1、NACA-TR-605NATIC_L TECHNICALIN_RMATION SERVICEU.S. DEPARTMENT OF COMME_PRIMFIF.k,D, VA. 22151Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REPORT No.
2、605p PRESUME AND ANALYSIS OF N. A. C. A. LATERALCONTROL RESEARCHBy FRED E. WEICK and ROBERT T. JONESLangley Memorial Aeronautical Laboratory1 g22-37- lProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NATIONAL ADVISORY COMMI_TEE FOR AERONAUTICSHEADQUAR
3、TERS. NAVY BU|LDING. WASHINGTON. D. C.LABORATORIES, LANGLEY FIELD, VA.Created by act of Congress approved March 3, 1915, for the supervision and direction of the scientificstudy of the problems of flight (U. S. Code, Title 50, Sec. 151). Its membership was increased to 15 byact approved March 2, 192
4、9. The members are appointed by the President, and serve as such withoutcompensation.JOSEPH S. AMES, Ph.D., Chai_man,Baltimore, Md.DAVID W. TAYLOR, D. Eng., Vice Chai_man,Washington, D. C.WILLIS RAY GREC_, Sc. D., Chairman, Ezecutive Committee,Chief, United States Weather Bureau.CHARLES G. ABBOT, So
5、. D.,Secretary, Smithsonian Institution.LYMAN J. Bm_s, Ph. D.,Director, National Bureau of Standards.ARTHUR B. COOK, Rear Admiral, United States Navy,Chief, Bureau of Aeronautics, Navy Department.FS_D D. FAC_, JL, J. D.,Director of Air Commerce, Department of Commerce.HARRY F. GUGGRNHEIM, M. A.,Port
6、 Washington, Long Island, N. t“.SYDNEY M. KRAUS, Captain, United States Navy,Bureau of Aeronautics, Navy Department.CHARLES A. LINDBERGH, LL.D.,New York City.WILLIAM P. MACCRACKEN, J. D.,Washington, D. C.AU6USTINE W. ROSINS, Brigadier General, United StatesArmy,Chief Materiel Division, Air Corps, Wr
7、ight Field, Day-ton, Ohio.EDw_ P. WARN_g, M. S.,Greenwich, Conn.OSCAR WESTOVER, Major General, United States Army,Chief of Air Corps, War Department.ORVrLLE WRIGHT, Sc. D.,Dayton, Ohio.GEORGE W. LEWIS, Director o/Aeronautical ResearchJOHN F. VICTORY, SecretaryHENRY J. E. REID, Engineer in Charge, La
8、ngley Memorial Aeronautioa_ Labo_ato_y, Langley Field, Va.JOHN J. IDE, Technical Assistant in Europe, Paris, FranceTECHNICAL COMMITTEESAERODYNZJmCS AIRCS._t_r STRUCTURESPOWER FLANTI$ FOR AIRCRAI_ AIRC_ ACCIDENTSAIRCRAFT MATERIALS INVENTIONS AND DESIGNSCoordination of Re_earch Needs of Milita_y and C
9、ivil A viationPzeparation of Research ProgramsAllocation of ProblemsPrevention of DuplicationConsideration of InventionsLANGLEY MEMORIAL AERONAUTICAL LABORATORY OFFICE OF AERONAUTICAL INTELLIGENCELANGLEY FIELD, VA_ WASHINGTON, D. C.Unified conduct, for all agencies, of Collection, classification, co
10、mpilation,scientific research on the fundamental and dissemination of scientific and tech-problems of flight, nicai information on aeronautics./Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REPORT No. 605RESUME AND ANALYSIS OF N. A. C. A. LATERAL C
11、ONTROL RESEARCHBy F_ED E. W_ICK and ROBERT T. JON_.SSUMMARYAn analysis of the principal results o/recent N. A. C. A.lateral control research is made by utilizing the experienceand progress gained during the course of the investigation.Two things are considered o/ primary importance injudging the eff
12、ectiveness of different control devices: The(calculated) banking and yawing motion of a typical smallairplane caused by a deflection o/ the control, and the stickforce required to produce this deflection. The report in-cludes a table in which a number _ different lateral controldew,ices are compared
13、 on these bases.Experience gained while testing various devices inflight with a Fairchild 22 airplane indicated that, follow-ing a sudden deflection o/the control at low speed, anangle of bank of 16 in I second represented a satisyactoryminimum degree of effectiveness/or this size o/airplane.Some de
14、vices capable o/ giving this degree of control were,however, considered to be not entirely satisfactory on ac-count of sluggishness in starting the motion. Deviceslocated near the trailing edge of the wings had no detectablesluggishness. Lateral control forces considered desirableby the test pilots
15、varied from 2 to 8 pounds; 15 pounds wasconsidered excessive.Test flights demonstrated that satisfactory lateral controlat high angles of attack depends as much on the retention oJstability as on aileron effectiveness.The aerodynamic characteristics of plain sealed aileronscould be accurately predic
16、ted by a modification of theaerodynamic theory utilizing the results of experimentswith sealed flaps. Straight narrow-chord sealed aileronscovering 60 to 80 percent of the semispan represented aboutthe most e_icient arrangement of plain unbalanced aileronsfrom considerations of operating force. The
17、stick force ofplain ailerons can be effectively reduced by the use of adifferential linkage in conjunction with a small fixed tabarranged to press the ailerons upward.INTRODUCTIONIn 1931 the Committee started a systematic wind-tunnel investigation of lateral control with specialreference to the impr
18、ovement of control at low airspeeds and at high angles of attack. Many differentailerons and other lateral control devices have beensubjected to the same systematic investigation in the7- by 10-foot wind tunnel. (See refers,me 1.) Thedevices that seemed most promising were tested inflight (reference
19、s 2 and 3). In many cases, however,devices that produced what seemed to be satisfactoryrolling moments and favorable yawing moments didnot give satisfactory control.An analytical study of control effectiveness wastherefore made (reference 4) taking into account anumber of secondary factors, includin
20、g the yawingmoments produced by the controls, the effect of thecontrols on the damping in rolling, the lateral-stabilityderivatives of the airplane, the moments of inertia, am!the time required for the control moments to becomeestablished after the deflection of the surfaces. Thecomputations consist
21、ed of step-by-step solutions of theequations of rolling and yawing motion for the condi-tions following a deflection of the controls. The resultsof these computations based on aerodynamic data ob-tained from wind-tunnel tests of wings incorporatingvarious devices agreed satisfactorily with the resul
22、tsmeasured in flight for widely different forms of control,such as ailerons and spoilers.The study of conditions above the stall indicatedthat satisfactory control could not be expected withoutsome provision to maintain the damping in rolling andthat a dangerous type of instability would arise if th
23、edamping were insufficient. Since damping in rollingdepends on an increase in the lift of the airfoil withincreasing angle of attack, it follows that, in order toobtain satisfactory lateral control, the outer or tip por-tions of the wing, which govern the rolling moments,must remain unstalled. If da
24、mping in rolling is re-tained, it is practically insured that control momentswill be retained as well.The progress of the investigation has thus led to amore accurate interpretation of the results of the wind-tunnel tests. In the present paper the experiencegained during the course of the investigat
25、ion is madethe basis of a revised method of comparison of lateralcontrol devices. Wind-tunnel measurements of controland stability factors (reference 1) are utilized in com-putations to show the banking and yawing motionsthat would be produced by the controls acting on asmall typical airplane. These
26、 computations follow themethod of analysis given in reference 4. In section I ofthe report the new basis of comparison is explained and1Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REPORTNO.605-NATIONALADVISORYCOMMITTEEFORAERONAUTICSa number of th
27、e devices that were tested in reference 1are analyzed and compared. The principal items ofcomparison are collected into a table. Section IIpresents an analysis of the rolling, yawing, and hingemoments of plain flap-type ailerons and deals with theapplication of these data in the design of controlsys
28、tems.I. COMPARISON OF LATERAL CONTROLDEVICESREVISED BASIS OF COMPARISONAIRPLANE USED IN COMPARISONThe procedure adopted in the lateral control investi-gation has comprised a wind-tunnel test program fol-lowed by flight tests of the different devices on theFairchild 22 airplane. Not all of the device
29、s testedin reference 1 have been tried in flight, however, andthe present reporL may be considered an analyticalextension of the flight-test procedure that was appliedto some of the devices. The procedure employed totest lateral controls in flight is simulated by means ofcomputation. Thus, the compa
30、rative criterions usedherein are based on application of tile devices to a hypo-thetical Fairchild 22 type of airplane, which is the typeused in the flight tests.The Fairchild 22 airplane was necessarily somewhatmodified for each different flight test and wings of differ-ent moment of inertia, plan
31、form, and section wereused in some cases. The wing of the hypothetical air-plane assumed in the computations represents an aver-age of the tested wings. Furthermore, since the char-acteristic ratios of dimensions (tail length, tail area,radii of gyration about various axes, etc.) used agreevery clos
32、ely with statistical averages of these quanti-ties, the assumed airplane may be considered to embodyaverage stability characteristics. The principal charac-teristics of the assumed airplane are as follows:Weight, W . 1,600 lb.Wing span, b 32 ft.Wing area, 8 . 171 sq. ft.Wing loading, W/,._ 9.4 lb. p
33、er sq. ft.Area of fin and rudder . 10.8 sq. ft.Tail length . 14.6 ft.Ix . 1,216 slug-ft.*lz . 1,700 slug-ft. 2ROLLING ACTIONIt is recognized that different types of airplanes re-quire different amounts of control. At the start ofthe wind-ttmnel investigation of lateral control devices(reference 1) a
34、 rolling _.iterion (RC=CJCL) represent-ing a conservative lower limit of rolling control for alltypes was assumed. The assumed satisfactory valueof the rolling criterion was 0.075, which corresponds toa lateral movement of the center of pressure of 7.5percent of the wing span. Recent experience indi
35、catesthat this value is likely to be ample for any conditionof flight that might be encountered and is therefore adesirable value to attain. Where a compromise mustbe made between the rolling moment and some othercharacteristic of the control system, particularly thecontrol force, a decidedly lower
36、value of the rollingcriterion may be used. It appears that a value pos-sibly as low as half the original one may be foundreasonably satisfactory for practically all conditions offlight with nonacrobatic airplanes.The criterion of rolling contiol used in the presentanalysis is the angle of bank attai
37、ned in 1 second fol-lowing a sudden deflection of the control. This criterionshows the actual amount of motion produced anddepends on both the acceleration at the start and thefinal rate of roll. It includes the effect of yawingmoment given by the control as wel_ as the stabilitycharacteristics and
38、moments of inertia of the airplane.The values of the criterion are found by computationand as such are applicable only to the particular typeof airplane (F-22) that has been assumed.Experience gained in flight tests of the Fairchild 22airplane with various lateral control devices indicateda minimum
39、satisfactory amount of rolling control car-responding to about 15 of bank in 1 second. (Seefig. 1.) Ailerons capable of giving this amount of bank( 161.97 foot-pounds (16)The control force is equal to twice the total work di-vided by the linear travel of the end of the stick, or. 3 94Stick Iorce=_=5
40、.4 pounds (17)The stick force at the partial deflection required for_= 15 is$15 7.42.3IX _2-_-_.5o= 2.31Xi-1.2o= 3.6pounds (18)These simplerelationsapply,ofcourse,only tolinearvariationofthe hinge moment and tonondifferentialgearing.Differentiallinkages.-Itappearsthata differentiallinkagecan,when pr
41、operlydesigned,beaveryeffectivemeans of reducing the operatingforceof flap-typeailerons(reference11). The reductionof operatingforceisaccomplishedby takingadvantage ofthe up-floatingtendency of the ailerons.With differentiallinkagetheaileronsonoppositetipsofthewing begintomove at differentratesimmed
42、iately afterthey aredeflectedfrom neutral,the downgoing aileronmovingmore slowlythantheupgoingone. The upgoingaileronthushas thegreatermechanicaladvantageatthecon-trol-stickconnection. Itisevidentthat the reducedA.LATERAL CONTROL RESEARCH 21upward pressure of the upgoing aileron is partly com-pensat
43、ed by its increased mechanical advantage andthat the increased upward pressure on the downgoingaileron is also partly compensated by its reducedmechanical advantage. At a certain deflection thedowngoing aileron reaches dead center and, regardlessof its aerodynamic pressure, cannot contribute to thes
44、tick force; if the upgoing aileron is then at the floatingangle (i. e., angle of zero hinge moment), the stick forcewill be zero.Ordinary ailerons show nearly straight-line hinge-/dr-0.0085)and in thiscasethemoment curves-_-=balancingeffectofa givendifferentiallinkagedependsonly on the upfloatingang
45、le. A formulafora differ-entialmotion that giveszero operatingforceover arange of deflectionsmay be obtainedby writingtheexpressionforthework ofdeflectionoftheaileronsandequatingittozeroateverypoint._.,= 4 (_/+ 6,)2- 2_.,2- 6.j, (19)where $, and _ are the upward and downward deflec-tions of the aile
46、rons and $,/is the floating angle meas-ured upward from the neutral position. A practicallimitation of this formula is reached when d$_/d_,approaches -1, for then both ailerons begin to movein the same direction and at the same rate.It should be appreciated that a differential designedin accordance
47、with equation (19) will give completebalance at the specified floating angle. It is, however,considered desirable not to eliminate completely thecontrol force at any flight condition, as the pilots feelof the control would be taken away. This conditioncan be avoided by designing the linkage for a fi
48、ctitiousfloating angle somewhat higher than the maximumactually reached in flight. If h_,t is the differencebetween the floating angle at which the differentialgives complete balance and the actual floating angleof the aileron in the given flight condition, the resultantstick coefficient Ch, will be
49、Stick moment p . dChd_, d_,_ (20)where 0 is the angular deflection of the control stick.In any given case the stick force can be balanced outat only one angle of attack and, in general, the balancingeffect diminishes as the angle of attack is reduced.Hence, if the stick force is made to become zero at anangle of attack above maximum lift, overbalance ofthe control in normal flight will be avoided.A more or less compl
