Dynamic Pressure Ratio

The dynamic pressure at the horizontal stabilizer, qH, is usually less than the free – stream value because of the loss of momentum due to the air passing around the main rotor hub and fuselage. This loss, of course, is greatest for the inboard regions of the stabilizer and is less outboard. For preliminary design purposes, Figure 8.9 can be used to obtain an estimate of what the average dynamic pressure

Source: Hoak, “USAF Stability and Control Datcom,” 1960.

Source: Harris et al., “Helicopter Performance Methodology at Bell Helicopter Textron,” AHS 35th Forum, 1979.

Source: Hoerner & Borst, Fluid-Dynamic Lift, published by Mrs. Hoerner.

ratio might be for a given configuration. One set of flight test data from reference

8.4 and three sets obtained during the powered wind tunnel tests reported in references 8.5, 8.6, and 8.7 are presented. Note that the presence of the rotor wake has a significant effect on the dynamic pressure ratio, making the distribution unsymmetrical and producing some values above unity, especially behind the advancing side of the rotor disc. The effect of the rotor can be even more dramatic for those speed conditions where the horizontal stabilizer is immersed in the rotor wake. As an illustration of this effect, Figure 8.10 presents the results of wind tunnel and flight tests reported in reference 8.8. The fact that the increase in dynamic pressure can be higher than the disc loading is a result of the distortion of the wake in low-speed flight discussed in Chapter 3 under "Correlation of Flapping with Test Results.”

If a wind tunnel model of a new design is being tested, the dynamic pressure ratio can be determined indirectly. This is done by holding the model at a constant angle of attack and varying the incidence of the horizontal stabilizer (or, as is more common, cross-plotting data from several pitch runs with different incidence settings). The change in measured pitching moment can then be used to evaluate the product of dynamic pressure ratio and the slope of the lift curve for the horizontal stabilizer:

A M

A/’h

This product will be valid for the stabilizer geometry tested and can be used directly in this form in the analysis. For design studies in which alternate stabilizers are being investigated, the ratio, qH/q can be found by assuming that the value of aH found from Figure 8.6 is valid.

Downwash Angle

The downwash angle at the horizontal stabilizer is generated by both the main rotor and the fuselage—although the latter effect is usually small unless the helicopter has a wing or has widely spaced engine nacelles with a combined span greater than the span of the stabilizer. The downwash due to the rotor can be estimated from the results of downwash surveys made in a wind tunnel and reported in reference 8.9. Figure 8.11 presents the test measurements, at a tip speed ratio of 0.23, which may be taken as typical of forward flight, for several

Source: Blake & Alansky, “Stability and Controi of the YUH-6IA,” JAHS 22-1, 1977.

vertical locations and two longitudinal locations. It may be seen that the induced velocity ratio, exceeds 2—the maximum theoretical value—in some

locations. The downwash angle due to the main rotor at the horizontal stabilizer, can be estimated from the curve of Figure 8.11 that most closely matches the longitudinal and vertical position of the stabilizer with respect to the main rotor. Using a mean value of vH/vl that corresponds to the span of the stabilizer, the downwash angle is:

or

vH D. L.

H vx Ц

As an illustration, the horizontal stabilizer of the example helicopter is located at X’/R = —1.08 and Z’/R = +0.3, and the semispan/radius ratio is 0.2. From Figure

8.11, the effective value of vH/vx is 1.5. At 115 knots and design gross weight, the corresponding downwash angle is 0.06 radians or 3.5°.

V’/p — __ c

Advancing Side Retreating Side

Source: Herson & Katzoff, “Induced Velocities near a Lifting Rotor with Non-Uniform Disc Loading,” NACA TR 1319, 1957.

Figure 8.11 also shows an effect that was not recognized as significant at the time of the test but has been recognized since then. It is the higher downwash behind the advancing side than behind the retreating side for locations just behind the hub. When helicopters with high disc loading and large horizontal stabilizers are flown, it is found that they have a coupling of pitch with sideslip as the stabilizer moves either to a high downwash region behind the advancing side or to a low downwash region behind the retreating side. Discussions of this effect are contained in references 8.10 and 8.11.

Another wind tunnel test in which the downwash angle at the horizontal stabilizer was measured directly is reported in reference 8.12. In these tests, a free-

0 .01 .02 .03 .04 .05 .06 .07 .08 .09 .10

Cjlo

FIGURE 8.12 Measured Downwash at Horizontal Stabilizer

Source: Bain & Landgrebe, “Investigation of Compound Helicopter Interference Effects," USAAVLABS TR 67-44, 1967.

floating stabilizer was used as a flow vane. Figure 8.12 shows the measured downwash angle for various combinations of fuselage, wing, and rotor at constant fuselage angle of attack as the rotor thrust was varied with collective pitch. The downwash angles due to rotor with and without the wing have been converted into induced-velocity ratios on the bottom portion of the figure by using the equations:

1Cl[12]Ct/o-0

2|i2

4.0 г 3.5 3.0 2.5 Vh ^ 2.0 1’5, / / / 1.0

It may be seen that for this configuration, the rotor-induced velocity ratio is essentially the fully developed value of 2.0, which would also be read from Figure

8.11 for this stabilizer position, and that the presence of the wing has essentially no effect. Figure 8.13 shows yet another set of downwash measurements from the wind tunnel tests of the powered model of reference 8.5. Again the asymmetry of the lateral distribution is evident.

A more recent wind tunnel test produced the data on Figure 8.14 for the flow conditions at the stabilator position of the Hughes AH-64. The powered model used a rotor from a previous test that was somewhat small for the size of the fuselage. Thus the flow survey was made at two positions: one close to the rotor at low speeds, and one further back at high speed, where the influence of the fuselage and wings was more significant. Results taken from reference 8.7 for both survey locations are shown in Figure 8.14. Note that at low’ test speeds, the rotor wake increased the local dynamic pressure significantly above that of the wind tunnel.

The equation for the downwash at the stabilizer due to the fuselage with or without a wing can be written:

Since wings on helicopters are relatively small compared to the fuselage, the charts prepared by the airplane people are generally not directly applicable. Figure

8.15 (page 500) gives some test results for the downwash as measured by the floating stabilizer of the model of reference 8.12 for the fuselage alone and for three different-sized wings. These can be used as a rough guide for estimating the effect during preliminary design. Configurations with external engine nacelles, such as the example helicopter, can be assumed to have almost the same downwash characteristics as the small wing of Figure 8.15. In another wind tunnel test, this time on the Sikorsky S-76, reported in reference 8.13, the value of dtF /daf was measured as 0.15.

If a wind tunnel model with adjustable horizontal stabilizer incidence is being tested without a rotor, the fuselage-induced downwash can be determined by the following procedure: •

FIGURE 8.14 Measured Downwash Angles

Source: Logan, Prouty, & Clark, “Wind Tunnel Tests of Large and Small Scale Rotor Hubs and Pylons,” USAVRADCOM TR-80-D-21, 1981.

Source: Bain & Landgrebe, “Investigation of Compound Helicopter Aerodynamic Interference Effects,” USAAVLABS TR 67-44, 1967.

methods used for wings are directly applicable unless the configuration is such that the stabilizer can be considered to be operating in clean air. However, if values are needed before wind tunnel tests are done, the wing method is the only method readily available. The theoretical span-efficiency factor, 5, as a function of aspect ratio and taper ratio is given in Figure 8.17, which was taken from reference 8.14.

The zero lift profile drag coefficient, CDq, of the horizontal stabilizer can be estimated from Figure 6.30 of Chapter 6. For this case, the Reynolds number is based on the chord.

Horizontal Stabilizer Characteristics of the Example Helicopter The preceding methods can be used to make an estimate of the aerodynamic characteristics of a horizontal stabilizer from its physical parameters. For the example helicopter, these parameters and characteristics are shown in Table 8.2.