u REPORT 86

CO \- CL. o o_ ADVISORY GROUP FOR AERONAUTICAL

RESEARCH AND DEVELOPMENT

REPORT 86

AMERICAN DEVELOPMENT IN STOL AND VTOL AIRCRAFT

by

C. W. MESHIER ROYAL A TT r |5| 12PH C1957 AUGUST 1956 LISP A fi y .1

NORTH ATLANTIC TREATY ORGANIZATION PALAIS DE CHAILLOT. PARIS 16

REPORT 86

NORTH ATLANTIC TREATY ORGANIZATION

ADVISORY GROUP FOR AERONAUTICAL RESEARCH AND DEVELOPMENT

AMERICAN DEVELOPMENT IN STOL AND VTOL AIRCRAFT

by

C.W. Meshier

This Report was presented at the Ninth Meeting of the Flight Test Techniques Panel held from 27th to 31st August, 1956, in Brussels, Belgium. SUMMARY

Current military requirements for aircraft which must not only operate from small unprepared forward areas, but also must achieve flight per­ formance beyond the limitations of present day helicopters, have led to greatly expanded research activity in the VTOL and STOL fields. To-date most of this activity has been directed toward development abd evaluation of basic configurations to achieve the desired flight performance. The modern aircraft gas turbine has solved the problems of obtaining suffi­ ciently high ratios of power to engine weight to make vertical flight of fixed wing aircraft possible. Recent developments in variable incidence and highly-flapped wings promise to allow such vertical flight at con­ ventional attitudes. Less successful effort, however, has until very recently been applied to the problem of controlling such aircraft at very low airspeeds.

The problem of providing control power for hovering flight in VTOL aircraft is an obvious one. Not so obvious, perhaps, is that presented by the STOL aircraft. Such aircraft, flying on the 'back side' of the power required curve during landing approach and immediately following take-off will be operating in a regime where gusts, wind shifts, and incorrect piloting techniques can very quickly place the aircraft at an airspeed where it cannot sustain itself at landing speed power levels. If it relies on conventional aerodynamic surfaces outside the slipstream for control the pilot can easily lose control under these conditions. Therefore even STOL aircraft which are normally operated at finite airspeed will require means of control effective at zero airspeed if practical operation from short fields under adverse weather conditions is to be realized.

533.6.015.1

ii SOMMAIRE

Les avions militaires n'ont a leur disposition, dans le voisinage des lignes de combat, que des terrains restreints et non amenages. Ils doivent cependant accomplir des performances de vol qui ne sont pas a la portee des helicopteres dont on dispose a 1'heure actuelle. Les carac­ teristiques ainsi exigees de ces appareils ont donne lieu a des travaux de recherche intenses dans le domaine de 1*atterrissage et du decollage a la verticale (VTOL) et sur courte distance (STOL). Ces travaux ont surtout porte, jusqu' ici, sur la realisation et 1'evaluation des con­ figurations de base necessaires a, 1' accomplissement des performances de vol requlses. L'emploi des turbines agaz dans 1'aviation moderne aresolu le probleme qui consiste a obtenir des rapports puissance-poids du moteur suffisamment eleves pour permettre le vol vertical des appareils a voileur fixe, Les progres recemment accomplis dans le domaine des ailes a in­ cidence variable et des ailes battantes laisse entrevoir la possibility de vols verticaux avec assiette normale. Cependant, jusqu'a une date r^cente, c'est avec moins de succ^s que 1'on a aborde le probleme du contrdle des appareils du type decrit a des vitesses tris faibles.

II est eVlderament necessaire de disposer d'un moyen de controler 1'appareil VTOL pendant la phase de vol stationnaire. Cette necessite est moins evidente en ce qui concerne les appareils STOL. Un avion de ce type, volant vers le debut de la courbe reprlsentant la puissance necessaire au vol, au cours des manoeuvres precedant 1'atterrissage ou suivant imm^diatement le dlcollage, fonctionnera a un regime oil toute rafale, tout changement de direction du vent ou toute faute technique de pilotage le mettra dans des conditions de vitesse telles que la susten- tation sera insuffisante compte tenu des puissances disponibles aux vitesses d'atterrissage. Si 1'appareil depend, pour sa regulation, des surfaces aeVodynamiques classiques en dehors du souffle de 1'hllice, le pilote peut facilement, dans ces conditions, perdre le controle de celui-ci. Par consequent, meme les appareils STOL qui sont normalement actionne*s a des vitesses limitees necessiteront des moyens de contrfile efficaces a une vitesse nulle si 1'on veut operer a partir de terrains restreints dans des conditions atmospheriques defavorables.

3c5m

iii CONTENTS

Page

SUMMARY i i

LIST OF FIGURES

NOTATION vi

1. INTRODUCTION

2. CONTROL REQUIREMENTS

3. HELICOPTER CONTROL

4. POSSIBLE CONTROL DEVICES

5. CURRENT VTOL/STOL CONFIGURATIONS

6. CONCLUSION

ACKNOWLEDGMENTS

REFERENCES

FIGURES

DISTRIBUTION

iv LIST OF FIGURES

Page

Fig.1 VTOL pitch angular acceleration capability

Fig.2 VTOL roll angular acceleration capability

Pig.3 VTOL yaw angular acceleration capability 10

Pig.4 VTOL pitch control. Maximum control fuel flow as percentage of lifting power fuel flow 11

Pig.5 Some American VTOL/STOL projects 12

Pig. 6 •» " 13

Fig. 7 14

Pig. 8 15 NOTATION

T/i__v component of rotor thrust perpendicular to rotor shaft with full control input

h distance from rotor hub to helicopter centre of gravity

lyy pitching inertia of helicopter

W gross weight

b span

A aspect ratio

I moment of inertia

M control moment

H translational acceleration a angular acceleration c,K constants

G.W. gross weight

Suffixes

The suffixes i and 2 denote two similar aircraft configurations.

vi AMERICAN DEVELOPMENT IN STOL AND VTOL AIRCRAFT*

C.W. Meshier*

1. INTRODUCTION

The Development of STOL/VTOL aircraft has been primarily concerned with the aero­ dynamics of achieving vertical lift and performance. The stability and control problem that these aircraft pose has been either temporarily shelved or slighted for later consideration. This review of recent American developments in STOL/VTOL aircraft discusses this flying qualities question.

The STOL concept is not new, but until quite recently, design attempts consciously directed at STOL performance were very few. However, progress can be rapid now.

As with helicopter development, which has followed generally a pattern of various rotor and control arrangements, coupled with the application of various available power plants, it is projected that VTOL/STOL developments will follow various known schemes for the derivation of vertical thrust, coupled with the application of various power plants to provide the thrust power. Successful STOL/VTOL aircraft will probably evolve as more efficient power plants are developed. Test bed VTOL/STOL aircraft, utilizing existing engines, are feasible at present. It is considered that useful tactical aircraft can be attained only after the efficient power plants are available. Several VTOL/STOL test bed programs are now procured through U.S. Army, Air Force, and Navy sponsorship. Engine developments are underway or planned to support the STOL/VTOL potential and to improve helicopter capability.

Proven gas-turbine engines which, on the average, develop about twice as much power per pound of engine as piston engines, will be ready for STOL/VTOL aircraft. The real headache will be the job of making these aircraft easily controlled and safe throughout all flight regimes.

2. CONTROL REQUIREMENTS

As in the development of the helicopter, a pattern of various rotor or propeller and control arrangements is suggested. The spectrum of propeller-driven types includes the ducted propeller, tilting wings, deflected slipstream, and tail sitters. The jet VTOL aircraft include the tilting jet, the tail sitter type, and the deflected jet.

There are obviously a considerable number of possible solutions to the aerodynamicists' one problem of developing lift at low-speed, but the related second problem of deve­ loping control forces and moments at low speed may, in some cases, even overshadow the lift problem.

* Lieut.-Commander, United States Navy, Bureau ofAeronautics, Navy Department, Washing­ ton. D.C., U.S.A. f'Short Take-off and Landing' and "Vertical Take-off and Landing'. The control capability of the STOL/VTOL aircraft should be one of the most important considerations in its design. Every concept to achieve vertical lift with the new engines and devices now available to the aerodynamicist requires some control compromise, and every proposal must be carefully examined in order to ascertain some probability of adequate control.

The question immediately arises, 'What is adequate control?' Some requirements for control in a hover and transition must be established in order that logical design decisions can be made. As no operational VTOL aircraft are in existence, no experience factors or specifications are available for the control designs. Experimental VTOL aircraft which have flown, on the other hand, have been noted for the marginal suitability of their control systems. It did not seem practical, therefore, to examine these aircraft as adequate criteria for control requirements. Since the only aircraft with successful operational VTOL experience is the helicopter, it appears expedient to review the control capability of this type in attempting to establish criteria for the control to be required of STOL/VTOL aircraft.

3. HELICOPTER CONTROL

Accordingly, a survey was made of the control power of existing helicopters. Inter­ views with test pilots indicated that maximum available angular acceleration was a good measure of controllability in hover. This appeared to be a reasonable yardstick. Angular accelerations of an aircraft produce proportional translational accelerations at the pilot's seat. These translational accelerations are one of the important items of sensory information used by the pilot in flight. These opinions are rein­ forced by results of helicopter flight activities carried out at the Langley Aeronau­ tical Laboratory of the National Advisory Committee for Aeronautics. Reference 1 presents one of the first published indications that angular acceleration response to control inputs was indeed an important measure of helicopter handling qualities. Sub­ sequent test work by Langley personnel, and Princeton University, as yet unreported, has further borne out these initial conclusions. The importance of angular accelera­ tion in yaw is officially recognized by the Military Specification for Helicopter Flying Qualities2, which specifies minimum levels of yawing angular acceleration for various gross weight ranges.

Figure 1 presents the results of a study of pitch maximum angular acceleration capability for several existing helicopters. This maximum angular acceleration was calculated using the component of rotor thrust perpendicular to the rotor shaft with full control input (±T#max) the distance from the rotor hub to the center of gravity of the helicopter (h), and the pitching inertia of the helicopter (lyy)*

T^max amax lyy '

Also presented on this figure are angular acceleration curves for typical VTOL air­ craft using a thrust-producing device at the tail of the aircraft. Accelerations are calculated for maximum thrusts at the tail of ±5% and ±20% of the aircraft gross weight. These accelerations follow the typical inverse of the square root of the gross weight ratio variation, developed as follows. At constant wing loading, wing span varies as the square root of the weight ratio, for constant aspect ratio.

•, W2 WtA W2A

Hence M - h

0I" bl

In general, other principal dimensions of the aircraft will vary directly with wing span. Since radii of gyration for similar aircraft are nearly the same percentage of aircraft principal dimensions, aircraft inertia varies as the square of the gross weight ratio. I, = k\*1. I, 2

IL2 T Substituting the span ratio.

Using a constant percentage of thrust at the tail or wing tips, the control moment ratio is

cW b Ml = l i

M2 = cW2b2

M2 ff2 Ml (?) Wl ffi The angular acceleration ratio can now be found as follows f2_ . *t/h 'd. The translational acceleration ratio at any similar points in the structure due to this angular acceleration is then found as

X2 b2a2 / b2

Xi bi*i \bi/ V*i

^ / Kr = 1.

The result indicated that, with the calculated variation of angular acceleration, the translational acceleration remains constant. Since the pitch angular acceleration capability for the helicopters surveyed very nearly follows the curve developed for ±20% gross weight thrust at the tail, and since the scale factor of these aircraft follows the square root of gross weight ratio, the vertical accelerations of the pilot's seat due to pitch control inputs will be nearly constant. This is another indication that angular acceleration is a good parameter for control comparison.

Figure 1 also includes the maximum pitching acceleration which can be developed by changing the deflection of a 30% chord full-span flap ±10°. This deflection limit was chosen to limit the change in aircraft lift coefficient to 20% of its nominal value (produced by 50° deflection of a 30% chord flap). Data from Figure 9 of Reference 3 was used to provide data for calculating these control moments. In order to develop this calculated moment, the 30% flap was assumed to carry no average deflection. This means another flap would have to be used to provide the resultant force turning angle required for normal sustenatlon of the aircraft. If 50° average deflection is assumed, motion of a 30% flap gives very small changes in moment, and hence is even less effective as a control device.

Figure 2 shows the results of the survey of existing helicopter roll angular accelera­ tion capability. Also shown are plots of angular acceleration with ±2% and ±10% gross weight thrust at each wing tip. This plot again shows that most present day helicopters follow the characteristic inverse square root of the gross weight ratio variation.

Figure 3 presents the yaw angular acceleration survey, as well as the requirements of Reference 2. Yaw acceleration capabilities for two values of thrust of ±2% and ±6% gross weight at the tail are also shown.

4. POSSIBLE CONTROL DEVICES

Selection of a VTOL/STOL control device capable of providing the controllability defined in the previous section involves many considerations, including such items as weight, reliability, complexity, ease of operation, asymmetric power and fuel con­ sumption. Some of the devices suggested have included light weight jet engines installed at the tail and wing tips, tail rotors, and engine driven compressors. Figure 4 shows a conjectured maximum fuel flow rate to such pitch control devices in percentage of fuel flow required for main propulsive and lifting power plants. The data for this figure were calculated on the basis of 4 lb of gross weight per brake horsepower of main propulsion, using the numerical values of specific fuel consumption in Ib/h.p.hr and lb/lb thrust. This assumption gives a somewhat optimistic view of jet engine use. It is seen that fuel requirements of a jet engine or high-pressure blower may con­ sume an unacceptable amount of fuel, jeopardizing the performance of the aircraft. A tail rotor requires less fuel but possesses attendant liabilities of complexity, drag, and weight.

Figure 2 shows that aerodynamic control surfaces may provide only marginal control capability for the aircraft. In addition it has been noted by N.A.C.A. personnel4 that such surfaces are less effective in close proximity to the ground. Leading edge slat devices have similar drawbacks.

The use of large propellers or rotors with flapping and lag hinges and cyclic control does offer the possibility of adequate control, but increased hub complexity, perhaps more weight, and higher maintenance might prove undesirable.

It is seen that whatever control device is utilized, features of the aircraft will be compromised. We must be certain that the compromise is not made with control.

5. CURRENT VTOL/STOL CONFIGURATIONS

These surveys establish the control capabilities of operational helicopters. Since VTOL/STOL control requirements will be similar to helicopter requirements, it seems prudent to provide such control power to any VTOL/STOL aircraft.

However, the hover control problem Is only part of the picture. Control in tran­ sitional flight, where such aircraft, flying on the 'back side' of the power required curve during landing approach and immediately following take-off, may be even more critical. References 4 and 5 report hover and transition flight tests by N.A.C.A. personnel of a model transport vertical-take-off airplane with tilting wing and propellers. Reference 6 reports transition flight tests of the same model in a high wing configuration with jet thrust added at the tail for pitch and yaw control. Control requirements in close proximity to the ground and in transition flight were discovered to far exceed those required for hover. Motion-picture film supplements present the results of these investigations more readily than they can be described. These films are available and can be shown on request.

For a look at some of the American STOL/VTOL efforts. Figures 5, 6, 7 and 8 show both actual test bed vehicles and artists' conceptions of operational aircraft. These figures all represent military sponsorship by either study or hardware contracts. The various configurations presented exhibit varying degrees of control capability con­ sistent with their particular design compromise. Actual flight tests of experimental test beds will, of course, tailor the controllability specifications eventually re­ quired to design operational STOL/VTOL aircraft. It is hoped that some of these test beds will be flying next year.

6. CONCLUSION

All the configuration variations of STOL/VTOL aircraft covered in the previous section may never become fully developed types. A phase program of development will eliminate some, while more may be added. All the configurations mentioned, as well as others, are potentially capable of producing the main force required for lift as well as the required amount of control power. However, some of them rely on the generation of large forces in addition to the main force required for lift and may impose addi­ tional power, fuel consumption, and weight penalties on the aircraft. Relatively inexpensive experimental test bed programs should provide the ground rules necessary to design high performance operational military and commercial STOL/VTOL aircraft, but these ground rules are certain to contain control requirements - control require­ ments that may not be possible or compatible with every configuration!

ACKNOWLEDGMENTS

The author acknowledges the assistance of the N.A.C.A., for the VTOL Transport Test Model motion-picture film supplements, of numerous authors of unclassified papers on STOL/VTOL aircraft, of Mr. D.W. Robinson, Jr., Kaman Aircraft Corporation, and of the aircraft manufacturers, who have provided illustrations for this paper. REFERENCES

1. Rieder, John P. Some Effects of Varying the Damping in Pitch and Roll Whitten, James B. on the Flying Qualities of a Small Single-Rotor Heli­ copter. N.A.C.A. T.N. 2459, January 1952.

2. Military Specification Helicopter Flying Qualities, Requirements for; MIL-H- 8501, dated 5 November 1952.

3. Kuhn, R.E. An Investigation of a Wing-Propeller Configuration Draper, John W. Employing Large Chord Plain Flaps and Large Diameter Propellers for Low Speed Flight and Vertical Take-off. N.A.C.A. T.N. 3307.

4. Lovell, Powell M., Jr. Hovering-Flight Tests of a Model of a Transport Vertical- Parlett. Lysle P. Take-off Airplane with Tilting Wing and Propellers. N.A.C.A. T.N. 3630, March 1956.

5. Lovell, Powell M., Jr. Transition - Flight Tests of a Model of a Low-Wing Parlett. Lysle P. Transport Vertical-Take-off Airplane with Tilting Wing and Propellers. N.A.C.A. T.N. 3745, May 1956.

6. Lovell, Powell M., Jr. Transition - Flight Tests of a Model of a High-Wing Parlett, Lysle P. Transport Vertical-Take-off Airplane with Tilting Wing and Propellers and Jet Thrust for Yaw and Pitch Control. N.A.C.A. Proposed T.N. ANGULAR ACCIL y 3£cx ^

3 (m -MISTING mXOPTffb PRORUlSIVtZ r?07VA? miZO 'A (7.W. THRUST AT TAIL

I- "- m. THRUST AT TAIL

0* ... i £ H- £ & 10 ZO 40 &.W.- THOUSANDS OF 135.

Fig.1 VTOL pitch angular acceleration capability ANGUlAP ACCU.

S -£x/srm HuicopTtt

±\OVoGW IV/A/G TIP THfOST

Z RfiQPCtlS,y£ RO

±mZ°/o GW rV/AJ& TIP TNtfOST 0 • i i i i z i 6 3/0 20 10 GK - 1000 's OF ji&s. Pig.2 VTOL roll angular acceleration capability ANGULAR ACCEL. LO­ RAP SEC - EXISTING- HELICOPTER OS pROPucs/ye /fOTO/f (AT f/2 CAFABJi./Tr) Oh-

0.1

0.2 *2% &W TAIL THRUST

O - £. t G 8 /0 ,30 iO C W. - /OOO's or 135.

Fig.3 VTOL yaw angular acceleration capability 11

COMTQOL FUEL noiM %r /.IF1IV6- FutL fkO*J /60 m

f20

100

so ±£0 7* GH J£T JEMGIME

Co

20 y*2Q V* GV TAJL '/fOlTO/? ^-P&OPC/LS/ite /=?OTO/R' Z 1 6 3 10 ZD 40 G.W* - 1000 fc Of IBS Fig.4 VTOL pitch control. Maximum control fuel flow as percentage of lifting power fuel flow 12

TURNING EFFECTIVENESS BASIC CONSIDERATIONS

DISK LOADING-W/4-L^/FT1

llltCT ISCEIT Piuciru

Fig.5 Some American VTOL/STOL projects 13

- -si,-'

•vflttf* uoctL *-ie • ' *'

Fig.6 Some American VTOL/STOL projects 14

RESCUE

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jr

ton It AC' NwltV UXt nonuOMAM TBAMKMT AIICIAn STUDY M-231 WITH BIC FMOUCH IMi OHICI '* N*v»l IIMIKII «Ot l« U. S. ARMY

Fig.7 Some American VTOL/STOL projects VTOL D-182ADucted Propeller I TEST AIRPLANE 4-

PIASECKI HELICOPTER CORPORATION MODEL - 76

Fig.8 Some American VTOL/STOL projects DISTRIBUTION

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GREECE Greek Nat. Def. Gen. Staff GRECE B. MEO Athens.

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