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BASIC AERODYNAMIC CHARACTERISTICS OF MuPAL-αααα’s DLC SYSTEM

Koki Hozumi＊ and Kazuya Masui† National Aerospace Laboratory of Japan, 6-13-1 Osawa, Mitaka-shi,Tokyo 181-0015, JAPAN

ABSTRACT The performance of the Direct Lift Control

system of NAL’s new in-flight simulator, “MuPAL-α”, was evaluated by wind tunnel and flight tests. As a result, it was proved that: 1) The vertical acceleration generated by the DLC

system satisfies the design requirements; 2) The maximum rolling and yawing moments

created by asymmetric deflection of the DLC flaps are easily cancelled by aileron deflection, so malfunction of the DLC system has little effect on the safety of the aircraft.

NOMENCLATURE Ax : Axial Acceleration (g’s) Az : Normal Acceleration (g’s) CL : Lift Coefficient CLact : Actual Lift Coefficient CLss : CL at Perfect Trimmed Condition CD : Drag Coefficient CDact : Actual Drag Coefficient CDss : CD at Perfect Trimmed Condition CT : Thrust Coefficient Cmδe : Pitching Moment Elevator Derivative Cmδh : Pitching Moment Horizontal

Stabilizer Derivative q : Dynamic Pressure S : Wing Area (Reference Area) W : Weight of the Vehicle α : Angle of attack ∆CL : Increment of Lift Coefficient

due to DLC Flap Deflection ∆CD : Increment of Drag Coefficient

due to DLC Flap Deflection ∆Cl : Increment of Rolling Moment Coefficient

due to DLC Flap Deflection ∆Cm : Increment of Pitching Moment Coefficient

due to DLC Flap Deflection ∆Cn : Increment of Yawing Moment Coefficient

due to DLC Flap Deflection ＊ Senior Researcher, Flying Qualities Group,

Flight Systems Research Center † Group Leader, Flight Experiment Group,

Flight Systems Research Center Copyright © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All right reserved.

δDLC : DLC Flap Deflection Angle δe : Elevator Deflection Angle δf : Landing Flaps Deflection Angle δh : Horizontal Stabilizer Deflection Angle γ : Flight Path Angle Abbreviations AOA : Angle of Attack DLC : Direct Lift Control MuPAL : Multi-Purpose Aviation Laboratory NAL : National Aerospace Laboratory of Japan

INTRODUCTION

The National Aerospace Laboratory of Japan (NAL) has developed a new in-flight simulator, “MuPAL (Multi-Purpose Aviation Laboratory) -α” based on a Dornier 228-200 (Fig. 1)1). The aircraft is equipped with a Direct Lift Control (DLC) system to give three degree-of-freedom longitudinal motion control. During the development and after the completion of MuPAL-α, wind tunnel tests2) and flight tests were conduced to confirm the static aerodynamic characteristics of the DLC system. In this paper, the basic characteristics of the DLC system obtained through these tests are presented, and the performance of the system is evaluated.

Figure 1 MuPAL-αααα

41st Aerospace Sciences Meeting and Exhibit6-9 January 2003, Reno, Nevada

AIAA 2003-396

Copyright © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

2

DIRECT LIFT CONTROL SYSTEM The DLC system is essential for the accurate simulation of aircraft longitudinal motion. The design requirements for the MuPAL-α’s DLC system are as follows: - The maximum speed during DLC operation should

be 150KIAS(77m/s). - The vertical acceleration induced by the DLC

system should be greater than 0.2g at 100KIAS(51m/s), which is equivalent to 0.45g at 150KIAS.

- The frequency response bandwidth of the DLC actuator should be greater than 2Hz.

To realize these requirements, a flap-in-flap system is adopted which actuates the trailing edges of original landing flaps as DLC flaps, Fig. 2 shows a schematic of the DLC system. The DLC flaps occupy 30% of the chord and 80% of the span of the original landing flaps, and have 3 segments per wing. All DLC flaps deflect 55° around the “preset” position (the “neutral” position of the DLC flaps when the DLC system is engaged: 3.5° trailing edge up), and each DLC flap is independently controlled by an electrical actuator and its controller according to a common signal from the DLC computer.

Figure 2 DLC Flaps

WIND TUNNEL TESTS Test Method

Test were conducted in NAL’s low-speed wind tunnel, which has a 6.5m-5.5m test section, using a

1/6 scaled model with DLC flaps and powered propellers (Fig.3). The six components of forces and moments were measured at a wind speed of 35m/s at various combinations of DLC flap deflection, angle of attack (AOA) and sideslip angle. Propeller pitch angle and rotation speed were set so that the thrust generated by propellers balanced the drag acting on the whole aircraft when both AOAof and the sideslip angle were zero. This condition correspond to CT = 0.057. As a part of the proof of airworthiness in the case of DLC system malfunction, the effects of asymmetric deflection of the DLC flaps were also measured.

Wind Tunnel Test Results 1. Symmetric DLC flaps Deflection

Figures 4-6 show test results for symmetric DLC flap deflection. The change of lift coefficient (∆CL / CL) corresponding to full DLC flap deflection (δDLC = +25° /–30°) at the DLC system’s design point (AOA α = 8°, flap angle δf = 0°) is greater than ±0.2 if the preset angle of the DLC flaps is at –3.62° , and this shows that DLC system satisfies the design requirement. The increments of drag and pitching moment coefficients corresponding to full DLC flap deflection under the same conditions are not so large. 2. Asymmetric DLC Flap Deflection

Table 1 shows the test case for evaluating the effects of DLC system malfunction, and Figures 7 and 8 show the corresponding results. A maximum rolling moment of about ±0.02 appears in case 5 and 6 in which the outboard-most DLC flaps deflect asymmetrically. Since the aileron and aileron trim can produce rolling moment of about ±0.06 and ±0.027, respectively, the maximum rolling moment induced by DLC malfunction may easily be cancelled by the Do228’s original aileron trim function. The change of yawing moment due to asymmetric deflection of the DLC flaps is fairly small.

Figure 3 Wind Tunnel Model

DLC#1#2 #3 #4 #5

#6Landing Flap

3

-40 -20 0 20 40-0.4

-0.2

0

0.2

0.4∆CL

CL

-0.2

120.

211

0.42

3

48

α (deg)

δDLC (deg)

(deg)-3.62

WIND TUNNEL TEST

Figure 4 ∆CL/CL versus δδδδDLC (δδδδf = 0°)

-40 -20 0 20 40

0.02

50.

030

-0.0

05

-0.06

-0.03

0

0.03

0.06∆CD

08

α (deg)

(deg)-3.62

δDLC (deg)

WIND TUNNEL TEST

Figure 5 ∆CD versus δδδδDLC (δδδδf = 0°)

-40 -20 0 20 40

0.100.

07-0

.03

-0.16

-0.08

0

0.08

0.16∆Cm

(deg)-3.62048

α (deg)

δDLC (deg)

WIND TUNNEL TEST

Figure 6 ∆Cm versus δδδδDLC (δδδδf = 0°)

Table 1 Test Cases for DLC Malfunction CASE DLC#1 #2 #3 #4 #5 #6

1 –30 0 2 +25 0 3 0 –304 0 +255 –30 0 +256 +25 0 –30

-10 0 10 20-0.04

-0.02

0

0.02

0.04

CASE 6

24

1

5

3

∆C

∆C = 0.027(trim)

α (deg)

WIND TUNNEL TEST

Figure 7 ∆Cl versus αααα (δδδδf = 0°)

-10 0 10 20

∆Cn

CASE 6

3 2

4 1

5

-0.006

-0.004

-0.002

0

0.002

0.004

0.006

α (deg)

WIND TUNNEL TEST

Figure 8 ∆Cn versus αααα (δδδδf = 0°)

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FLIGHT TEST Test Method

In the flight tests, only aerodynamic characteristics for cases of symmetric DLC flap deflection were obtained. Indicated air speed was varied quasi-statically between 80 and 140 KIAS (40 and 70 m/s) while maintaining a trimmed flight condition at DLC flap angles of –30°, 0°, and +25°.

The judgment of the trimmed flight state was made by comparing the actual lift and drag coefficients CLact and CDact with the values corresponding to completely trimmed flight, CLss and CDss.

( )α−α−= sinAcosAqSWC xzactL

( )α+α−= sinAcosAqSWC zxactD

γ= cosqSWC ssL

γ−= cosqSWC ssD

It was judged that trimmed flight was realized when both | CLact – CLss |, and | CDact – CDss | were less than a certain value3).

Results of Flight Test Figures 9-11 show the longitudinal aerodynamic

coefficients CL, CD, and Cm estimated from the flight test results based on the following equations

actLL CC =

α+= cosCCC TactDD

eemhhmm CCC δ+δ= δδ

where Cmδe and Cmδh are derived from wind tunnel data, and CT is calculated based on an existing thrust model.

The change of lift (∆CL/CL) corresponding to full deflection of the DLC flaps (δDLC = +25°/–30°) at the design point of the DLC system (AOA α=8°, flap angle δf = 0°) extracted from the flight test data was evaluated. As shown in Fig. 12, the actual DLC system satisfies the design requirements (∆CL/CL = 0.224 at δDLC = +25° and ∆CL/CL = –0.254 at δDLC = –30°) without presetting of the DLC flaps. Figures 13 and 14 show the increments of CD and Cm due to full deflection of the DLC flaps. As in the results of the wind tunnel test, these are not so large.

α (deg)

CL

0

0.5

1.0

1.5

2.0

-5 0 5 10 15

FLIGHT TESTWIND TUNNEL

0

-30

(deg)δDLC= 25

Figure 9 CL versus αααα (δδδδf = 0°)

0

0.05

0.10

0.15

0.20CD

α (deg)-5 0 5 10 15

FLIGHT TESTWIND TUNNEL

0(deg)δDLC= 25

0

25-30

-30

Figure 10 CD versus αααα (δδδδf = 0°)

α (deg)-5 0 5 10 15

-0.4

-0.2

0

0.2

FLIGHT TESTWIND TUNNEL

Cm

0

(deg)δDLC= 25

-30

Figure 11 Cm versus αααα (δδδδf = 0°)

5

-40 -20 0 20 40

∆CL

CL

-0.4

-0.2

0

0.2

0.4

48

α (deg)

-0.2

540.

224

0.47

8

δDLC (deg)

FLIGHT TEST

Figure 12 ∆∆∆∆CL /CL versus δδδδDLC (δδδδf = 0°)

-40 -20 0 20 40-0.06

-0.03

0

0.03

0.06∆CD

48

α (deg)

0.03

60.

050

-0.0

14

FLIGHT TEST

δDLC (deg)

Figure 13 ∆∆∆∆CD versus δδδδDLC (δδδδf = 0°)

-40 -20 0 20 40

0.15

0.09

-0.0

6

-0.16

-0.08

0

0.08

0.16∆Cm

FLIGHT TEST

δDLC (deg)

48

α (deg)

Figure 14 ∆∆∆∆Cm versus δδδδDLC (δδδδf = 0°)

COMPARISON OF TEST RESULTS

The value of ∆CL/CL derived from the flight tests is larger than that obtained from the wind tunnel tests. Although there are many causes of error, such as interference of the model support and the tunnel walls in wind tunnel tests, and dynamic effects and measurement error in flight tests, in this case the cause of the discrepancy seems to be the power effect. CT was maintained at a constant value while AOA was changed in the wind tunnel tests, Consequently, CT increases as the AOA increased in the flight tests, but power effect is not reflected in the wind tunnel tests.

For the drag coefficient and the pitching moment coefficient, there is a difference between the results of the wind tunnel and flight test due to the power effect, but the tendency is very similar.

CONCLUSION

The performance of the MuPAL-α’s DLC system was evaluated by the wind tunnel and flight tests. As a result, the followings were proved: 1) The vertical acceleration generated by the DLC

flaps satisfies the design requirement. Although the wind tunnel tests indicate that the “presetting” (biasing) the DLC flaps is necessary, the flight tests show that the DLC system satisfies the design requirements without any presetting. This indicates the possibility to considerably simplify the operation of the DLC system, and should be investigated in further.

2) As the maximum rolling and yawing moments induced by asymmetric deflection of the DLC flaps can be easily cancelled by the aileron deflection, malfunction of the DLC flaps has little effect on the safety of the aircraft.

REFERENCES 1）Masui. K., Tsukano. Y., “Development of New

In-Flight Simulator MuPAL- α”, AIAA Paper 2000-4575, Aug.2000.

2）Hozumi. K., Shirai. M., “Basic Aerodynamic Characteristics of MuPAL’s DLC flaps Derived from Low-Speed Wind Tunnel Testing ”, NAL TM-756, Feb.2001. (In Japanese)

3）Victor C. Stevens., “A Technique for Determining Powered-Lift STOL Aircraft Performance at Sea Level from Flight Data Taken at Altitude ”, Society of Flight Test Engineers (SFTE), 13th Annual Symposium Proceedings, Sep.1982.