Document
NASA TN D-2847
Hard copy (HC)
Microfiche (M F) ' 53
STABILITY CHARACTERISTICS OF A
TANDEM-ROTOR TRANSPORT HELICOPTER
Langley Research Center
LaHgky Station, Hamppton, Vu.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASH1NGTON, 0. C. JUNE 1965 NASA TN D-2847 .
. .
STABILITY CHARACTERISTICS O F A TANDEM-ROTOR TRANSPORT HELICOPTER AS DETERMINED BY FLIGHT, TEST By J a m e s R. Kelly and Matthew M. Winston Langley Research Center Langley Station, Hampton, Va.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For s a l e by t h e C l e a r i n g h o u s e for F e d e r a l Scientific and T e c h n i c a l Information Springfield, V i r g i n i o 22151 - Price $2.00 I STABILITY CHARACTERISTICS OF A TANDEM-ROTOR TRANSPORT HELICOPTER AS DETERMINED B Y FLIGHT TEST By James R. Kelly andMatthew M. Winston Langley Research Center I Selected, unaugmented stability characteristics of a modern tandem-rotor transport helicopter were determined by a flight investigation. The angle-of- l attack instability was the predominant factor which res-dted in unacceptable l maneuver stability characteristics. Also present were speed and directional I inst8bilities.
I Based on pilots' comments, the current V/STOL specifications concerning , handling qualities appeared applicable to a helicopter of this size and con- figuration. Theoretical calculations of the pitch-and roll damping showed good agreement with flight measurements.
I I I I IUTRODUCTION The advantage of having information available on the unaugnented stability characteristics of a specific aircraft configuration is w e l l recognized. This information has several applications; it provides designers with an indication of the inherent stability characteristics of a particular configuration and, when the physical Characteristics of the aircraft are available, it supplies a source of data for comparisons with theory. Where comparisons with theory 1 accompany the data, the designer can obtain an insight into the degree of cor- relation expected from applying a particular theory to similar configurations.
Knowledge of the basic stability characteristics of a given configuration also provides preliminary information as to the degree of artificial stabilization required to provide satisfactory handling characteristics.
The helicopter configuration used in the present study has been independ- ently evaluated in two previous investigations (refs. 1 and 2 ) ; however, the emphasis in both references was placed on evaluation of the stability augmenta- tion system (SAS) installed in the aircraft. As a result, only limited SAS-off stability data were presented (ref. 2 ) . I n an attempt to fill the existing void, this report presents additional SAS-off stability characteristics and a relatively complete listing of the pertinent physical characteristics of the configuration.
The helicopter which was evaluated in references 1 and 2 was obt'ained by the Langley Research Center for use as a variable-stability aircraft. Prior to the application of the computer model simulation technique (described in ref. 3 ) to the variable-stability helicopter, a brief investigation of the basic unaug- mented stability characteristics of the helicopter was made in what were con- sidered to be possibly critical areas. The results of this investigation have been documented and are presented herein. Although it is not within the scope of this report to present a comprehensive theoretical treatment of the stability characteristics of the configuration investigated, the pitch and r o l l damping obtained from flight data have been compared with theory. Pilots' comments on the handling characteristics have also been included where appropriate in order to determine whether existing handling-qualities criteria are applicable to a helicopter of this size and configuration. It should be emphasized that the pilots' comments refer to the basic unaugmented characteristics since the dual stability augmentation system installed in the helicopter was not engaged during the present flight investigation.
TEST HELICOPTER AND INSTRUMENTATION Helicopter The test helicopter (fig. 1) is a modern twin-turbine tandem-rotor config- uration. A three-view drawing is shown in figure 2. Physical dimensions and characteristics are given in table I .
Control moments are generated in the following manner: Pitching moments are created by longitudinal displacement of the center stick to produce differ- ential collective pitch on the front and rear rotors; rolling moments, by lat- eral displacement of the center stick to introduce lateral cyclic pitch to both rotors; and yawing moments, by pedal inputs to produce differential lateral cyclic pitch of the front and rear rotors.
The aircraft is equipped with a dual stability augmentation system (SAS) which is employed to improve the basic handling characteristics. Since the SAS was not engaged during the present investigation, it is not described.
Lnstrumentation The helicopter was instrumented to record angular velocities, angular accelerations, and linear accelerations about the principal inertia axes. Con- trol positions, airspeed, altitude, and rotor rotational speed were also recorded. The sideslip-angle and angle-of-attack sensors were boom-mounted 3 ) and provided reliable information above approximately 25 knots; below (fig.
this speed the sensors were affected by the rotor downwash. Standard NASA recorders equipped with synchronized timers were employed.
RESULTS AND DISCUSSION .
Sensitivity and Damping The sensitivity (initial angular acceleration per inch of control) and angular-velocity damping-to-inertia ratios about the pitch, roll, and yaw axes were determined for the hovering-flight condition. Several step inputs were made independently for each axis and the pilots' control inputs and resultant angular velocities were simultaneously recorded. Typical response time his- tories axe shown in figures 4, 5, and 6 for the pitch, roll, and yaw axes, respectively. The pitch and roll time histories in figures 4 and 5 indicate that the step input is preceded by a slight control input in the opposite direc- tion. The initial control displacement is an intentional input referred to as a "false start" and is used to obtain a step input of longer duration and to minimize linear velocity effects. This practice yields more precise results.
Standard methods were used in arriving at the sensitivity and +,he damping- to-inertia ratios, and the results given iri tabie I1 represent the average values obtaioed from several step inputs in the pitch, roll, and yaw axes. For convenient reference, table I1 includes sensitivity and damping requirements obtained from references 4 and 5 . Table I1 a l s o includes the yaw control power requirements. (Control power is defined as the maximum angular acceleration which can be produced from a trimmed flight condition.) The damping require- ments given by references 4 and 5 are a function of helicopter inertia only and are therefore readily obtainable. On the other hand, the sensitivity require- ments are expressed in terns of an angular-displacement requirement in a given time interval following a 1-inch input from trim (control power is given in similar terms). For this report the angular-displacement requirement was con- verted to an angular-acceleration requirement by the following equation which assumes a first-order system (that is, a system containing only a mass and a damper ) :
% - - el
I where sensitivity, (moment per unit control to moment of inertia), required
Ms -
to produce the given displacement I angular-velocity damping-to-inertia ratio (negative values indicate
Mil -
stable damping)
r
given time (specified in refs. 4 and 5 ) tl given angular displacement after time tl (specified in refs. 4 and 3 ) It should be emphasized that the value specified in references 4 and 5 for the I angular-velocity damping-to-inertia ratio (Q/I) is used rather than the meas- ured value.
When inputs were made in the longitudinal direction, the pilot stated that the helicopter was "touchy," especially in forward flight. Although many fac- tors enter into this characteristic, the main contribution in the present case is believed to be the high longitudinal control sensitivity, which is indicated by the measured value presented in table 11.
The pitch-sensitivity requirements given in references 4 and 5 are minimum values, and no mention is made of a maximum allowable sensitivity. Since the helicopter appears to be approaching some sort of maximum it seems desirable that future criteria should consider a limitation on the maximum allowable pitch sensitivity. The fact that the heli- copter became more sensitive in forward flight is attributed primarily to the angle-of-attack instability, which is discussed in a subsequent section.
The pilot commented that the lateral response was somewhat high but satis- factory. Reference 4 states that the lateral control effectiveness (used in this context to mean the angular roll rate) shall be considered excessive if the maximum rate of roll per inch of stick displacement is greater than 20 degrees per second. By assuming that the aircraft response is described by a first- order system, the steady-state angular-rate capability about the roll axis, obtained by dividing the lateral control sensitivity by the damping-to-inertia ratio, is approximately 30 degrees per second. Reduction of the lateral con- trol sensitivity or an increase in damping would be necessary to reduce the steady-state rate capability to less than 20 degrees per second. It is inter- esting to note that if the roll damping-to-inertia value met the minimum visual- flight requirement of reference 4, the roll-rate capability for the existing roll sensitivity would be slightly less than 20 degrees per second. This con- dition implies that the sensitivity in itself is satisfactory and the high response is due primarily to the low damping. The indication that the sensi- tivity is satisfactory is also substantiated by the results of a recent inves- tigation (ref. 6 ) . In reference 6, for the minimum damping-to-inertia value required by AGARD in reference 5 (which is more than twice the damping-to- inertia ratio of the basic unaugmented helicopter), approximately the same sen- sitivity as that of the present helicopter was investigated and found to be satisfactory.
When inputs were made to the yaw axis, the pilot stated that even though the helicopter was relatively more powerful and had better directional control than earlier tandem-rotor helicopters, he would prefer at least twice the existing control power. From the comparison between the measured yaw control power and the yaw control power requirements of references 4 and 5 (see table 1 1 ) , it can be seen that the requirements for yaw control power are from 1 1 approximately 1 - to 2-times the measured value depending upon which require- 2 2 ment is used.
The investigation reported in reference 6 indicates that the minimum yaw control power requirement for maneuvering under visual flight con- ditions should be somewhere between the requirements of MIL-H-850U (ref. 4) and AGARD (ref. 5 ) . The pilots' comments obtained during the present investi- gation tend to support this view, Comparison of Calculated Pitch and Roll Damping With Measured Values Theoretical pitch and r o l l damping-to-inertia ratios were calculated for the helicopter having a gross weight of 15 500 pounds and a center of gravity located approximately 15 inches forward of the center l i n e between the rotors.
In the past, pitch-damping theory f o r tandem-rotor helicopters that neglected the change in induced velocity due t o the vertical velocity of the rotor disks was reported t o provide a good estimate of the pitch damping for an overlapped tandem helicopter configuration. This theory, when applied t o the present con- figuration, resulted i n a calculated pitch damping which was three times the actual value. Consequently, the pitch damping w a s calculated by the theory presented in reference 7 i n which the induced-velocity effect was included.
Although intended primarily f o r tail-rotor studies, reference 7 is a l s o appli- cable t o studies of the pitching response of a tandem-rotor helicopter. The
computed pitch damping-to-inertia r a t i o of -0.60 @- w a s fwnd t o be i n
good sgreenent with the measured value of -0.50 sJ. The damping-to-
inertia r a t i o about the r o l l axis w a s computed by the method of reference 8.
However, since the method derived i n reference 8 was for rotors having flapping hinges on the r o t o r shaft, it was necessary t o add the damping contribution due t o the offset flapping hinges of the present configuration.
The computed r o l l rad sec
damping-to-inertia r a t i o was found t o be -0.82 2, which agreed closely
raci/sec with the measured value of -0.76 Speed Stability The speed s t a b i l i t y of the helicopter was investigated for a wide range of airspeeds at six different power settings. The speed range f r m 30 t o 105 knots was covered by all six power conditions.
Data were obtained for airspeeds as l o w as 25 knots for three power conditions and as high as 145 knots f o r one power condition. These data are presented in figure 7, which shows the varia- tion of longitudinal stick position with airspeed. The figure indicates that the aircraft i s unstable with speed for a l l conditions investigated. One p i l o t stated that the instability with speed was annoying since constant retrimming was necessary t o hold a given airspeed; nevertheless, the pilot did not consider the instability dangerous because a sufficient, control margin for maneuvering existed at all trim speeds.
i Angle-of -Attack Stability
Measured.- An attempt was made t o measure the angle-of-attack s t a b i l i t y The value of the of the aircraft by the procedure described i n reference 9.
I stability derivative determined by this method is approximately 0 . 8 rad and indicates that the aircraft is unstable. It should be noted, however, that in arriving at this value a severe limitation in the application of the proce- dure of reference 9 was encountered. One requirement of this technique is that rotor speed and forward speed be varied in direct proportion. However, the speed-governing system installed on the engines of this helicopter limits the range of obtainable rotor speeds which, in turn, limits the change in forward speed allowed for the measurement. Consequently, the angle-of-attack stability could not be determined with any degree of precision.
Pull-and-hold maneuver.- The pull-and-hold maneuver was employed to deter- mine the maneuver stability characteristics of the helicopter, and to provide a Several pull-and-hold qualitative measure of the angle-of-attack instability.
maneuvers were performed at trim airspeeds ranging from approximately 40 to 85 knots. During the maneuvers, time histories of the longitudinal stick posi- tion, pitching angular velocity, and normal acceleration were recorded and are presented in figure 8 for trim airspeeds from 50 to 80 knots. The time history of normal acceleration is divergent during the maneuver for all trim airspeeds.
With a given input, the time history of pitching angular velocity does not appear to be reaching any maximum rate for any of the conditions investigated.
The pilots commented that these maneuver stability characteristics were definitely unsatisfactory. One pilot stated that these characteristics would be dangerous at higher trim airspeeds (about 120 knots and up) especially if the pilot was flying on instruments, since the aircraft could "get away" from the pilot before he could initiate corrective action. Current requirements for acceptable maneuver stability characteristics (refs. 4 and 5) state, in part, that the time history of normal acceleration shall become concave downward within 2 seconds following the start of the maneuver and remain concave down- ward until the attainment of maximum acceleration. Based on the time histories presented in figure 8, the characteristics shown do not satisfy the current requirements. Since the pilots considered these maneuver stability character- istics unacceptable, the current requirements appear to be adequate in this area.
Longitudinal Trim Change With Power The longitudinal trim change with power was measured at a forward speed of 20 knots. The results of this measurement are given in figure 9 where longi- tudinal stick position is plotted against vertical velocity; the vertical veloc- ity is used as a measure of power.
The pilot commented that the magnitude of the trim change (approximately 2.3 inches) was acceptable under the conditions investigated. The direction of the trim change, however, would definitely be undesirable for accomplishing certain tasks since the helicopter pitched up with a decrease in power and pitched down with an increase in power in the low-rate-of-descent region
(a ft/se&). This feature was found t o be somewhat annoying during a low-speed
1owLangle instrument landing system approach. For example, when the pilot was above the glide slope while holding a desired airspeed and made a correction by decreasing power, the aircraft pitched up due t o the t r i m change effect. This pitch-up caused the helicopter t o f l a r e and thus resulted i n i t i a l l y in an even greater deviation from the glide slope and a loss of airspeed. The f l a r e w a s eventually followed by the desired increase i n rate of descent necessary t o acquire the glide slope. Present requirements only specify a maximum allowable trim.change independent of direction.
Based on the pilot comments noted herein, it appears that future c r i t e r i a should possibly give consideration t o the direc- tion as w e l l as the magnitude of a trim change.
D i h e d r a l Effect The dihedral effect at various flight conditions is presented i n fig- ure 10. The variation of lateral control displacement with sideslip angle is approximately linear and indicates that the helicopter is stable at the various flight conditions of t h i s investigation. A t any trim airspeed the dihedral effect i s essentially independent of power condition; however, there is a def- i n i t e increase i n dihedral effect w i t h an increase i n airspeed.
Directional Stability The results of the directional-stability measurements are given in fig- ure 1 1 where the variation of pedal position with sideslip angle i s presented for several flight conditions. Because of structural limitations on the air- craft the maximum sideslip angles investigated were +26O. The results indicate that the helicopter is unstable for l e f t sideslip angles and f o r small right sideslip angles for all the flight conditions. The helicopter may be either stable or unstable for large right sideslip angles, depending upon the power condition and trim airspeed considered. In general, there is a tendency for the helicopter t o become more unstable directionally as the speed i s increased.
The p i l o t commented that for visual flight operations the directional s t a b i l i t y was unsatisfactory but not necessarily dangerous f o r the range of sideslip angles covered. However, under instrument conditions, the pilot con- sidered that the directional s t a b i l i t y characteristics would be unacceptable.
This comment is substantiated by the investigation of reference 10 i n that the directional s t a b i l i t y and directional angular velocity damping are i n the unacceptable region of figure 3 i n reference 10. The p i l o t noted that since the helicopter constantly diverged from a desired heading, it would be very tiresome t o f l y i n the navigation mode even under v i s u a l flight conditions.
CONCLUDING FEMARKS An investigation of the unaugmented stability characteristics of a modern tandem-rotor helicopter has been conducted. The predominant characteristic of the configuration is the angle-of-attack instability which results i n unacceptable maneuver stability characteristics.
Also present are speed and directional instabilities, but these instabilities are of a much lesser degree.
Within the flight conditions covered by this investigation, current V/STOL specifications concerning handling qualities appear applicable to a helicopter of this size and configuration. It is shown that a good estimate of the pitch and roll damping is provided by existing theory.
Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., March 16, 1965.
REFERENCES 1. Taylor, F. W.; and Wirt, W. 0.: Navy Evaluation of the Model YHC-IA Heli- copter. Rept. no. 1 (PTR AD 3 0 9 7 ) , Flight Test Div., U.S. Naval Air Test Center (Patuxent River, M d . ) , May 5, 1960.
2 . Crawford, Charles C . ; and Hodgson, Walter J . : YHC-IA Flight Evaluation.
AFFTC-TR-61-1, U.S. Air Force, Feb. 1 9 6 1 .
3 . Garren, John F . , Jr.; and Kelly, James R.: Description of an Analog Com- puter Approach to V/STOL Simulation Ehploying a Variable-Stability Heli- copter. NASA 'I" D-1970, 1 9 6 4 .
4 . Anon.: Helicopter Flying and Ground Handling Qualities; General Require- ments for. Mil. Specification MIL-H-85OlA, Sept. 7, 1961; Amendment 1, Apr. 3, 1962.
5 . Anon.: Recommendations for V/STOL Handling Qualities. AGARD Rept. 408, oct. 1962.
6. Garren, John F., Jr.; Kelly, James R.; and Reeder, John P.: A Visual Flight Investigation of Hovering and Low-Speed VTOL Control Requirements.
NASA TN D-2788, 1-96?.
7. her, Kenneth B . ; and Gessow, Alfred: Charts for Estimating Tail-Rotor Contribution to Helicopter Directional Stability and Control in Low-Speed Flight. NACA Rept. 1216, 1955. (Supersedes NACA TN 3156.)
8 . her, Kenneth B . : Theory of Helicopter Damping in Pitch or Roll and a Com- parison With Flight Measurements. NACA TN 2136, 1950.
9 . her, Kenneth B.: Method for Studying Helicopter Longitudinal Maneuver Stability. NACA Rept. 1200, 1954. (Supersedes NACA TN 3022.)
1 0 . Garren, John F., Jr.; Kelly, James R . ; and Reeder, John P . : Effects of Gross Changes in Static Directional Stability on V/STOL Handling Charac- teristics Based on a Flight Investigation. NASA TN D-2477, 1964.
TABLE I.- PHPSICAZ, CHARnCTERIsTICS OF TEST EUICOPER O v e r a l l dimensions (blade turning): Length, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81.66 Width. ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48.33 Height. ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18.00 Rotor characteristics: Distance between rotors, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33.33 lhrher of b M e s per rotor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
m e airfoil section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HACA 0012 BLadechord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5 madetwist, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.05 maim, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24.17 map hinge offset, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0 . 9 3 s o l i a l t y . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.0593 Sweptdiskaree, f t 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3670 Pmjecteddiskarea, f t 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33u0 liowal rotational speed, rpm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 268 m o m t i p speed, ft/sec . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 676 Percent overlap, 100(1- D i s ~ ~ r b e ~ ~ ~ e ~ ~ O r S ) . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Rotor shaft forward tilt (with respect t o fuselage reference line): Front rotor, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.5 Rear rotor, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.0 Rotor height (above fuselage reference line): Front rotor, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -3.3 Rear rotor, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 . 3 4.24 mmal a s k loading, lb/ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
BLadeweight, l b . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -1% Inclination of principal inertia axis relative t o fuselage reference line, deg . . . . . . . . . 4 . 5 (nose down) Power-plant rating (two turbine engines), hp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1050 Xionnal parer loading, lb/hp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.15 Maximum take-off weight: lformal.,lb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15500 Overload, l b . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16600 operating gross Wight, u , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . - 1 3 000 Fuelcapscity, gal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 0 Most forward center-of-gravity location (ahead of midpoint between rotors), ft . . . . . . . . . . . . . .
2-35 Most aft center-of-gravity location (behind midpoint between rotors), ft 0.83 . . . . . . . . . . . . . . . . .
Moments of inertia: Piteh, Slug-ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75000 R o l l , slug-ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9200 Yaw, slug-ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71000 Control travel: Longitudinal stick, in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25.5 Lateral stick, in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . k3.6 Pedal,in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . k2.3 Collective pitch lever, in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.8 made travel: (at 0.75 radius), deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 t o 17 Collective pitch Longitudinal differential collective pitch, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23 Longitudinal cyclic trim: Front rotor (fixed), deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -2 Rear rotor (fixed for test, normally variable), deg . . . . . . . . . . . . . . . . . . . . . . . . . . 0.5 Lateral cyclic pitch: Front rotor, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . k6.15 R e a r rotor, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.4.60 Directional differential l a t e r a l cyclic: Front rotor, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . k7.13 Rear rotor, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . k7.13 TABLE 11.- HOVERING CONTROL CHARACTERISTICS (MEASURED AND SPECDIED BY CRITERIA) AGARD Rept. 408 MIL-H-850I-A (ref. 4) Control characteristic Measured (ref. 5 ) Instrument visual Pitch axis rad/sec 2 0.12 0.12 . . - 0 - 3 1 Sensitivity, in.
-0.28
am ping, rad/sec2 . . . . . -0.50 -0.52 -0.52
Inertia rad/sec Roll axis
-I
rad/sec2 . . . 0.40 Sensitivity, in.
Damping, radFec2
-1.63 o'22 T l -1.63 -1.17 . . . . . -0.76
Inertia rad sec Yaw axis rad/sec2
. . . Sensitivity, 0.11 0.11 0.20
in.
Control power, rad/sec* . . 0.25 0.61
0.33 Damping, rad/sec2 . . . , .
=O
-0 95 -0.95 1 -0 95
Inertia rad/sec
T (D (D L
a Nose up .1 .rl 0 0 .1 .2 -P .rl PI Nose down .3
2 R e a r w a r d 1
.rl Forward 1
3 0 .a 1 . 6 2 . 4 3 . 2 4 . 0
T i m e , see Figure 4.- Typical time history of a longitudinal step input (hovering).
co f f cu .
m
i
f cu \D rl co m cu rl 0 rl .
L
-I 0 rl rl 0 rl o?
c, 1 6 Full aft Reference stable slope -.
\ I M Full fonard Full aft Full forward Full aft 2 0 Full forward 0 I O 20 30 40 50 60 70 80 90 100 1 1 0 120 Airspeed, knots (a) rim airspeed, 45 bots.
Figure 7.- Variation of longitudinal stick position with airspeed.
Full forward FUII a i i Full f o m r d FUH a n Full forward 0 I O 20 30 40 50 60 70 80 90 100 1 1 0 120 130 140 150 Airspeed, knots (b) TrFm airspeed, 80 knots.
Figure 7.- Concluded.
.G I I I I I I I I I I
Nose down . I I
TI PI Rearward 1 . 0
r
. 3 bo c 1 . 0 Forward 1.5 0 .8 1.6 2.4 4.0 Time, sec (a) Trim airspeed, 50 knots.
Figure 8. - Pull-and-hold maneuver.
h
2 Nose up .2 -
V rl
* : -
.1 [o
3 2 _-
% .A 0 _--
4 4
5 3 .6, I I I I 1 I I I I I I 1 $Nose dom.1 ! & Rearward 1.0 d 1 I I I I I I I I 1 Forward 1 . 5 .8 1 . 6 2.4 3.2 4.0 4.8 Time, sec (b) T r i m airspeed, 60 knots.
Figure 8.- Continued.
.rl iJ c d k - aJ 1;2 A i J aJ .rl O d 1 . 0 -
; $
E
I I I I I I I I t I t 1 . 1
5 Nose down . 1
* -rl PI Rearward 1 . 0 Forward 1.5 0 .8 1 . 6 2.4 3.2 4.0 4.8 Time, sec ( c ) rim airspeed, 70 h o t s .
Figure 8.- Continued.
k a ,
-
1.2 2 3 : E l /
~-
;rl & 1 . 0 E k ? I I I I 1 I I I I I I I .8
-
.1 / c- 1 I 1 I I I 1 I I I 1 I 4 Nose down .1 %
-
Rearward 1 . 0
-
. 5 - .5 - 1.0 c L I I L I I I 1 I I 1 I Forward 1 . 5 0 .a 1.6 2.4 3.2 4.0 4.8 Time, sec ( d ) ~ i m airspeed, 80 knots.
Figure 8.- Concluded.
z 3 s
I Full right 4 0 ~ u i i i e f t l I I I I I I I I I I I I I I I I I I I I I I I I I I I I -28 -24 -20 -16 -12 -8 -4 0 4 8 12 1 6 20 24 28 Left Sideslip angle, B, deg Right (a) Airspeed, 45 knots.
Figure 10.- Variation of l a t e r a l s t i c k position with sideslip angle.
Full right 800-ftlrnin rate of descent I Full lefl -28 -24 -20 -16 -12 -8 -4 0 4 8 1 2 I6 20 24 28 Left Sideslip angle, fi, deg Right (b) Airspeed, 60 knots.
Figure 10.- Continued.
L
-28 +I -20 -16 -12 -a -4 o 4 a 12 16 20 24 28
Left Sideslip angle, B, deg Right ( c ) Airspeed, 80 knots.
Figure 1 0 . - Concluded.
. Full rqM
H10-ft/min rate of climb Full left Full right " 2 0 - - IW0-ft/min rate of descent
Full left , , , I , 1 1 I I I I I I
-28 -24 -20 -16 -12 -8 -4 0 4 8 1 2 1 6 20 24 28 Left Sideslip angle, 0. deg Right (a) Airspeed, 45 knots.
Figure ll.- Variation of pedal position with sideslip angle.
.
Full right U M - 500-ftfmin rate of climb -
F u l l l e f t , , I I I I I I I I I I I , . I I I I
Full right 800-fVmin rate of descent -28 -24 -20 -16 -12 -a -4 o 4 a 12 16 20 24 2a Left Sideslip angle, 8, deg Right (b) Airspeed, 60 knots.
Figure 11.- Continued.
Full I Full ( c ) Airspeed, 80 knots.
Figure ll.- Continued.
.
Full right - f m - c C I 0 .- - 5 9 B+
-
- 0
&
L Full left -28 -24 -20 -16 -12 -8 -4 0 4 8 1 2 1 6 20 24 28 Left Sideslip angle, B, deg Right (d) Airspeed, 60 knots.
Figure 11.- Concluded.
NASA-Langley, 1965 L-4415