conversion schedule at 70deg. and 0deg. The error of impact pressure, Δq, was fitted by the leastInt’l square J. of Aeronautical method as expressed & Space Sci. in 17(1),Eq. ( 7120–131). (2016)

∆q ≈ 0.1608× qc (7)

As a result, the corrected VTAS is calibrated as Eq. (6) and was scheduled by airspeed in order to reduce the amount As a result, the corrected V is calibrated as Eq. (8) and shown in Fig. 8 (b). TAS shown in Fig. 8 (b). of output in the integrator controller of the rotor governor

[12]. When the rotor speed increased, the collective (θ0) cal (6) VVTAS =1.1608 × TAS (8) also increased due to the rotor governor (θ0G), which

increased the airspeed (VCAS) and the scheduled collective 3.3 Airplane Mode 3.3 Airplane Mode as well. Consequently, the rotor speed dropped back due to the excessive collective, and vice versa, as shown in Fig. No critical problems in handling qualities occurred during the flight tests in airplane mode. This No critical problems in handling qualities occurred during 9 (a). Therefore, the parameter for collective trim schedule the flight tests in airplane mode. This was because the fidelity ∆ was because the fidelity of nonlinear dynamic model was high enough to design an accurate control was changed from qthec vs . measured q airspeed to the airspeed of nonlinear dynamic model was high enough to design an command500 of the IP which has rate limiter. The other cause law of Smart UAV [13]. The only issues were the tendencies of low damping in rotor speed and roll ∆ q accurate control law of Smart UAV [13]. The only issues were of this issue was the low value of the proportional gain in the 400 Lleast Square Fit rate responses. the tendencies of low damping in rotor speed and roll rate rotor speed feedback. The gain was scheduled to be reduced responses. in300 airplane mode in order to avoid the speed sensitivity q

3.3.1 Rotor RPM Oscillation ∆ which is caused by the active collective command (θ0G) of the 200 3.3.1 Rotor RPM Oscillation ∆q = 0.1608 x q A noteworthy tendency of TR-100 was the oscillation of rotor speed in the airplane mode as shown rotor governor [12]. c A noteworthy tendency of TR-100 was the oscillation 100The final rotor speed overshoot was reduced to under in Fig. 9 (a). The maximum overshoot ofof the rotor RPM speed in airplane in the mode airplane was approximatelymode as shown ±2% in, Fig.which 9 (a). ±0.5% RPM at the same airspeed region as shown in Fig. 0 The maximum overshoot of the RPM in airplane mode 9 (b).0 Note500 that1000 Fig. 91500 (b) shows2000 higher2500 3000 airspeed command caused distinctive rotor noise. One of the root causes was that the collective angle was scheduled by q was approximately ±2%, which caused distinctive rotor to match the airspeedc with Fig. 9 (a) after calibration of airspeed in order to reduce the amount of noise.output Onein the of integrator the root controller causes was of the thatrotor the governor collective [12] angle. aerodynamic(a) Dynamic lift coefficient. pressure correction

When the rotor speed increased, the collective (θ0) also increased due to the rotor governor (θ0G), 300 V ∆ GPS qc vs . q which increased the airspeed (VCAS) and the scheduled collective as well. Consequently, the rotor Calibrated V 500 TAS 250 ∆ V speed dropped back due to the excessive collective, and vice versa, as shown q in Fig. 9 (a). Therefore, TAS 400 Lleast Square Fit 200 the parameter for collective trim schedule was changed from the measured airspeed to the airspeed 300 q ∆ 150 command of the IP which has rate limiter. The other cause of this issue was the low value of the Speed [KPH] 200 ∆ q = 0.1608 x qc proportional gain in the rotor speed feedback. The gain was scheduled to be reduced in airplane mode 100 100 in order to avoid the speed sensitivity which is caused by the active collective command (θ0G) of the 0 0 500 1000 1500 2000 2500 3000 50 3400 3600 3800 4000 4200 4400 q c Time [sec] 11

(a) Dynamic pressure correction (b) Calibration results of true airspeed (VTAS)

Fig. 8. Calibration results of true air speed 300 Fig. 8 Calibration results of true air speed V GPS Calibrated V RPM Oscillation RPM RPM Oscillation RPM 100 TAS TA RGET 100 TA RGET 250 RPM V RPM TAS 98 98 RPM [%] RPM [%] θ -28 200 0 96 θ 96 θ 1684 1686 1688 1690 1692 16940-28 1970 1972 1974 1976 1978 19800G 1 1 θ 0G 150 Speed [KPH] 0 0 Collective [deg] Collective [deg] -1 -1 COM 1684 1686 1688 1690 1692 1694 1970 1972 1974 1976 1978 1980V 100 COM T 310 VT 315 VCA S V 305 CA S 310 [km/h] [km/h] 50 300 305 CAS 3400 3600 3800 4000 4200 4400 CAS V V 295 Time [sec] 300 1684 1686 1688 1690 1692 1694 1970 1972 1974 197627 1978 1980 Time [sec] Time [sec] (b) Calibration results of true airspeed (VTAS) (a) Large rotor speed oscillation (b) Improved rotor speed oscillation Fig. 8 Calibration results of true air speed Fig. 9. Rotor speed oscillation in airplane mode Fig. 9 Rotor speed oscillation in airplane mode

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28 Youngshin Kang Development of Flight Control System and Troubleshooting on Flight Test of a Tilt-Rotor Unmanned ....

3.3.2 Slightly Low Damping in Roll Axis 3.4 Comparison with Flight Envelop of Eagle Eye The last anomaly of TR-100 was its slightly low damping The full flight envelope for the maximum speed sortie in the roll axis under a high speed in airplane mode as of the Smart UAV is compared with that of the TR-911X shown in Fig. 10 (a). This tendency is similar to that of the Eagle Eye in Fig. 11 [2, 5]. Note that the flight trajectories of TR-40, the 40% scaled model which showed low damping Eagle Eye are approximately regenerated according to the in the roll axis of airplane mode. The aerodynamic model flight test data [5]. The maximum true airspeed of the flight yielded an over estimated rolling moment parameter

against roll rate (Lp), so that the control gains were updated 10000 based on the parameter identification. The resultant Smart UAV 9000 Eagle Eye damping in roll axis was improved as shown in Fig. 10 8000 (b). The related stability margin of the both original and Smart UAV 7000 modified roll axis SCAS model are shown in Fig. 10 (c) Eagle Eye Engine Limit and (d). The original gain margin is 9.25dB and the phase 6000

margin is 56.6deg. as shown in Fig. 10 (c). These margins 5000

were reduced to the minimum requirement of 6dB and ALT[m] 4000 45deg. at the modified model in order to increase the roll damping by increasing roll axis feedback gains as shown in 3000 Fig. 10 (d). This means that the Smart UAV could meet the 2000

handing quality requirements in all flight envelopes with 1000 enough stability margins. Based on this experience, the 0 control gains of the other variants of the tilt-rotor aircraft 0 50 100 150 200 250 300 350 400 450 500 V [km/h] are designed to have optimized minimum gain margin (≥ TAS 6 dB) and phase margin (≥ 45 deg.) [16]. Fig. 11. FigComparison . 11 Comparison of flight of flight envelops envelops ofof SmartSmart UAV UAV and and Eagle Eagle Eye Eye [2, 5[2,] 5]

Roll Axis 35ΦΦ 35Φ Roll Axis Roll AxisRoll Axis 35 COMCOM COM Roll AxisRoll Axis 1010 10 ΦΦ Φ 2020 20 pp p 3535ΦΦ 35Φ 00 0 COMCOM COM 1010 10 ΦΦ Φ [deg] [deg] [deg] [deg] [deg] [deg] Φ Φ Φ Φ

Φ p Φ -10-10 -10 p p 0 0 COM 0 VVCOM VCOM -20-20 -20 TT T 19701970 1970 19751975 1975 19801980 1980 19851985 1985 19901990 1990 33103310 3310 33153315 3315 33203320 3320 33253325 3325 33303330 3330 V V 365365 365 VCACA S S CA S 365365 365

360360 360 360360 360 COM [km/h] COM COM [km/h] [km/h] [km/h] [km/h] [km/h] VVT V 355 T T

355 CAS 355 355 CAS CAS

355 CAS

355 CAS CAS V V V V

V V V

V CA S V CA S CA S 350350 350 350350 350 19701970 1970 19751975 1975 19801980 1980 19851985 1985 19901990 1990 33103310 3310 33153315 3315 33203320 3320 33253325 3325 33303330 3330 TimeTime [sec] [sec]Time [sec] TimeTime [sec] [sec]Time [sec]

(a)(a) LowLow(a) Low dampingdamping damping inin thethe in rollroll the axis axisroll axis (b)(b) ImprovedImproved(b) Improved dampingdamping damping inin thethe in rollroll the axis axisroll axis

BodeBode Diagram DiagramBode Diagram BodeBode Diagram DiagramBode Diagram GmGm = = 9.25 9.25Gm dB =dB 9.25 (at (at 17.4 dB17.4 (at rad/s) rad/s) 17.4 , ,rad/s) Pm Pm = =, 56.6 56.6Pm deg =deg 56.6 (at (at deg6.5 6.5 rad/s) (atrad/s) 6.5 rad/s) GmGm = = 6.07 6.07Gm dB=dB 6.07(at (at 17 17dB rad/s) rad/s)(at 17 , ,rad/s) Pm Pm = = ,45 45Pm deg deg = (at 45(at 9.31deg 9.31 (atrad/s) rad/s) 9.31 rad/s) 5050 50 100100 100

00 0 30 00 0 -50-50 -50

-100-100 -100 -100-100 -100 Magnitude (dB) Magnitude (dB) Magnitude (dB) Magnitude (dB) Magnitude (dB) Magnitude (dB) -150-150 -150

-200-200 -200 -200-200 -200 00 0 00 0

-90-90 -90 -90-90 -90

-180-180 -180 -180-180 -180 Phase (deg) Phase (deg) Phase (deg)

Phase (deg) -270 Phase (deg) Phase (deg) -270-270 -270 -270 -270

-360 -360 -360 -360 -360 -360 -2 0 2 4 -2-2 -2 00 0 22 2 44 4 -2 -2 0 0 2 2 4 4 1010 10 1010 10 1010 10 1010 10 1010 10 1010 10 1010 10 1010 10 FrequencyFrequencyFrequency (rad/s) (rad/s) (rad/s) FrequencyFrequencyFrequency (rad/s) (rad/s) (rad/s)

(c)(c) OriginalOriginal(c) Original sstabilitytability stability mmarginargin margin ininrollroll in axisaxisroll axis(d) (d) ModifiedModified(d) Modified sstabilitytability stabilitymmarginarginmargin inin rollroll in axis axisroll axis

Fig. 10. Damping of the roll axis in airplane modeFig.Fig. 1Fig.100DampingDamping 10 Damping ofof thethe of rollroll the axis axisroll inaxisin airplane airplane in airplane modemode mode

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2929 29 Int’l J. of Aeronautical & Space Sci. 17(1), 120–131 (2016)

envelop of the Smart UAV is over 440km/h which is greater production type of the tilt-rotor UAV, designated as the TR- by an amount of over 70km/h at an altitude of 3km than that 60 which is scaled down to roughly 60% of the TR-100. It of the Eagle Eye, of which the maximum speed is 370km/h will be used for the military and civil missions of the Korean (i.e. 200kts). As the maximum altitude limit is dependent on government. the engine limit of Fig. 11, it is expected that both aircrafts New research topics of the tilt-rotor UAV will be conducted yield similar levels of altitude performance, which could not as follows: First, the research will focus on improving the be verified due to airspace limitation in Korea. However, the range and endurance of the tilt-rotor UAV with the nacelle- tested flight envelope shows that the Smart UAV is capable of mounted wing extension study [16]. reaching its maximum speed of over 480km/h at an altitude Second, the ship deck operation capability will be of 5km. investigated based on the HILS. The simulation will be conducted using the unsteady wake model around the ship that is generated by a computational fluid dynamics (CFD) 4. Conclusion tool.

The flight test of the Smart UAV in helicopter, conversion, and airplane modes were completed. The Smart UAV Acknowledgement

accomplished a maximum speed of VTAS 440km/h at an altitude of 3km [2]. Many unexpected problems that occurred This study was supported by the research project on the during the envelope expansion positively contributed to Technology development for shipboard operation of 200kg improvement and led the Smart UAV to a better aircraft. class Tilt-Rotor UAV of the Ministry of Trade, Industry and The stable and controllable characteristics of the control Energy (MOTIE). The flight test results are the property law of the Smart UAV were verified in all flight envelopes. In of all members of the Smart UAV development center. The addition, it was shown that the vehicle was easy to handle for contributions with all their might led to the success of the the IP without receiving assistance from the EP due to the full Smart UAV program. autonomous functions, which include take-off and landing. Therefore, the proven process of control law design could be applied to the family of tilt-rotor type UAVs in KARI [16]. References The reliability of the newly developed FCS was evaluated under the dual redundancy and fault management system. [1] Kang, Y. S., Park, B. J., Cho, A., Yoo, C. S. and Koo, S. O., All the faults that occurred during the flight test were “Flight Test of Flight Control Performance for Airplane Mode th effectively isolated and successfully replaced with the of Smart UAV”, 12 Conference of International Conference on normal backup system. As a result, no fatal accident due to Control Automation and Systems, Jeju Island, Korea, 19 Oct., a single failure of the FCS system occurred during the flight 2012., pp.1738-1741. tests. The production type of the DFCC will be integrated [2] Kang, Y. S., Park, B. J., Cho, A. Yoo, C. S. and Koo, S. with the onboard ADC, resulting in the dual redundancy of O., “Analysis of Flight Test Result for Control Performance of the air data system for the dual DFCC. Smart UAV”, Aerospace Engineering and Technology, KARI, The hardware in the loop simulation (HILS) and tethered ISSN 1598-4168, Vol.12, No.1, July 2013, pp. 22-31. hover test that followed the changes of the OFP contributed [3] Ahn, O.S., “Advanced VTOL Concept for Smart UAV to the detection and filtering of many hidden software errors. Program”, Nagoya-KyungSang Aero-Conference, 2003. The IP, EP, and the flight control engineers trained themselves [4] Fortenbaugh, R. L., Builta, K. E., Schulte, K. J., during the HILS and changed the details of the flight plans “Development and Testing of Flying Qualities for Manual to enable a safer and more effective test. The fidelity of the Operation of a Tiltorotor UAV”, 51st Annual Forum of mathematical model used for the design of the control law American Helicopter Society, Fort Worth, TX, USA, May 9-11, was high enough to solve the flying quality problems during 1995., pp.299-320. [5] Settle, B. and Wise, T. “Bell Eagle Eye TR–911X – the flight test. The aerodynamic derivatives including CL, CD, Tiltrotor Unmanned Aerial Vehicle: Recent Developments, and Clp which showed some modeling error compared to the flight test results, were updated after the flight tests. Autoland Integration, and Flight Test Demonstrations”, 56th The Smart UAV program was completed successfully Annual Forum of American Helicopter Society. Virginia, USA, without any crash during the whole flight test, and the May 2-4, 2000, pp. 306-319. proven flight control technology is transferred to the new [6] Wyatt, D., Eagle Eye Pocket Guide, Bell Helicopter

DOI: http://dx.doi.org/10.5139/IJASS.2016.17.1.120 130

(120~131)15-069.indd 130 2016-04-04 오전 10:58:31 Youngshin Kang Development of Flight Control System and Troubleshooting on Flight Test of a Tilt-Rotor Unmanned ....

Textron Inc., 2004. [16] Kang, Y.S., Park, B.J., Cho, A., Yoo, C.S., “Control Law [7] Schareffer, J. Alwang, R. and Joglekar, M. “V-22 thrust Design for the Tilt-rotor Unmanned Aerial Vehicle with power management control law development”, 47th Annual Nacelle Mounted Wing Extension (WE)”, Journal of Institute Forum of the American Helicopter Society, Phoenix, Arizona, of Control, Robotics and Systems, 2014., Vol.20, No.11, pp. USA, 1991, pp. 1093-1100. 1103-1111. [8] Fortenbaugh, R., Hopkins, R. II and King, D., “BA609 [17] Kang, Y. S., Park, B. J., Cho, A., Yoo, C. S. and Koo, S. First Flight VSTOL Handling Qualities,” 60th Annual Forum O., “Guidance Flight Test for Helicopter Mode of Smart UAV”, of the American Helicopter Society, Baltimore, MD, USA, June The Korean Society for Aeronautical & Space Sciences (KSAS) 7-10, 2004, pp.800-815. Conference, FC2-3, 11 Nov. 2011, pp. 248-251. [9] Choi, S.W., Kang, Y.S., Chang, S.H., Koo, S.O., Kim, J.M., [18] Kang, Y. S., Park, B. J., Cho, A., Yoo, C. S. and Koo, S. O., “Development and Conversion Flight Test of a Small Tilt- “Automatic Flight Test to Reduce Test Risk of Smart UAV in rotor Unmanned Aerial Vehicle”, Journal of Aircraft, Vol. 47, Helicopter, Conversion, and Airplane Mode”, 1st Asia-Pacific No.2, 2010, pp. 730-731. International Symposium on Aerospace Technology, FC-I- [10] Kang, Y. S., Park, B. J., Yoo, C. S., Kim, Y. S. and 7.5.1., Jeju Korea., Nov. 2012. Koo, S. O., “Fully Automatic Flight Test of Small Scaled [19] Park, B. J., Kang, Y. S., Cho, A., Chang, S. H. and Yoo, Tilt Rotor Aircraft”, 2nd International Forum on Rotorcraft C. S., “Development of HILS System for Camera Guidance Multidisciplinary Technology, FP03-1, Oct. 19-20. 2009. pp. Mode Test of Smart UAV”, The Korean Society for Aeronautical 31-34. & Space Sciences (KSAS) conference, FG3-2, 11 April 2012. [11] Kang, Y. S., Park, B. J., Yoo, C. S., Koo, S.O. and Lee, J. pp.500-505 . H.“Trouble Shooting for Fully Automatic Flight Test of Small [20] Park, B. J., Yoo, C. S., Cho, A., Kang, Y. S. and Koo, S. O., Scaled Tiltrotor UAV”, Aerospace Engineering and Technology, “Development of HILS System for Collision Avoidance Test KARI, ISSN 1598-4168, Vol.8 No.1, 2009. pp. 1-9. of Smart UAV”, The Korean Society for Aeronautical & Space [12] Harendra, P.B., Joglkar, M.J., Gafey, T.M., Marr, R.L., Sciences (KSAS) conference, TB5-5, 11 Nov 2012. pp. 715-720. V/STOL Tilt Rotor Study - A Mathematical Model for Real [21] Fossen, T.I., Marine Control Systems – Guidance, Time Flight Simulation of The Bell 301 Tilt Rotor Research Navigation and Control of Ships, Rigs and Underwater Vehicles. Aircraft, NASA-CR-114614, NASA, April 1973. Marine Cybernetics (http://www.marinecybernetics.com), [13] Yoo, C. S., Choi, H. S., Park, B. J., Ahn, S. J., Kang, Y. S., 2002, Norway, pp. 167-169. “Development of Simulation Program for Tilt Rotor Aircraft”, [22] https://en.wikipedia.org/wiki/Undulation_of_the_geoid Journal of Institute of Control, Robotics and Systems, Republic [23] Kang, Y. S., Choi, S. W., Hwang, S. J., Ahn, O. S., of Korea, Vol. 11, No. 3, March 2005., pp. 193-199. Koo, S. O., Kim, J. M., “KARI Tiltrotor UAV Flight Test and [14] Kang, Y. S., Park, B. J., Cho, A., Yoo, C. S., Koo, S. O., Performance Enhancement Study”, 38th European Rotorcraft “Trouble Shooting of Anomalies during Flight Test of Smart Forum, UAS-082, Amsterdam, Netherland, 4-7 Sep 2012. U A V ”, The Korean Society for Aeronautical & Space Sciences [24] Wayne M. O., Aircraft Performance Flight Testing, (KSAS) conference, TA4-1, 11 April 2013. AFFTC-TIH-99-01, Technical Information Handbook, A, Sep. [15] Yoo, C. S., Ryu, S. D., Park, B. J., Kang, Y. S. and Jung, S. 2000., pp. 31-32. B., “Actuator Controller Based on Fuzzy Sliding Mode Control [25] Cho, A., Kang, Y. S., Park, B. J., Yoo, C. S., Koo, S. O., “Air of Tilt Rotor Unmanned Aerial Vehicle”, International Journal data System Calibration Using GPS Velocity information”, of Control, Automation and Systems, May 29 2014 (accepted 12th International Conference on Control, Automation and for publication). Systems in ICC, Jeju Island, Korea, Oct. 17-21, 2012.

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