Rotorcraft Aerodynamics Group

Investigation of Aerodynamic Interactions Between a Rotor and a T-Tail Empennage

Erwin Moedersheim and Dr. J. Gordon Leishman

This is the HTML version of a paper presented at:
The American Helicopter Society International Specialists' Meeting, Stratford, Connecticut, October 11-13, 1995.

Rotor and Tail Instrumentation

Instrumentation was provided on the rotor and the empennage. Instrumentation was not provided on the body in these particular tests, since it had been extensively instrumented in a previous series of tests (Ref. 18-20,). The rotor balance permitted the measurement of three mutually perpendicular time-averaged force components (thrust, drag, and side-force), along with the corresponding moments. Rotor power was measured using a torque disk coupled to the rotor shaft. Hall-effect sensors were located at the blade hinges to monitor flap and lead/lag displacements. Details of other rotor instrumentation are given in Ref. 21.

Time-averaged pressure measurements were obtained from thirty-two static pressure taps, which were located along rows at the leading-edge (x/c = 0.04) and trailing-edge (x/c = 0.81) on both the upper and lower surface of the stabilizer, as given in Table 2. The leading-edge pressure taps were positioned to detect the leading-edge pressure peak, and to give an indication of the local lift coefficient at which the tail section was operating. The trailing-edge pressure taps helped provide information on whether the flow was separated over the trailing-edge.

Pressure tap No. x/c 2 y/b
Top Bottom
1 17 0.04 0.80
2 18 0.04 0.60
3 19 0.04 0.40
4 20 0.04 0.20
5 21 0.04 -0.20
6 22 0.04 -0.40
7 23 0.04 -0.60
8 24 0.04 -0.80
9 25 0.81 0.80
10 26 0.81 0.60
11 27 0.81 0.40
12 28 0.81 0.20
13 29 0.81 -0.20
14 30 0.81 -0.40
15 31 0.81 -0.60
16 32 0.81 -0.80

Table 2: Locations of static pressure taps on the horizontal stabilizer

The time-averaged pressures were measured using a digital multi-channel pressure system. The pressure tubes were connected to a thermally insulated multiplexer module mounted inside the fuselage. This module contained miniature pressure transducers, analog multiplexers, and analog-to-digital (A/D) converters. A miniature pneumatic valving system inside the module permitted rapid on-line recalibration of the pressure sensors with the test in progress. Frequent recalibration was essential to maintain a measurement accuracy of ± 7.5 Pa (0.0011 lb/in² ) over the tunnel run times.

Time-dependent pressures were measured using pressure transducers located at sixteen locations, and grouped at two spanwise stations - see Table 3. However, due to physical constraints, it was impractical to co-locate the pressure transducers at the same chordwise or spanwise locations as the static pressure taps. Based on step response tests, the resonant frequency of the sensors was approximately 12 kHz, which was well in excess of the frequencies measured.

Transducer No. x/c 2 y/b
Top Bottom
1 9 0.075 0.50
2 10 0.263 0.50
3 11 0.494 0.50
4 12 0.725 0.50
5 13 0.075 -0.50
6 14 0.263 -0.50
7 15 0.494 -0.50
8 16 0.725 -0.50

Table 3: Locations of pressure transducers on the horizontal stabilizer

Data Acquisition System

 Data acquisition was controlled by several computer systems. Time-averaged pressure data were acquired from the pressure measuring system over a GPIB interface. Quasi-steady data, such as rotor balance loads, were logged directly by a mini-computer through a 16-channel multiplexer and a 14-bit A/D converter. Unsteady pressure data were logged using a high-speed multi-channel 12-bit A/D converter system. This data acquisition system was interfaced with a separate computer over a GPIB interface. Triggering for all the unsteady measurements was synchronized to 1-per-revolution and 256-per-revolution pulses initiated from an encoder on the rotor shaft.

The time-averaged pressure measurements were made by averaging 256 samples at each location over an interval corresponding to about 200 rotor revolutions. Time-histories of the pressure transducer responses were logged continuously over up to 20 rotor revolutions at a sampling resolution of 256 data frames per channel per revolution, i.e., an azimuth resolution of 1.4 deg. All unsteady time-history data presented in this paper are time-averaged, i.e., the data were ensemble averaged over ten or more rotor revolutions. The estimated error in C_p' was ± 0.01. The maximum likely estimated error in C_p^u was about ± 0.1.

Test Conditions

The experiments were performed in the University of Maryland's Glenn L. Martin wind tunnel, which has a 2.36 × 3.35 m (8 × 11 ft) working section. The rotor was tested at a rotational speed of 2100 rpm (35 Hz), which corresponded to a nominal tip Mach number of 0.52 in hover. Collective, lateral cyclic and longitudinal cyclic blade pitch were set remotely by means of swashplate actuators. The forward flight trim procedure was performed by minimizing the 1-per-revolution blade flapping response relative to the shaft, which resulted in a rotor tip-path-plane (TPP) was perpendicular to the rotor shaft axis.

Over seventy test conditions comprising variations in rotor thrust, advance ratio, shaft angle, and tail position were examined; the range of test parameters being summarized in Table 4. Comparative studies were conducted at a constant rotor thrust for different advance ratios and shaft tilt angles. Only selected results can be shown in this article, with most data being for a blade loading coefficient of 0.075.

Parameter Test values
Advance ratio, µ 0.05 to 0.30
Rotor shaft angle, alpha_s -6° to +4°
Rotor rotational speed 2100 RPM (35 Hz)
Rotor hover tip Mach number 0.52
Blade loading, BL 0.075 to 0.085

Table 4: Range of Test Parameters

Flowfield Survey

Flowfield survey data, which were origionally obtained by Leishman and Bi (Ref. 17, 21), were used as a reference for the current tests. These data were obtained during a test with the same rotor, both with and without the body. Miniature seven-hole probes measured the time-averaged total pressure and three components of velocity at three advance ratios ( µ = 0.075, 0.10, and 0.20) and at a blade loading of BL = 0.075. The probes were mounted on a traversing rig, and were moved in three horizontal planes located at z_h/R = -0.14, -0.29, and -0.45. While these planes were somewhat below the plane of the horizontal stabilizer used in the present test, the data provided considerable information about the stabilizer environment. The measurement grid contained 896 measurement points (28 × 32 grid) in any one plane, mostly equispaced at 7.62 cm (3.0 in).

Flow Visualization

The wide-field shadowgraph technique was used to visualize the locations of the rotor tip vortices relative to the lifting surface. The basic components of the shadowgraph system are a point light source strobe, a retroreflective screen, and a video or still camera. The tip vortices created by the rotor cause small changes in the flow density and index of refraction. Therefore, the light rays from the strobe are refracted as they pass through these vortices, causing (magnified) shadows to be cast on the screen. By examining the wake for successive rotor azimuth positions, and by using a grid system on the screen, it was possible to quantify the locations of the leading- and trailing-edges of the rotor wake boundary relative to the rotor and the empennage. Further details of the wide-field shadowgraph technique as applied in this experiment can be found in Ref. 22.

Continue with the Results and Discussion...

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Erwin Moedersheim
Dr. J. Gordon Leishman / leishman@eng.umd.edu