~ RFORMANCE REPORT I♦ National Resean ·t"I Council Canada C. l.S . T.J. Ae ronaut i ea 1 and Mechanical Engineering Library Consetl national de recherchf Canada l.C . l.S. T. Bibl ioteque d aeronautique et de genie mecanique OtTE Re po rt n O • : ----=Q=-=C=---'----'4L.:v'.....IR,._D=---_,(:=--P::-'-/-=-0-=5'---'-{V).:....!...l.(<.>...---'/ __ ha s been D downgraded to : _________ _ @ de-clas sified by (Na me) : S. ll'l . L). 1-/ev\ r: e (Dept.): Di\ill Coordi o,;,/-o<" - /kcess lo T"ICer"".._1-,·., ,... Date: ,4uo '/- ''z.2 Signature A RCRAFl CF-105 F II EN T ru R EVISION No AV RpE CANADA LIM IT ED N 1 A I~ 0 TECHNICAL DEPARTMENT (Aircraft ) REPORT N< Monthly_!te por_!__No, l No nr SHCETS --- l ' REVISEO BV CF-105 MnHTHLY PF.RFC'R 1.1Al/(;E REPORT ( I ssued Mi d-l'onthly) This i~ Copy Number 9 Issued to1 R.C.A.F. 1:io•oo • oo •• • •• o • ~ • • Cl'3oeo o Date October 15th, 1955. ('IC, () () ()• •,,- O • V ~.Jo:i..; o OIJOQ W PREPARED BY HECKED BY 14 , 1951 _,)SUPERVISED BY APPROVED BY APPPQVED BV DATE Oct. 1955 , DATE' ]KTRODUCTION CF-105 MONTHLY PER1"0fWANr:E !< :) ~ ~ ~rr .. I 1 o __ I 1 ,. I I •• ·- ' I CJl I{) (') r u z ' '_j <[ 7 . I I 4- -- I : I ' LIMB 4 ~I KT . G P/PER F/105 I .l------1f-------4 I I. l ··--f-·- .: 1~- -- - N. I .oopo- 1 -t -- t - - j --1- I I ~, I , +- . - I ·-1- - ~ -& 01 45 --,.. I I - --l-- i, - I - -+ i t -t .. 7 1B. CF-105 PSRFOR!f,' ·;ss ,HTH ORE!'IDI\. PS 13 ENGINES (C,G. = 29% M.~.C.) The following CF-105 - PS 13 performance estimute is based on the wind tunnel configuration designated B2V1~1E1oll.sDg-4 over the subsonic portion, and configuration h9, NA5, B4, G3, V2, Rs, over the supersonic range. The particul~r feature of the former configuration is tho extended, notched, and carrbered leading edge of the wing, The drag of this configuration is st!mma rized, (Extract P/Aero Data/58), and is presented in SAction 2 of this report, The latter CQnfigurstion differs chiefl7 by not having a cambered leading edge. This drag d ta is given in P/Aero Data/48 but has not been summarized for this report. This constitutes little change under sµpersonic cruise conditions, and only decreases the supersonic drag by about 4% at maximum 1g 1 due to less elevator angle for trimo Thus, the performance does reasonably represent that for the one confiirurlltion, B2V1 tV1 E10N5D8 - 4, The PS 13 e11ei11e data is in a more incomplete state. The engine data above the tropopause was taken from the Deca 854 Memo, (Ref, Orenda Pll-1-1) on the PS 13, with the exceptio11 of the cruise operation at ,92 M.~I. and 40,000 Ft., where insufficient data was avai l able from the lt.emo, and we were forced to use the original PS 13 Brochure (EMS 8) Ap~ il 151, The memo of Dec. 1 54 assumes a 6.5 Sq. Ft. intake, and pressure recov-3ry cm·ve from P/Power/23 APP/r/10. It also considers the effect of a J? 11 ejector, as well as a bypass which o:i;ens to 118 sq. Inch~s. For engine performance below the tropopause the original PS 13 Brochure was used. The above mentioned pressure recovery corre-ction werA applied to this data 9 but no account was taken of the bypass effect. It should be noted that revised thrust estimates now being prepared indicate an increase in maximum thrust at 1. 5 M, 'I. of approximately J.%. T'nis offsets the slightly optimistic supersonic drags used in this report for the performance of the PS 13 engines version. SECREl LOADING AND PERFORMANCE P/ Peif/ 102 Performance Under N.A.C. A. Standard Atmospheric Conditions SECRET To R.c.A.F. Specification AIR 7-4 w~th Two PS 13 Enginee WEIGHT: Take-Off Weight with 15,510 Lb. Fuel (78.2% Max.) ••••••••••••••• Lb. Operational Weight Empti ••••••••••••••••••••••••••••••••••••••• Lb. Combat Weight (i/2 Fuel) ••••••••••••••••••••••••••••••••••••••• Lb. Landing Weight (With Reserve Fuel + Missiles) •••••••••• ••. ••• •• • Lb. Wing Loading at Normal Take-Off Weight •••••••••••••• •• •••••• Lb./Sq.Ft. Power Loading at Normal Take•Off Weight ••••••••••••••• Lb./Lb. Thrust SPEED True Air Speed in Level Flight At Sea Level at Combat Weight Maximum Thrust ••••••••••••••••••••••••••••••••••••••••••••••••• Kts. Military Thrust••••••••••••••••••••••••••••••••••••••••••••••••• Kts. True Air Speed in Level Flight At 50,000 Ft. at Combat Weight Maximum Thrust ••••••••••••••••••••••••••••••••••••••••••••o••••• Kts. CEILING Combat Ceiling at Combat Weight, Rate of Climb= 500 F.P.M. Maximum Thrust at 1.5 M.N. •••••••••••••••••••••••••••••••••••••• Ft. RATE OF CLIMB Steady Rate of Climb at Sea Level, Combat Weight Maximum Thrust at M.N. = .92 ••••••••••••••••••••••••••••• F.P.M. Military Thrust at 530 Kta. ••••••••••••••••••••••••••••• F.P.M. 55,889 40,379 48,130 42,200 44.5 1.19 t 720 650 1,110 62,200 50,000 25,200 Steady Rate of Climb at 50,000 Ft., Combat Weight Maximum Thrust at M.N ... 1.5 •••••••••••••••••••••••••••••• F.P.M. 11,500 TH'E TO HEIGHT Time to 50,000 Ft. M.N. • 1.5 from Engine Start at.Take-Off Weight~ 55,889 Lb. Maximum Thrust ••••••••••••••••••••••••••••••••••••••••••• Mins. 4°1 MANOEUVRABILITY Combat Load Factor at Combat We!ght Maximum Thrust at M.N. = 1.50 at 50,000 Ft. Placard Speed= 720 Kts. TAKE-oFF DJSTANCE SECKt:.T Take-Off Distance over 50 Ft. Obstacle at Sea Level••••••••••••• Take-Off Weight• 55,889 Lb. ' Maximum Thrust •••••••••••••••~••••••••••••••••••••••••o•••• Ft. Military Thrust•••••••••••••••••••••••••••••••••••••••••••• Ft. Maximum Thrust Hot Day •••••••o••••••••••••••••••••••••••••• Ft. 2,500 3,800 3,300 LM'DING DISTANCE Landing Distance over 50 Ft. Obstacle at Sea Level at Combat Weight Ft. 5,000 STALLING SPEED RANGE True Stalling Speed in Landing Configuration at Combat l7eight at Sea Level ••••••••••••••••••••••••••••••••••••••••••e•••••••••• Kts. 105 Combat Radius of Action at 50,000 Ft., Climb at M.N. = .92, Cruise out at Ll. N. = 1.5, Combat for 5 mins. at M. ll , "' 1.50, Cruise Back at M. N • ., .92, 15 Min. Stack at 40,000 Ft., 5Min. Fuel Reserve on Landing High Speed Mission with 15,510 Lb. Fuel•••••••••••••••••••• N.M. 200 High Speed Mission with Full Internal Fuel••••••••••••••••• N. M. 318 Combat Radius of Action at 50,000 Ft., Mission as above except Cruise Out at M.N. = .92 Maximum Range Mission with 15,510 Lb. Fuel••••••••••••••••• N. M. Maximum Range tlission with Full Internal Fuel •••• ••. ••• •••• N. M. Combat Radius of Action at Sea Level, Cruise Out at .6 M. N. and Combat at M.N. "' .92 at Sea Level, Cruise Back at .92 M.N. at 40,000 Ft., 15 Min. Stack, 5 Min. Fuel Reserve on Landing Sea Level Mission with 15,510 Lb. of Fuel Sea Level Mission with Full Internal Fuel •••••••••••••••••• .................. Ferry Range Mlssion at Economical Cruise Speed (M = .92 and Height, including 15 Mins. Stacking at 40,000 Ft., 5 Min. Fuel Reserve on Landing Range with Full Internal Fuel and 500 Gal. - External Tank • N.M. 315 491 217 318 1,675 IT~~ - :--- f -~ j + j I ~ -1 -~-sE4RE -.--~~1-+ -.-...---.----1-------, ~----'-X__.I. ---'Vi-'-v'\,J-'-_lj _G~. -+-.S_t.lS.;.._Tt.~....;::E=D _ t- -1-- l- _ _ _ _ ~ .... JN L£11£L VRN I lcoMf6.ilr w£1a-11r ' . ~ .L. 1+ -t- @,l.2Jt LB¾~ - I I I I • __ d= 1 a , r +- 1 t - I &c}c i._ .5.P..fEOJ+ -,- _r _ - I t I • -·•---l...-·+--~-+----+--+---+---'-· l . ~ i - -1 - ' ,_+ I • + • I ...,.__'-------+-_____ _.... , I -, • : • " • i ·r -~+~ 1:--i ~I t ~ - • J - - - -- --! T• -' +-- I I -+- I -1- 1 --j- -~ G.2 -0 ' I -4--- ----; ' '+ --+- - - _, _. ~ l I ~ ~. -~ • • ~ l f ~ -t - ' '" - ,::5 - •i -f-.- . . _j__r--17-t - - r- I- +--1 v• l ~ 1£)00 1 1 ' . j l t---r-~ r l -- _.J i I -+- - i .(40 _· ' : • ' ' f -1' ·•-- ,_,,!_~ _-r_~---+-----+--+- ~ - ' 4---"-'--1--l---l---i---'-.• -4---•--+-' ---+--J.1--+-f~-- l -- i- - -~ I C w - r ! l - r • t J Q, 000. p0,000 L8 __ ...._, I _ 6q00Q LB,---tl. ,-r--- I I i--:-d 10,boo 11 - ..1.-+ ,:_ -· ~ ~l---+--+---1-1 --+-----'-1_1 ___;t--+---+-+- I • I CuMs Ar -~1 M.N, MMll'IUI"\ THRUST ---4-5,000 LB. ,-~ _ .,.____.,.__50,000 LB. t ; ,.___...,._..-,.__55,000 1.8. i L I t-L; ' 'I T t- r I IQ.re:: OF C1..111a ; _j_ -----1---l---+- 60000 f I t -r- 1 + £CH£. ONc ~A'-F M1Nur£ I ALLOWED FROM E:.NTRIM CM 0 CM5 Extract P/Aero D~ta/58 (Continued) - Total Drag - Lb . - Ambient Pressure - Lb. / Sq.In. SECRET - Mach Number - Air~rart Effectjve Weight (Normal load factor x ajrcraft we1ght) - Lb. - Aerodynanic centre% M.A.C. - Centre of Grav:ity % M.A.C. - EJevator AneJe for Trim cf Aircraft at effective weight - degrees - Elev&tor Angle for Minimum drbg - deGrees - Non-linearity factor for CMI:> - Minimum Drbg Coefficient - Lift Coefficient at CDMIN corresponding to l:>TRIM - Aircraft Lift Coefficient - Aerodynamic Drag Efficjency Factor at OTRIM - Pitching Moment Coefficient at Zero L:lft - Elevator Pitching Effectiveness at constant CL From the ~rag curve, trus determinPd, the coeffic"ents can be approxi mate]y reduced to tr.e more conventional form. co~co + ~ intercept SECRET l , • I J I r I I l I I 1 I L -t t I ~ I • - t - - t r I ,- I ~u w i= 0 SECRET I I - t _.l,_,.. __ ..,_ t-- • ♦ I ♦- - - -+- t -----+- +- • 1 I L I I . ) . -- - ... __._ ...... I -+ ' l- 1 ( P/H v 51 1 T ECHNIC A CF-105 D PAR T'Vlf f/FT!iOD OF F.STil,.ATI!\G EJECTCR PFRFORl,'AKCE J.R, Y.onk Oct, 1 55 SECRET 1.0 INTRODUCTION The following report gives a metho::l. of using the experi mental results obtained by the fl.A.C.A. ejector tests to estimate the thrust of the Pratt & Whitney JT 4A/25 fitted with an ejector. It has been founc thut under certain conditions the ejector gives a significant increase of thrust, particularly at high speeds. The ejector also serves to give a cooling 1irflow ever tne engine. 2.0 THE JT 4.A/25 EJF.CTOR - L ~ ► ----... --==-~ i (Dp) l PR;~~y D A, e (Jlp1" ~ I /A/.._, ~ s Ee:: 1 JO ARY ___ _____ R+b---1 / __ -;, AIRFLQ'.,J The JT 4A/25 has a two position prim·1ry nozzle. The '1ozzle diameter DP is 34.'5 in. for afterburning and 28 in. for no after burning. On the CF-105, it is proposed to instal a secondary shroud of diameter D6 of 39 in. which projects beyond the end of the jet pipe by a distance L. This secondary shroud together with the jet pipe nozzle make up the ejector, The second'l.ry airflow cushions the expansion of the primary jet within the shrotrl a,'ld under certain conditions it gives an increase of thrust. The increase of thrust is due physically to the action of secondary pressure acting over the annulus area of the ejector. II /,(I 4/1/(R II I I it, I u TECHNICAL DEPARTMENT Aircraft CF-105 l11:THOD OF ESTIMATING EJECTOR PERFOffiMNCE 2 J.R. Monk Oct. '55 SECRET 2.1 Theoretically a larger secondary shroud diameter will give a greater gain in thrust but the increase is limited by two effects. (a.) The shroud must ron filled at its exit otherwise the secondarj pressure will fall to ambient pressure or lower a.oo there will be a. loss of thrust. The length of the shroud L, is here an important factor and there is an optimum length for any given shroud diameter. (b) The primary jet must not be over-expanded otherwise there will be a. shock a.rrl loss of thrust within or at the exit of the secondary shroud. 2.2 The secondary shroud diameter has been chosen as a. com promise between thrust boost when using afterburner and minimum loss of throat with afterburner off where effect (a.), mentioned above, operates. The shroud length L ha.a been chosen to give optimum perfonnance with the particular diameter chosen. Note that too long a shroud~ive rise to unnecessary friction loss. 2.3 It is convenient to express the ejector geometry in non dimensional fora by using diameter ratio D/Dp and spacing ratio r/Dp• The properties of the ejector a.re grouped into four parameters, secondary pressure ratio P /Pa, primary pressure ratio Pi/'Pa, mass flow ratio corrected II NO 1//l{R If I /111 1 ,//1 TfCHNICAL DEPARTMENT Aircraft CF-105 l.'.ETHOD OF ESTi liATING EJECT Ofl PFRFORl,.ANCT J . R. Monk Oct, 1 55 , E for effect of temperature ms fTs and gross thrusS.ECiR£T ~ fTp F8 y'Fjet which we cun regard as the ratio of ejector gross thrust to simple convergent nozzle gross thrust. 2.4 The characteristics of the JT 4A-25 ejector have been taken from NACA RM E52 L24 and are presented in carpet form in Figures 1, 2, 3 and 4 as follows:- Related Parameters:• A/B NO A/B lrls JTs , p s , Pp Fig. 1 Fig. 2 - N'P np Pa Pa r/Dp = 0.58 ' 0.52 DJDp = 1.13 1.43 Fig. 3 Fig. 4 Fej/Fjet, Ps!Pa 'Pp/%, r/Dp = 0.58 0.52 Ds/Up "' l.lJ 1.41 TABLE 1. Summary of the graphs of nozzle characteristics used in this repor t. Note that the above geometrical ratios tlo not correspond exactly with thosP now proposed for the JT 4A/25 to be fitted in the CF-105, but the differences are not thought to be significant. I CF-105 l/ll(/, II D PARTMENT A1rc1aft METHOD OF F.STH'. \Tim EJF.CTOR PERFORMANCE J.R. Pfonk P /PO"IF 11/ 51 4 Oct . 1 55 3 .o ES':'IM \TION OF THRUST BOOST GIVEN BY EJECTOR ( See Appendix I) The basic problem is to find the prim•t!) and second ·1ry pressures and the prim.1.ry and secondary airflows which act on the ejector. Once we know these ope:ating conditions, it is a simple matter to fin:l the increase of thrust by comparison with the experimental data given in NACA RN E52L24 an:l Figures 1-4. 3.1 The gross thrust ratio has been found as outlined above and is plotted in Figure 12 for.maximum afterburning and military thrust (maximum non-afterburning). The gross thrust has also been computed for partial afterburning con:liticns at M "' 1.5 and partial non-aftPrburning at M = 0,92. The calculations are listed in Part II of 1teport P/Power/51. We are here c?ncPrned primarily with thrust but the calculations have also enabled estimations to be made of secondary mass flow and pressure recovery in duct of engine airflow. 4.0 CONVERSIOI' OF GROS" TlffiUST R \TIO INTO NET THRUST '.'TITH ALLo;vA.NCE FOR SPILL \GE DRAG AND Il\'PUT IWl.1Et:rvr: Dll\G The starting point of the calculation is the net thrust given in the engine brochure. From this, we find the gross thrust of the basic engine by subtracting the total pressure loss and adding the input momentum. The basic gross thrust is then factored by the gross thrust ratio to give the actual gross thrust with ejector from . which we can find the actual net thrust (FAe) by subtracting the input momentum and spill11ge dr;9.g, ' llR /Rf/ DE PART ME~ T Aircraft J.J::(. Monk Oct. 1 55 CF-105 MtTHOD OF ESTIM 'TING EJECTOR PERFORMANCE SECRE1 The procedure of this part of the estim,1ticn is given in Appendix II. The actual net thrust was obtained for maximum after burning and military conditions over a range of Mach number and altitude and is plotted in Figures lJ and lJa for A/B lit and Figure 14 for no A/B. The thrust has also been estimated for partial afterburning conditions at M ~ 1.5 and partial thrust, no A/Bat M • 0.92. The thrust is also plotted as the ratio with atmospheric pressure an:1 as its r~tio with the basic net thrust with no ejector. 5.0 PROPOSED FUTUBE WORK It is hoped to issue a report in the near future to examine the following topics, some of which will represent a further examin ation of the somewhat arbitrary assumption used in the present report. (a) The interdependence of duct inlPt area, bypass inlet area and ejector ~ize and their relation to the thrust-drag summation. (b) A more rigorous determin3tion of bypass inlet total pressure. (c) A comparison of the ejector thrust-rain with that obtained from a convergent divergent nozzle of similar dimensions using one dimensional theory. L TECHNICAL DE'PARTME T Aircraft D P/Pu rn/51 6 CF-105 1.!ETHOD OF ESTIL \TING EJECTOR PERFORllANCE APPENDIX I J.R. Monk Oct.' '55 METHOD OF r~TIY'TING THE FQUILIBRIUM BYP\SS AIRFLOW SECRET A1ID THE A.'lSOCI 1.Tr:D PRE~nURf.S AND THRUST FOR GIVEN F.NGINE \JID FLIGHT CONDIT!• NS INTRODUCTION An iterative process is employed to establish the duct, engine arrl ejector operating point. Basically this involves finding the secondar,r airflow at which the pres".ure losses in the duct and bypass just account for the difference between the intake pressure and the secorrlary pressure at the ejector at the same secondary airflow. We require to know the following variables:- Flight Parameters Aircraft Mach Number Aircraft altitude and appropriRte ambient pres~ure arrl temperature Engine throttle setting (convenient to define as a percentage of ideal brochure thrust) Geometry Duct and bypass inlet areas EJector diameter ratio appropriate to afterburner lit or not lit and spacing ratio Presisure Losses Pre-entry shock loss Inlet separation loss Duct friction loss Pressure lose in bypass (at supersonic speed) (at low subsonic speed) TE CHNI CF-105 • M A !1.ETHOD UF FTB \TH G ~:JECTOR PERFORMANCE Oct. '55 SECRET Puge 13 shows 'i sped.rr,en c•1lcul!lqon whicl is exam~ ned, column by column in the following de cription:- Colunn Aircraft Mach mmber and altitude 1. Aircraft speed in knots 2. Aircraft speed in ft./ sec. J. Ide,11 engine airflo,r, read from engine brochure 4. Bypass area c9mmunicating with ejector, in sq.in~. 5. St~gn.tion spee~ of sound (I t l) '":" ' ~- 1- i" 6. Square root of stagnation temp. I \ t- ,_, ,__, . 7. Ambient atmospheric pressure Pa, lb./sq.in. 8. Star,nation pressure Po ., Pa , I \ \ 9. Shock loss ratio, ratio of tot'll pressur~ be ► ind shock structure to starn1tion pressure, P1, read from Fir,ure 8 10. 11. Po or at speeds of less than M = o. 5, there is 1 separ,1tion lofs, caused by breukway of the flow round the intake lips , given in Fi17,Ure 9. Total pressure at inlet PI = / P1tP 0 \PI 0 ~atio of total pressure of air entering compre~sor to inlet totnl prPssure, Pc. Note that when the bypa~s rills open Pr at approxirrntely !,'. .. 0.6, the duct boundary lay"r is t .ken into the bypasc so thit under thP"e conditions, the core of ,dr entering the engine suffer very little friction loss. 't r HCHNICAL DEP.ARTPi,f~T A1rc1df> METHOD OF 1:,TI!,: TING FJECTOR PJ'.RFORI,' \NCE J ,J:l.. Monk Oct . 1 55 CF-105 METHOD Cont'd. SECRET ,r r When tre duct inlet is choked it is possible to get a shock wave in the diffusing part of the duct. Thus it is conven ient to define Pc/Pr in terms of the duct choking parameter ma 0 /Pr by means of the following diagram:- Note: l/ ,, , r er, < • >- Ii, p 4 4 ~ 4,4, corrf>sporrls to choking mass flow in an inlet area of 5.6 sq.ft. or m" 1,0. m v vihen the bypass gills are closed use a value of Pcft'I = .965. 12. Ratio of total pregsure ut cornpres cor to stagnation i,r, (Co:. 9) x (Col. 11) 13. Actual engine airflow me= m 1 • Pc/P0 i.e. (Col. 3) x Col, 12), Duct inlet choking pararooter mac. l where m is total duct Pr 144g airflow. (i.e. m= engine airflow+ byp~ss airflow+ cooler airflow) ( Colunn 13) (Colunn 21) (Column 22) Make an initial gt:ess for ma 0 based on the engine airflow PT plus five or ten percent depending on the t,lach number -·~ 5€.CR~ \ ,.11,, q/f/fli-llt /l''llttl TECH~ICAL DEPARTMENT Aircratt PjPOj'iER/ 51 9 A CF-105 METHOD OF FSTIMATING EJECTOR PERFORMANCE J.R. Monk Oct . 1 55 NETHOD Cont'd. Column SECRE1 - U.. and then when the secondary airfldw is establisheq the initial guess can be checked. 15. 16. The ratio of total pressure of the air entering the bypass to the inlet total pressure, PB /Pr• Note that the secondary airflow is made up largely of the duct boundary layer. The total pressure of this flow has been assumed to lie half-way between the mean total pressure and the static pressure calculated from the mean duct Yach number at the compressor face. This assumption is justi- ~- fiable where the rr~an duct Mach number is moderate (for JT 4A/25 Mnuct f .6) PB/ Pr has been plotted in Figure 10. 17. The total pressure at the bypass PB= (PB /Pr) • PI• 18. The bypass airflow parameter m 8 ffo (where A: 74 sq.in~.). APBg This has a maximum value of .0123 when the inlet gills are choked. The quantity refers to the air going to-the ejector only and not to the cooler airflow. A suitable initial guess is .01. It is then adjusted in subsequent lines until the pressure l_osses through· the bypass just makes up the differ ence between the bypass inlet pressure PB and the ejector secondary pressure Ps. CF-105 CAL EPA T ENT Atr raf 1.:ETEOD OF J.'3TI!,,\TING EJECTOR PERFORliAllCE P /Pu711:J\' 51 10 Oct. 1 55 IIETI OD Cont'd . SECRET Column 18. The pressure loss through the bypass is Pn - P 4 so ms ffc APB must be adjusted until P4 = Ps, (i.e. P4 = Ps or column 20 Pa Pa = column 26) 19. The ratio of the static pressure in t}~ rear engine bay to the bypass inlet total pressure, P4 /PB• Hote th9.t the croos sectional area of the re,r engine bay is very laree, consequently the velocity is small arrl it me...· be assumed Fiirure 11 rivEs the ratio p4 /PB in terms . of the parameter m8 /T 20. p P4 Using column 7 and 17 convert _!:. into - • P3 Pa 21. Secondary airflow throui:rh the bypass, lb./ sec. Find from 22. columns 4, 6, 17 and 18. Oil cooler airflow, lb./sec. The inlet to the oil cooler is asRumed to be choked above a ~.ach nur.iber of 1.2, falling line,1rly to zero flow at lr' .. 0.5. T!":us the cooler airflow 0 can be represented in the follolling diagram:- I' •~ I I"- I 1/\ Arbitrary diagram of oil cooler airflow. CF-105 1R(k111 I 111,10 DEPARTMENT Aircraft METHOD OF ESTIN. ~TING EJECTOR PERFORMANCE J.P.. Monk P /PO",'EP/ 51 11 Oct. 1 55 METHOD Cont'd. SECRET Column 23. The primary jet pressure ratio Pp /Pa is calculated from the Pratt & l'lhitney Brochure (Curve 17227 for, no after burner, letter ,~ f Mnr'n for afterburner). For convenience 2 4 ~::. these have been replotted in Figures 5 and 6. To get the actual Pp /pa, the ideal Pp /pa from the brochure is factored by the duct loss r atio Pc /P0 • 24. The temperature corrPction factor .JTs to be applied to the Tp secondary and primary airflow. Ts has been defined as Ts = T0 + LJ T where l:l T is given in the following table:• TABIE II Mach Number 0 - 1.3 1.5 - 2.0 A/B Lit 15°c . 50°C 6T No A/B 100°c 75°c 25. Corrected mass flow ratio :: •R 26. Secondary pressure ratio P s /pa • This is found from Figure 1 or 2 at the particular primary pressure r 'ltio and corrected mass flow ratio corresponding to the value chosen , for ms ffo in column 18. As mentioned under column ( 18) APi3 s _\., I' I II r ~ I /£ ,J P/POr.FR/51 12 TEC HN ICAL DEPART MENT Aire raft CF-105 METHOD OF ESTildTING EJECTOR PERFORl✓. ANCE J.R. Monk METHOD Cont I d. ~ 26. the value of the secondary airflow parameter, ms /T, is AP p p irr.proved by iteration in subsequent lines until 8 = 4 Pa Pa 27. r· Gros s thrust ratio :.:u_. This is read from Figure 12 for the F Jet value of primary pressure ratio an::l the value of secon::lary pressure established in Column 26 above. Part II of :he Appen::lix details the procedure for finding the value of the thrust increase from the gross thrust r atio. Oct. '55 • • • . . " 25,oco 451000 {{)A£_, A/6 {it M= f,?] {;;{, ::: I ·-/1 4-/ I 0 To - 2 3 {ri&-j;; f&c 11/iaJ e,tn..t a i:.,ft,:1. .. > ff) 8'f- ISXJ 33 4- 1038 17 51 ZSI I O 3 7-5 1 7 S I • I ~ 4- aH'~ 1 71u:,, tJ1., n) /:, (N:U-, 21.;.J:J / •. ~C,,,1,1, ' , 606/i Cl .77 90 - 0 0 4 5 6 7 8 9 10 A ffo P0v ?o · s}igk Clo {fk J/J P.x. /I Ii/A,, /J 7 '-t- !30f ;9.g 5- ~52 31·3 .C/2=> zq . . > It r------tt---- -1----11---~--+---~--+---+---+--- -+--~ 0 ~ 111, n) CJ ( ;•"<' /;. ., A VR 24 J:J 'txci rytk. - •. ~/;, ,1,i.l:e ri o-(- ~( (. &.,;,,.. ~~J tl~tMt .to.,lio , AIRC 90 WEI< C. G b ' ' 7 8 9 10 11 12 13 14 15 16 17 18 19 20 ::- P0v fo·· SMk Pc., Pc t?'l~L~i.L /ilao Pa Pe, m5 1To P4- P4 ll"~ -P:r. I 0 r»,;, f~ tl I h Pr f.,o Pr O Pz A f&-g, ft, • Pa- JhjS ~o:: F,.'\ -( ,'. '✓ 1 !!J( ( rn V ~ F} } FA-e '-== ( Fn '_ 6Fo + (Y\ . _+ 9 ) FJo 1 t- F} } ~.,-'J ( x -A(/.) Y)14-4-- . A - rn WEIGHT - ( rn V + l.1 Fs = "° o.'° e~ 9 fv'l.,, 9 C. G. POSIT 7 8 9 ' 10 11 12 13 14 15 16 17 18 19 20 21 (h ::- I V / h V Fn ·-/:Jo Fei {1 3)•-(/1+ ) a p o() f,J.) Ao y A,,_, m e.. rn5, f!' i fl1 +ms+ rn,;_ ff -f m V .:,<.) f1jse, _ __, /tlvc. , I ' 3'1-~ &<:,' ) \ • O[~ 31'1 3t,, I lo I lfo 3· I G,J /7, If 2 A V ROE. CANADA LIMITED TECHNICAL DEPARTMENT (Aircraft) I AIRCRAFT CF /05 J" 75 EJECTOR FINALISED DIMENSIONS ""1-6 ·6' I/ Ri.PL..H N SHEET No f'~ P_uwu 51- 17 --- -- - -- - - - 'P REPARED SY OAT~ :r MONK 31 Al)u 5c ,____ _______ ,___ •• CHECKED BY 0 0:::, ~ I DATE I ------ --- - ---- --- SPACING RATIO DIA. 1 RATIO A/ B ON O· 6/2 I· 13 A/B OFF O· 507 I· 3 9 AV ROie CANADA L IMITED RE.POR T l\i0 P/ POWER/ 51 NTARFO TECHNICAL DEPARTMENT (Aircraft) SHEET No __ _!!_ _ PRE P AR E D B Y CATE I A IR C R A F T J.R. Monk Oct. 1 55 CF-105 JT 4A-25 >----- - ~I ~ >-:.. ..QI L EI-L j l iii? , I 4 . [() "" z 0 N tl 8 ,ril -- - ~L~ ..,j (~~--- C H ECKED B Y DATE ----+---- l\ SECRET cl. \\ 0 v w h w \ \ C z \lj ~ ~ ll)I L I l!) 0 I ._ I -';:; (lJ / ., .. a,1,, w JI ,_ z Q) ~ 0-, <..'.) C LU c; z t-.s= LLl -.) <./) ~ >-- (/) 1.. C' I-fU ...t.. I 0 i: "' . cY \D <11 ~ .J J t') • r JI z 6.2l • (}. 16 0. t4 0 iO L T - f -r FIG 1 _ MASS FLOY'{ RAtlO AGA.ULSff _ FOR, DIAMf.:TER RATIO = I. I SPACING RATIO =0.5 +-- + ·- -+- 1, r T- - - .- ' ! L Q J.L L • 4~l i _J , - - : I I i I I ! ,_ I I - - .j.. t \ - • - [_ -- NOTE - FOR. SP~CING !---i----1 --i---1------1-----1--l- "-'--+-1-,-+----!----l • : -+ . I • -- - - i - - j -7 ~ F rt.J _l_ FJc1 - --r-_L ' • • t -----,-1- r --Ti,. 1· L 7 1· + -~-t -~ - -+-- 1 •·-- + . .-- - l ..... -i. ; +. i - ~ • - - I· -t O;O X HE "" 359 14L • ~--~--~----,.---- -r -,- r-,-'!""....,. -- --------..,. T..,."" T - r- " + F1G. 13. - ~ J75 NET THR'UST f"Ao AGAINST 1 MACH NUMBER A/B LIT " -l BY PASS AREA = 74- o DIAMETER RATIO"' H3 , t ~ SPACING RATIO :: •5'if + - -c ----t· t + -t L1+- r t- T j. j. I • t - + - -J- ...... t- T t-. "i" t I + t -I- - t -+- ~ +- - t- -.:j- - + - / -- ~ t • t- - -'--- T +- T I- t j. -.--t- - ' ~ - I ,j -+-- ---+ ~r--- t I ' T-, t t 4- - I 1--- -+- - __:;_ - t---r·~- 1 I· I ~ t 7 1 t· : -+- -+-- - r-.·- t T I t-- -+- +- 3!'jooo' I t -+ r I I 1~ _J_ t- l 0- t - l \j) I t I L ~ t t } ~ . ~ -+· I ' -- ~-r ~ - j FtG I~ a.. T \ f T ' l ; SECRET- • • Sl:CRET 1-· \ \ ..,_ \ t l i - · r • 1 f - - t- 1 I I r1- + -+ l ' ' ' .... ·'f.l. -l •• -- I 0) I{) "1 ; r ~ - ) -1- HO T f FJG ) 3 c ~ r I I .j.. • '._J - .t -~L I +- .j - --~ :_1 I I -. r ~-l-- f • r I L --j f ' -~- 1 ...j... I • - . - __ L ., 5.L. - , • 15 000 cµ 2~000' :r 35,000 1 4 't5,000' a 551ood r ~=rl-t =.J==.: r;;_,_,_r_r1..__·...._7_,_Q. tlt11L . v_._ M }Jo · ffi ,A)..I LW •·--4...........+---.- ~+---1----i L_ -+ + 1_·-+-r -+-_.__+-----.----1--~...--.__.__. 1 -~l __ j_l I ---- I I+· • ·+ + + ! -I -+---+------.--· i -~ . :i ~ - ~~--+--+"--..___.l__,____._ rt ~ -, _J_ . f- :a -i--...+-...c..........;,----"----+-~-+-· -r-· w z . r -, I 1 ~ r -? ~------+----+-----------+' ---+---1- 1 r~ ' -"--I-..._ ~ f+. r -~-~-1----+--!---l----l------1----+---------.--4--..;..... -1---~ l.- ~ ~ -~~~~---~ ' - . -· - - --- - --- I ! --,. - -· : --- • ~ljl'--'---'---'-----'----'-~~~~-=======~-=~~........_.., I + ' A j__ - 4 ' I I i - _I_. "t ~ I -i- - I -i i I ' ..... _ -+--- i --•i- ............._ __ +--- + -l-----i--- + t 1 I ·--t-----!'-'-·- 1 8 ---~ SECRET f T - -t, -r I -~->----+-+--+-,-+-- t-l---+-~;..-.+~--'-----4-+- .. ___ ! i +--+---.--!---+-+---'--..__._- • Pc. Po ·'i? '96 - ,'fS: - ·"It • -~3 ·72, ... " C -

CF-105 monthly performance report no. 1

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