FlLE lN Vf\ULT. u .. ,n.JLfts,sr1ED AV R 0 SECRET C A N A D A SECURITY WARNING This document. is classified SECRET and is intended solely for the use of the recipient and may be used only in connection with work carried out for, or on behalf of, the Government of Canada. THE UNAUTHORIZED RETENTION OR DESTRUCTION OF THIS DOCUMENT IS PRO HIBITED. At all times this Document must be hand led in accordance with the existing regulations pertaining to information classified SECRET published by the Canadian or U.S. Governments. This brochure has been specially pre pared for the Canadian and United States Governments and their attendant Services. The information contained herein is the property of A. V. Roe Canada Limited and the Canadian Government and the recipient shall not transfer, copy, reprint, divulge or use any portion of it without the prior written consent of A. V. Roe Canada Limited. This is Copy Number ••• 'J ._,, . . 3 9 ... •·• ..... ANALYZED . j, , .) . inAG J.r1fH ~ Issued to: .• . f/J.. ,. ~ ... ~{./LP. Date: 1 6 .ft/lr l'/5 4 . .............. \· ......... . AV R 0 C A N A D A CF-105 TWIN ENGINE SUPERSONIC ALL-WEATHER FIGHTER JULY 1954 A. V. R O E • C A N A D A L I M I T E D AIRCRAFT DIVISION MALTON - ONTARIO SECRET AV R 0 C A N A D A I ' CF-105 SUPERSONIC ALL-WEATHER FIGHTER AV R 0 PARA. 2 3 3. 1 3. 1. 2 3. 1. 3 3. 1. 4 3. 1. 5 3. 1. 6 3. 1. 7 3. 2 3. 3 3.4 3. 5 3. 6 3. 7 3. 8 3. 10 3. 11 3. 12 3. 14 3. 15 3. 15. 1 3. 15. 2 3. 16 3. 17 3. 18 3. 19 3. 20 3. 22 4 5 C A N A D A TABLE OF CONTENTS SUBJECT SCOPE APPLICABLE SPECIFICATIONS AND OTHER PUBLICATIONS REQUIREMENTS: Characteristics Performance Weights Center of Gravity Areas Dimensions and General Data Control Surface and Corresponding Control Movements General Aerodynamics Structure Design Criteria Wing Group Tail Group Body Group Alighting Gear Surface Control System Engine Section Propulsion Instruments and Navigational Equipment Hydraulic, Emergency Air and Pneumatic Systems Hydraulic and Emergency Air Systems Pneumatic System Electrical Electronics Armament Furnishings and Equipment Air Conditioning and Anti-Icing Equipment Auxiliary Gear SAMPLING, INSPECTION AND TEST PROCEDURES PREPARATION FOR DELIVERY PAGE 1 3 3 3 27 30 30 30 32 32 38 52 55 58 59 61 64 65 67 76 76 76 80 81 86 92 9-4 96 97 97 97 AV R 0 FIG. NO. l 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 C A N A D A TABLE OF CONTENTS {Cont'd) LIST OF ILLUSTRATIONS TITLE 3 View General Arrangement Level Flight True Air Speed Level Flight Mach Number Maximum Steady Rate of Climb Time to Height from Engine Start Distance to Height from Engine Start Fuel to Height from Engine Start Take-off Distances - Sea Leve-1 Take-off Distances - 3, 000 Ft. Landing Distances High Speed Mission Maximum Range Mission Long Range Ferry Mission Acceleration Time in Level Flight Manoeuvrability Performance at Altitude Performance Flight Envelope at Sea Level Performance Flight Envelope at 40, 000 Ft. Performance Flight Envelope at 50, 000 Ft. Equipment Arrangement CF-105 Component Breakdown Profile Drag Coefficient Drag Efficiency Effect of Elevator Deflection on Minimum Drag Coefficient Lift at Minimum Drag Structural Flight Envelope at Sea Level Structural Flight Envelope at 30, 000 Ft. Structure Arrangement Engine Installation and Removal Engine Airflow and Cooling Fuel System - Diagrammatic of Fuel Feed and Transfer Fuel Syi,tem - Diagrammatic of Fuel Pressurization Fuel System - Diagrammatic of Refuelling and Defuelling Hydraulics Schematic Flying Controls System Hydraulics Schematic Utility System Low Pressure Pneumatic System Schematic PAGE 2 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 33 34 42 44 45 46 53 54 56 66 70 72 73 74 78 79 82 Electrics Circuit Diagram 83 Armament and Electronic Equipment Installations - Falcon 9 3 Armament and Electronic Equipment Installations - Sparrow 2 95 AV R 0 C A N A D A 1. 1 Details of the following airplane are covered in this brochure: - Service Model Designation •••••.•••..•• Designer's Name and Model Designation ......................... . Number and Places for Crew .•.••••..•• Number and Kind of Engines ..•...••.••• Supersonic, all-weather fighter A. V. Roe Canada Ltd., - CF-105 Two (2)-Pilot & Radar Operator Prototypes Two Curtiss-Wright· Corp., turbojet engines, YJ. 67-W-1 fitted with afterburners Production Aircraft To be decided. 1. 1.1 The main role of the aircraft is high altitude, all-weather, night and day interception and destruction of enemy bomber aircraft. 1.1. 2 The secondary role of the aircraft is low altitude, all-weather, night and day interception and destruction of enemy bomber aircraft, However, the air craft will be designed to fulfil its primary role and limitations will be accepted in the fulfilment of its secondary role. 2 APPLICABLE SPECIFICATIONS AND OTHER PUBLICATIONS 2. 1 Specifications and publications used in the preparation of this brochure are as follows: - SECRET (a) Handbook of Instructions for Aircraft designers, ARDCM 80-1 Edition including revisions up to and including January 1954. (b) Air Force (USAF) Model Specification MIL-1-6252 B(ASG) dated 15 April, 1953. (c) Manual of aircraft requirements for the Royal Canadian Air Force CAP 479, dated 20 May; 1954. (d) Engine Model specification YJ. 67-W-l, Curtiss-Wright W.A.D. Specification No. 880-A, revised 6 February, 1952. (e) Installation data, YJ. 67-W-1 turbojet engine, Curtiss-Wright publication A. E. D. -26-B, dated November, 1953. AV R 0 C A N A D A I SECRET CJ .. .. 1111 Ill AVIONICS ARMAMENT EQUIPMENT FUEL ENGINE 1---- - --------- --------- ------ - 66'-6 t------------------------ -----73'-1.6! FIG,. 1 3 VIEW GENERAL ARRANGI AV R 0 C A N A D A 3 REQUIREMENTS 3.1 Characteristics : 3.1. l Refer to Figure l on the preceding page for information on the configuration of the air plane. 3.1. 2 Performance: 3.1.2.l Tabulated Performance: TABLE l PERFORMANCE UNDER N. A. C.A. STANDARD ATMOSPHERIC CONDITIONS True Air Speed in Level Flight at Sea Level at Combat Weight (45,000 lb.): Maximum Thrust with Afterburners ..•.•...••••• • ••• •. •. Military Thrust ...••.•.•••..•. • .•. • .••......••.•..•... True Air Spee.din L evel Flight at 50,000 ft. Altitude at Combat Weight {45,000 lb.): Maximum Thrust with Afterburners .•••• , .. • •.••••...... Combat Ceiling (rate of climb = 500 f. p. m.) at Combat Weight (45,000 lb.): Maximum Thrust with Afterburners at Mach No. = l. 50 Maximum Thrust with Afterburners at Mach No. = 1. 75 Steady Rate of Climb at Sea Level at Combat Weight (45,000 lb): Maximum Thrust with Afterburners at Mach No. = . 95 ..•... Military Thrust at Ma ch No. = . 95 ..••. • ................. . Steady Rate of Climb at 50,000 ft. Altitude at Combat Weight (45,000 lb.): Maximum Thrust with Afterburners at Mach No. = l. 50 •.. Time to 50,000 ft. Altitude and Mach No. = l. 50 from Engine start at normal take-off gross weight= 51, 0481b. Maximum Thrust with Afterburners .••..•.•••..•.••..... SECRET 780 knots 655 knots 1240 knots 61, 5 00 ft. 62, 500 ft. 46, 500 f. p. m. 11, 700 f. p. m. 10, 500 f. p. m. 4. 3 min. 3 AV R 0 C A N A D A TABLE 1 (CONTINUED) Combat Load Factor at Combat Weight {45,000 lb.): Maximum Thrust with Afterburners at Mach No. =1. 50 at 50,000 ft. 1. 78 g. Maximum Thrust with Afterburners at Mach No. =1. 75 at 50, 000 ft~ 1. 90 g. ,,-, Take-off Distance over 50 ft. Obstacle at Sea Level at normal take -off gross weight (51, 048 lb.) Maximum Thrust with Afterburners ..•.••••.• • •.•••••.. Military Thrust ..•...•. • .••••.•• • • . • .••• • .•. • • • ..•.•.. Landing Distance over 50 ft. Obstacle at Sea Level at Combat Weight {45,000 lb.) ...•.•••••.• •• ••••• • •. • •••••• True Stalling Speed in Landing Configuration at Sea Level at Combat Weight (45,000 lb.) ..•.•• • •.•.•• • ••••.•• • •... Combat Radius of A ction with Combat at 50, 000 ft. Altitude: High Speed Mission (Table 3 - RCAF Spee AIR-7-4) .....•.. High Speed Mission {Full Internal Fuel) •.• •. • •••• • ••••••• • Maximum Range Mission (Table 4 - RCAF Spec.AIR-7-4) ... Maximum Range Mission (Full Internal Fuel) .•..•.•.•..•• , Ferry Range Mission {Table 5 - RCAF Spee. AIR-7-4) •••••. 3,1. 2.1.l Engine Performance: 3000 ft. 5300 ft. 5900 ft. 104 knots 2 00 naut. mi. 436 naut. mi. 306 naut. mi, 655 naut. mi. 1869 naut. mi. TABLE 2 - PERFORMANCE RATINGS OF YJ. 67-W-1 ENGINE WITH AFTERBURNER UNDER N.A.C.A, STANDARD SEA LEVEL CONDITIONS Condition Low Pressure Thrust Sp. Fuel Consumption Cor:1 press or R.P.M. Lb. Lb. /Hr. /Lb. Thrust Maximum 6,175 21, 500 2.05 with A/B Military 6,175 13,200 .795 Normal 6,175 11, 700 .760 A V R O C A N A D A 3. 1. 2. 1. 2 Combat Radius - High Speed Mission: TABLE 3 - COMBAT RADIUS OF ACTION - HIGH SPEED MISSION (TOR. C.A. F. SPEC. AIR-7-4) Condition A. Start Weight B. Engine Start C~ Take-off to Unstick, Mil. Thrust D. Accel. to . 75 M, N. Mil, Thrust to . 95 M. N. Max. Thrust with A / B E. ClimbatM.N. =.95to40,000', Max, Thrust with A / B F. Ac eel. to 1. 5, Max. Thrust with A/B G. Climb to 50,000', 1.5 M.N. Max. Thrust with A / B H. Cruise out at 50,000', 1. 5 M. N. I. Combat50,000 1, l.5M.N.,Max. Thrust with A / B J. Descend to 40,000' K. Cruis e Back at speed for Max, Range L, Stack at 40, 000 1 at speed for Max. Enduranc e M, Descend to S. L. N. Land with 5 Min. Fuel (Max.End.) TOTAL Distance Naut. Mi. 7.4 12.3 15,7 10. 5 154. 1 200.0 400 Combat Radius of Action = 200 naut. mi, Time Mins. • 50 .36 1.26 1.46 1. 35 . 70 10.70 5.00 1. 10 22.00 15.oo 4.63 5.00 69.06 Fuel Aircraft Consumed Weight Lb. Lb. - 51, 048. 100 50,948 132 50, 816 730 50; 086 1560 48,526 895 47,631 460 47,171 - 2735 44,436 2540 36 1660 1000 192 580 12, 620 40, 840 * 40,804 39, 144 38,144 37,952 37,372 *l, 056 lb._ of armament fired NOTE: With full internal fuel the combat radius is increased to 436 naut, mi. 5 A V R O C A N A D A 3, 1. 2, 1. 3 Combat Radius - Maximum Range Mission: A. B. c. D. E. F. G. H. 1. J. K. L. M. N. TABLE 4 - COMBAT RAD.iUS OF ACTION - MAXIMUM RANGE MISSION (TOR, C,A, F, SPEC, AlR-7-4) Fuel Aircraft Condition Distance Time Consumed Weight Naut. Mi. Mins. Lb. Lb. Start Weight - - - 51,048 Engine Start - . 50 100 50,948 Take -off to Unsti ck, Mil, Thrust - .36 132 50,816 Acee!, to. 95 M, N,, Mil. Thrust ll,6 1.71 675 50,141 Climb at • 95 M. N,, Mil. Thrust 50,0 5.45 1320 48,821 Cruise Out at 40, 000 1 (Max, Range) 220.0 24.20 2155 46,666 Accel. to 1. 5 M. N, at 40, 000 1 15,0 1.28 850 45,816 Max, Thrust with Afterburner Climb at 1. 5 M. N, to 50, 000 1, 9.4 . 65 440 45,376 Max. Thrust with Afterburner Combat atl,5 M,N, at 50,000 1 , - 5,00 2540 41, 780* Max, Thrust with Afterburner Descend to 40, 000 1 (Idling R.P.M.) - 1. 10 36 41,744 Cruise Back at 40, 000 1 (Max.Range) 306,0 33.60 2600 39,144 Stack at 40, 000 1 (Max, Endurance) - 15, 00 1000 38,144 Descend to Sea Level - 4.63 192 37,952 Land with 5 Min, Fuel (Max. End.) - 5. 00 580 37,372 TOTAL 612 98,48 12, 620 Combat Radius of Action = 306 naut. mi, *l, 056 lb. of armament fired NOTE: With full internal fuel the combat radius increases to 655 naut. mi. A V R O C A N A D A 3,1. 2.1. 4 Ferry Range Mission with Full Internal Fuel plus 500 Imp, Gall, under-fuselage external tank, TABLE 5 - FERRY RANGE MISSION (TO R.C.A.F, SPEC. AIR-7-4,) Fuel Aircraft Condition Distance Time Consumed Wei ght Naut, Mi. Mins. Lb. Lb. A. Start Weight - - - 62,808 .- B. Engine Start - . 50 100 62,708 c. Take-off to Unstick, Mil. Thrust - . 54 210 62,498 D. Accel. to . 95 M, N,, Mil, Thrust 16.0 2.28 935 61,563 E. Climb to 40,000', Mil. Thrust 73.0 8,0 1, 890 59,673 F. Gruis e at 40, 000 1, (Max. Range) 1,780 198.05 18,972 40,701 G. Stack at 40,000' (Max. End.) - 15.0 1035 39, 66(; H. Descend to S. L. (Idling) - 4.63 198 39,468 I. Land with 5 Min. Fuel (Max, End,) - 5. 0 600 38,868 TOTAL l, 869 234 23,940" ' ;Ferry Range = 1869 naut. mi. NOTE~ Armament, and external tank are carried throughout mission, 7 AV R 0 C A N A D A 3.1.2.1.5 Combat Fuel Allowances: The following table sets forth the combat fuel allowance based on 5 minutes at a Mac h number of 1. 5 at maximum thrust, together with the equivalent cruise range at a Mach number of 1, 5 at 50,000 ft, and at maximum range speed at 40, 000 ft, Comparative allowance for turning through 180 degrees at the r espe ctive altitudes is also given, TABLE 6 - COMBAT FUEL ALLOWANCES AT COMBAT WEIGHT ( 45, 000lb,) Altitude Combat Fuel Range at Max. Range Fuel for 180° Allowanc e M = 1. 5 at 40, 000 Ft. Turn at M= 1. 5 Ft. - lb, a t 50 , 000 ft. Naut, Mi, - Lb, Naut. Mi. 55,000 2,000 115 215 970 50,000 2,540 145 2 75 820 ' 45,000 3,220 185 350 750 40,000 4, 120 235 445 710 35,000 5, 160 290 560 740 3.1, 2 , 1. 6. The performance specified herein is based on estimated specific fuel consumption of the two Y J. 67-W-l engines, each one fitted with an afterburner. Turbojet engine fuel of spe c ific gravity = • 75 is assumed for performanc e esti mations. 3, 1, 2.1. 7 Drag Estimate: Performanc e estimations specified herein are based on the dra g e stimate detailed in sub-paragraph 3, 3 8 AV R 0 C A N A D A '3,1.2.1.8 CombatRadius - SeaLevelMission: TABLE 7 COMBAT RADIUS OF ACTION - SEA LEVEL MISSION (TWO ENGINE CRUISE} Fuel Condition Distance Time Consumed Naut. Mi. Mins. Lb. A. Start Weight - - - B. Engine Start - . 50 100 c. Take-Off to Unstick, Mil, Thrust - . 36 132 D. Accel. to Cruise Speed, Mil. Thrust 4.0 . 85 340 E. Cruise out at S. L. (Max. Range) 210.0 31. 80 5,250 F. Accel. to ,95 M,N. at S.L., 6.0 . 78 310 Mil. Thrust G. Combat at . 95 M, N. at S, L., - 5,00 2,150 Mil. Thrust H. Climbto40,000', Mil. Thrust 40,0 4,40 1,080 I. Cruise Back at 40,000' (Max. Range) 180,0 19. 75 1,486 • J. Stack at 40, 000 1 (Max, Endurance) - 15,00 1,000 K. Descend to Sea Level - 4. 63 192 L, Land with 5 Min, Fuel (Max, End.) - 5,00 580 " TOTAL 440 88,07 ~ 2,620 Aircraft Weight Lb. 51,048 50,948 50,816 50,476 45,226 44,916 41,710>\' 40,630 39,144 38,144 37,952 37,372 Combat Radius of Action = 220 naut. mi. *l, 056 lb, of armament fired NOTE: With full internal fuel the combat radius increases to 433 naut. mi, AV R 0 60 so tu 40 ~ I.I. 0 "' C z ~ :::, 0 ~ 30 ' w C :::, ,_ ;:: .... c( 20 10 CANADA MILITARY THRUST MAXIMUM THRUST-- WITH AFTERBURNER COMBAT WEIGHT=45,000 LB. S.L. -+------.-------,-------r-----.-,r-------,------------ 0 200 400 600 1,000 1,200 TRUE AIRSPEED - KNOTS FIG. 2 LEVEL FLIGHT TRUE AIR SPEED 10 AV R 0 80 io 60 50 t; ~ ... 0 "' Q z ~ ::, 40 0 :i: ... "' Q ::, ... 5 C. k::. -t----. 60 50 I ... w w ... 40 I ... 0 .,, C z < .,, ::, 0 :r ... ' w C 30 ::, ... ;:: ... < 20 10 0 + - --+ - J ~ ~ ---t" - --+- t .. ACCELERATION --- - .. I i t ~ .. - "t ,: -r L DISTANCE FOR Ml~~MiE~G~~ME I DISTANCE FOR CLIMB USED IN HIGH SPEED MISSION -+----4 I ~ DISTANCE FO CLIM~D IN MAXIMUM R NGE MISSION t----t--;---+- --i----t-- --- I 1 t t ! .. t -- I TAKE-OFF WEIGHT =51,048 LB. 1 .... i I ------+---------,---.;....---,-------,---__; 20 40 60 80 100 DISTANCE TO HEIGHT - NAUTICAL MILES FIG. 6 DISTANCE TO HEIGHT FROM ENGINE START 14 AV R 0 C A N A D A .... "' "' ... ... 0 .,, C z :,i ::, 0 :r .... ' "' 60 50 40 C 30 ::, .... ;:: ... < 20 10 0 SECRET FUEL FOR CLIMB USED IN HIGH SPEED MISSION~ ACCRERATION TO 1.5 MACH NO. 7 FUEL FOR CLIMB USED IN MAXIMUM RANGE MISSION 1,000 2,000 ...._______ FUEL FOR MINIMUM TIME TO HEIGHT \_ ACCELERATION TO 1.5 MACH NO. TAKE-OFF WEIGHT - 51,048 LB. 3,000 4,000 5,000 FUEL TO HEIGHT - (LB.) FIG. 7 FUEL TO HEIGHT FROM ENGINE START 15 AV R 0 8 7 6 ... "' "' u.. u.. 0 5 .,, 0 z c( .,, ::) 0 ::c ... ' "' 4 u z c( ... .,, c 3 2 40 42 C A N A D A r f LEGEND :i AT MAXIMUM THRUST WITH A/B AT MILITARY THRUST ---- TO CLEAR A SO FT. OBSTACLE \ GROUND RUN N. A. C. A. SEA LEVEL STANDARD 44 46 48 50 52 54 56 58 60 62 64 A/C WEIGHT - THOUSANDS OF POUNDS FIG. 8 TAKE- OFF DISTANCES - SEA LEVEL 66 16 AV R 0 ... ... ll! ... 8 7 6 0 5 :g z ~ :::, 0 :c ... I ~ 4 z ~ .,, 0 3 2 40 C A N A D A TO CLEAR A 50 FT. OBSTACLE ~ GROUND RUN MAXIMUM THRUST WITH A/B N. A. C. A. 3,000 FT. STANDARD (+40' F) 44 48 52 S6 60 64 TAKE-OFF WEIGHT - THOUSANDS OF POUNDS FIG. 9 TAKE-OFF DISTANCES - 3,000 FT. 17 8 7 'II -a . ... ti 0 ... ... ... 06 .. "' • 0 z z ~ a ::, - 0 z :c a ~ a ... - ~ 5 "' <( ~ .. "' • i5 z " Ill "' 4 3 ii 36 40 44 48 52 A/C WEIGHT - THOUSANDS OF POUNDS GROUND RUN N. A. C. A. SEA LEVEL STANDARD USING BRAKE PARACHUTE AT TOUCHDOWN 56 60 64 .. < :1111 0 ft .. z .. ~ .. :0 'II -Q .. .. z -Q z UI ,. Ill Ill a a -UI UI -0 z 60 t:; ~ so ... 0 Ill Q ~ Ill 5 40 :c ... ' l!l :::, ... !:; 30 " ~ " 8 c 20 z " ... Ill 5 a.: V 10 S. L. J 0 HIGH SPEED MISSION WITH FUEL FOR 200 N. 2 CO~ T 1 RADII - HIGH SPEED MISSION WITH FULL INTERNAL FUEL -----------------~ --------1-- - : -~-------- MISSION AS SPECIFIED CANNOT BE ACCOMPLISHED BELOW APPROX . 35,000 FT. so 100 150 : I : "\ 1 I I t I I I ,' 120 .,l 2s : Jo ,, I , I ,, : ,,,' ~ .... ,, TIME TO COMBAT FROM ENGINE START (MINUTES) 1 ------------ --------- I 200 CONDITIONS 1. START TAKE - OFF AND ACCELERATE TO .95 M. N. 2. CLIMB TO 40,000 FT. WITH MAXIMUM THRUST + A/B 3. ACCELERATE TO 1.5 M. N. WITH MAXIMUM THRUST + A/B 4. CLIMB TO COMBAT ALTITUDE WITH MAXIMUM THRUST + A/B S. CRUISE OUT AT 1.5 M. N. AT COMBAT ALTITUDE 6. COMBAT FOR 5 MINS. WITH MAXIMUM THRUST + A/B 7. DESCEND TO 40,000 FT. 8. CRUISE BACK OVER BASE AT .95 M. N. AT 40,000 FT. 9. STACK FOR 15 MINS. AT 40,000 FT. 10. DESCEND TO SEA LEVEL, APPROACH AND LAND 250 300 350 400 COMBAT RADIUS -NAUTICAL MILES 1 .. < 11111 0 " .. z .. a .. • -a . -N ... ... ... u. u. :E 0 ,. "' C >C z - ;,; :E :::, C 0 :i:: :E ... ' ... :Ill C ,. :::, !:: z ... ... a < Ill ... < ea :E 2: - 0 "' u "' -0 z 60 50 ~ t 40 30 20 10 ~ _/ MAX. RANGE MISSION WITH FUEL FOR 200 N. M. HIGH SPEED COMBAT RADII -+- 20 .. t / +- ' I ' f ~ ~ ¼ :>0"'4 + +- +- I I r'" .L.., MISSION AS SPECIFIED CANNOT BE ACCOMPLISHED -BELOW APPROX. 30,000 FT. I MAX. RANGE MISSION WITH FULL INTERNAL FUEL I ---~----------:---------;-i----------.. . I I +- I I '\ I I I I I +; - t ~ -:- t - : .. ·~ : I I I : I I I I I I 40 1 I 50 I 60 II 70 I I I I I t I I I I I ·! I : :/ I I .!; ~ I I ,/ rl I ,, J ,, 1 : 1 , , I I ,, t ~ 1 - :.,/-- TIME TO COMBAT I )-' I ;' , , , FROM ENGINE START (MINUTES) .. . .._ ; ~ i t CONDITIONS 1. START TAKE-OFF AND ACCELERATE TO BEST CLIMB SPEED. 2. CLIMB TO 40,000 FT., MIL. THRUST. 3. CRUISE-OUT AT .95 M. N. AT 40,000 FT. 4. ACCELERATE TO 1.5 M. N. WITH MAXIMUM THRUST+ A/B 5. CLIMB TO COMBAT ALTITUDE WITH MAXIMUM THRUST+ A/B 6. COMBAT FOR 5 MINS. WITH MAXIMUM THRUST+ A/B 7. DESCEND TO 40,000 FT. 8. CRUISE BACK OVER BASE AT .95 M. N. AT 40,000 FT. 9. STACK FOR 15 MIN. AT 40,000 FT. 10. DESCEND TO SEA LEVEL, APPROACH AND LAND. s. L.-t--i......-=-------,---_;~--r-------,---__:.--r---,:__ __ ...,.... _______ :.,__ ___ _ 0 100 200 300 400 500 600 COMBAT RADIUS • NAUTICAL MILES .. < Ill 0 " .. z .. 11:11 .. A V R 0 t:; ~ ... 0 "' Q z < "' ::, 0 :c ... I w Q ::, ... ;:: ... < w "' 5 "' u C A N A D A 50 40 30 20 10 4 CONDITIONS 1. START ENGINE, TAKE-OFF AND ACCELERATE TO BEST CLIMB SPEED AT MIL RATING 2. CLIMB TO CRUISE ALTITUDE, MIL.RATING 3. CRUISE-OUT AT MAX. RANGE PROCEDURE 4. CLIMB OR DESCEND TO 40,000 FT. 5. STACK FOR 15 MINS. AT 40,000 FT. 6. DESCEND TO SEA LEVEL, APPROACH AND LAND TAKE-OFF WEIGHT = 62,808 LB. TOUCHDOWN (MINUTES) NOTE : - AMMUNITION AND 500 GAL EXTERNAL TANK CARRIED THROUGHOUT MISSION, NO CREDIT GIVEN FOR DISTANCE COVERED IN DESCENT, AT LOW ALTITUDE RANGE COULD BE IMPROVED BY CRUISING ON ONE ENGINE S. L . ..,..----,,--...;.,----,~--f----,---"'T"'"--"T"--""T"--""'T----, 0 200 400 600 800 1,000 1,200 1,400 1,600 1,800 2,000 RANGE - NAUTICAL MILES FIG. 13 LONG RANGE FERRY MISSION s- < 60 _J i t t r I I: ACCEL. FROM 1.2 M. N. TO 1.5 M. N. I ~ I- ~ " "II ··i • s--Q ,+ I j . I z .. ... ACCEL. FROM .95 M. N. TO 1.5 M. N. s- • I I / I I 1: " I I .... " w w 40 Ill u.. .. u.. Ill 0 llll .,, • C z ... <( - .,, 0 :) 0 z ::c .... 30 ... I w - C .. :) .... Ill ;:: ... - <( z .. 20 Ill < Ill .. "II .. -Q :z: 10 ... I I ~ COMBAT WEIGHT - 45,000 LB. MAXIMUM THRUST WITH AFTERBURNER , s. L. 0 1 2 3 4 5 6 7 ... TIME (MINUTES) ... 'II - " . -0, 'II ffl :a 'II 0 :a I: ,. z " ffl 'II .. - " z ... ffl z < ffl .. 0 'II ffl ,. ... "' ffl ,. .. ffl < ffl .. 8 7 6 5 4 I:.'. ~ 3 ... u ~ ~ 2 0 ... ~ C. G. AT .29 M.A. C. COMBAT WEIGHT=45,000 LB. ~ ---POWER LIMITED LEVEL TURN TEMPERATURE LIMITATION LIMITED BY MAX. ELEVATOR HINGE MOMENT = + 60,000 LB. FT. /SIDE o i l I I .4 ,6 -1 -2 -3 .8 MACH NO 1.0 1.2 LIMITED BY MAX. ELEVATOR HINGE MOMENT = - 60,000 LB. FT. / SIDE .. < 1111 0 ft .. z .. a .. ,. < 1111 'II Si C. G. AT .29 M.A. C. L 1 i 1 . l l . l l + I 10 - COMBAT WEIGHT =45,000 LB . a . .. .... 7 ., + " MAX. ELEVATOR DEFLECTION TEMPERATURE LIMITATION ,. ,, Ill z ::a 6 f t 'II t l MAX. ELEVATOR HINGE MOMENT ,. 0 ,- " ::a I :I 5 ~ + t + + + + + t ~ + ,. I ,. z " 4 Ill 'II ~ .. - "' a 2 3 z ;,~ \ ... + Ill I z ,. t < POWER LIMITED Ill LEVEL TURN .. MAX. THRUST WITH AFTERBURNER 0 ,, Ill ,. 0 ... I ~ t .5 1.0 1.5 ~ 2.0 .. MACH NO 0 .. -1 0 -2 ~ ~ I /i-MA>t aEVATO• HONGE MOMEN• 0 l 0 I ... r /\. t ... I J . I -31 ' ---- .,. ____ ____ ..;.. _ _, I j "' u, A VRO CANADA 3.1. 3 Weights: Following is a list of the component weights of this airplane. TABLE 8 - WEIGHT STATEMENT Item Weight Percent of Lb, Gross Weight STRUCTURE GROUP 18,366 35. 97% POWER PLANT GROUP 11,616 22. 76% FIXED EQUIPMENT GROUP 6,695 13. 12% WEIGHT EMPTY 36, 677 71.85% CREW, RESIDUAL FUEL, OIL 695 1.36% FUEL FOR 200 NAUTICAL MILES 12,620 24. 72% RADIUS HIGH SPEED MISSION ARMAMENT (EXPENDABLE} 1,056 2. 07o/o USEFUL LOAD 14,371 28.15% NORMAL TAKE-OFF GROSS WEIGHT 51, 048 100.00% TABLE 9 - STRUCTURE WEIGHT Item Weight Lb, WING GROUP: 9,560 Inne r Wing 5,924 Outer Wing 2,948 Elevators 436 Ailerons 252 TAIL GROUP: 841 Fin 690 Rudder 151 BODY GROUP: 5,860 Front Fuselage 4,500 R e ar Fuse lage 1,045 Detachable Rear Fairing 315 LANDING GEAR: 2, 105 Main Undercarriage 1,780 Nose Undercarriage 325 TOTAL STRUCTURE WEIGHT 18,366 27 AV R 0 CANADA TABLE 12 - USEFUL LOAD Item Weight Lb, Crew 430 Oil 40 Residual Fuel' 225 Fuel for 200 Nautical Miles Radius High Speed Mission 12,620 Armament (Expendable) -- 8 Guided Missiles 1,056 TOTAL USEFUL LOAD ,,. . 14,371 TABLE 13 - UNIT WEIGHTS Component Unit Weight Wing Group (Gross Area 1225 sq. ft,) 7, 81 lb, per sq.ft, Tail Group (Gross Area 138 sq.ft.) 6. 09 lb, per sq, ft, Weight of Fuel System (2692 Imp. Gall. of fuel) . 2 7 lb, per Imp. Gall. TABLE 14 - DESIGN INFORMATION Length - max, Height - max, . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Span ..•.....•.......................•.•....... Thickness - root chord .......................... Thickness - tip chord . , ..... , ..•.••••••• • •••.•••. Wing area - net ..•••............•..•.... , , , . , . • Taper ratio ........•••••••.....•. , • , , • • • • • • • • • • Length - root chord ........••..•. • .•..... • • • • • • Length - tip chord ......•.............•.••• • • • • • Maximum fuselage depth ...........• • ..••••••. • . • Maximum fuselage width .................••••.•.• Normal take-off gross weight ... , .• , .••• • .••.•.•• Stressing weight and load factor: At combat weight ..........••••.•• • • • • • • • • • • · .,, • Ultimate load fa c tor ...•.•••.•.••......• • • •. • • Lirni t load factor ..•..•.•.•••• • . •. • • • • • • • • • • • • Factor of safety ...•••.••.••• , , • .. • • • • • • • • • • • 73 ft . 20 ft. 50 ft. 3. 5% 3.8% 766. 4 sq. ft, . 08889 45 ft. 4 ft. 5 ft. 9 ft. 51, 000 lb, * 45,000 lb.* 10,00g.* 7. 33 g.* 1. 364 * 1. 65 in, 9. 5 in . 0 in, 0 in, 0 in. 1 o. 5 in. 9 in. * These values are the current estimated weight and corresponding R. C, A, F. specification load factors. For structural design purposes a design gross weight of 55,000 lb. and a combat stressing weight of 47, 000 lb. are actually used with corresponding load factors of 10, 00 g. ultimate and 7. 33 g. limit. Consequently at a stressing weight of 45, 000 lb, an ultimate load factor of 10, 45 g. and a factor of safety of 1, 43 can be realised. 29 AV R 0 C A N A D A 3.1.3.1 Gross weight estimations are as follows:- Normal take-off g ross weight , ........ . 51, 000 lb. 3.1. 4 Center of Gravity: 3, 1. 5 3, 1. 6 The aircraft will be balanced with C. G. limits between 27% and 31% of the mean aerodynamic chor e. . Ar e as: Vvi.ng area, total including ailerons and elevators Elevator area aft of hing e line (each) ..... , ..• Aileron a rea aft of hinge lin e (ea c h) ........ . Vertical tail area, total ........ , , .. , , .... , . Fin - to rudder hinge ......•..•...•. , , . , , . Rudder, aft of hinge ........ ••·••••• · •·•• Speed brakes: Gross area (each) .. , , • , .• , , ••. , .....••.. , NOTE: The control surfaces are power-operated and do not incorp orate aerodynamic balance aids. Dimensions and General Data: Wings: Span, maximum . , ....... , ..... •. • • • • • • • • • Chord: At root ...... . ....... .,, . , . , .•• :a ••• , • • • •• At construction tip (theoreti c al extended 5 e C t ion at tip) 1 ♦ 1 1 • 1 • t t I a • I • a I II II O I • a • tl I • M ean aerodynamic ... " ........ , .. ,., •• • Airfoil section desi gnation and thickness A t root ..............••• •., • • • • • • • • • • • • • • At tip ..................... , ...•••••••• I • Incidenc e : At root••··•••·•••••·• • •••·••·•••••····• At construction tip , ..... , , • • • • • • • • • • • • • • • Sweepbac.k at 25% c hord .... , ..• , ••• , , •. • •. • • Anhedral .......... , . , • , • • • • • • • • • • • • • • • • • • Aspect ratio . , ........ , • •, • • • • • • • • • • • • • • • • • 1225 sq. ft. 53,39 sq, ft. 33.25 sq, ft. 138 sq, ft. 99.4 sq, ft, 38, 6 sq. ft. 8. 85 sq, ft. 50 ft. 0 in. 45 ft. 0 in. 4 ft, 0 in. 30 ft, Z, 612 in. NACA0003. 5-6-3. 7 Mod, Camber , 0075 Mod, NACA 0003, 8-6-3, 7 Mod, Camber . 0075 Mod, 0 deg. 0 deg. 55 deg, 0 min. 4 deg. 0 min. 2,04 30 AV R 0 C A N A D A Ailerons: Span ( each) •.••..••..... • • • • • •.•.• • • • • •. Chord (average percent wing chord} ..•.•••. Elevators: Span (each) ..•.............• • • • .. • •.. • •• Chord (average percent wing chord} ••.•.•.• Speed brakes (fuselage - lower) Span (each) ••••••••••. •. •. • • •. •• • •.••••. Chord ( each) •••••• • ••• • .• • •. • • •. • • • . •. • • Tail - vertical: Airfoil section designation and thickness At root .......•..............•......... At tip . . . . . . . . . ...................... . Sweep of leading edge ..•.•...••.••••••.. Aspect ratio ................. • ....••... Height over highest fixed part of airplane - fin Height - wing tip ...• • • • • • . • • • • • ••. • • • • •••• Height to top of cockpit •••..•••..•..•••.••• NOTE: The above heights are taken with the airplane in its n@rmal ground attitude - shock - absorber struts static, Length, maximum: Reference line level ..................... Distance from wing m.a, c, quarter chord point to vertical tail m. a. c, quarter chord point .. •• Angle between reference line and wing zero lift line .. •• ...... • ... • ... • .. • • • • • • • • • • • • Ground angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wheel rim size: Main wheels Nose wheels Tire size: Main wheels Nose wheels ............................ ............................ ............................ ............................ Tread of main wheels ...........••.••••••.. Wheel base ...........••• • • • • • • • • • • • • • • • • • 10 ft. 0 in. 27.2% 10 ft. 2 in. 18. 2 % 2 ft. 2 in, 4 ft. 1 in, NACA 0003. 5 -6-3. 7 Mod. NACA 0003. 8 -6-3. 7 Mod. 60.12 deg. 1.043 20 ft. 9. 5 in. 7 ft. 3 in. 14 ft. 3 in, 73 ft. 1. 65 in. 15 ft. 10 in. 0 deg. 4 deg . 15 in. 8 in. 29"x7.7" 1811 X 5. 511 26 ft. 6 in. 29 ft. 11. 85 in. 31 A Y R 0 C A N A D A Vertical travel of axle from extended to fully compressed position: Main wheels ... • .••...•.... • ......•.• • •. 12 in, 10 in, Nose wheels ............................ 3, l, 7 Control Surface and Corresponding Control Movements: Following is a table of control surface and control movements on each side of neutral position for full movement as limited by stops: - TABLE 15 Surfaces Control Movement Rudder Surface 30 deg. RIGHT, 30 deg. LEFT - Pedals According to USAF Spee. AR DCM 80-1 Ailerons Surface 19 deg. UP 19 deg. DOWN - Stick According to USAF Spee. AR DCM 80-1 Elevators Surface 30 deg. UP 20 deg. DOWN - Stick According to USAF Spee. AR DCM 80-1 Speed Brakes: Surface 6 0 deg. maximum DOWN. 3,2 General: 3, 2, l General Interior Arrangement: Refer to Figs, l and 19, These illustra tions show the disposition of crew, armament, power plant, fuel and the avionics equipment superimposed on a general arrangement drawing and a perspective draw ing respectively. Additional illustrations are included in the text to show the main items of equipment in detail together with the access to same for purposes of main tenance, These comprise:- Fig, 27 Fig, 28 Fig, 37 & 38 Structure Arrangement Engine Installation and Removal Armament and Electronic Equipment Installations Special attention has been given to ease of access for those installations which :equire frequent servicing. This aspect is enlarged upon in the paragraphs deal ing with equipment in detail. 3. 2. 2 3, 2, 3 Materials: These will conform to specifications approved by the RCAF. Workmanship: This will conform to the usual high grade airplane practice. 32 AV R 0 CANADA FIG. 19 EQUIPMENT ARRANGEMENT SECRET A V R 0 CANADA FIG. 20 CF-105 COMPONENT BREAKDOW~ SECRET A Y R 0 C A N A D A 3.2.4 Production, Maintenance and Repair: The design of the aircraft is such as will ensure ease of production, simple and rapid installation of the engines and equipment, and ease of general maintenance. Special attention has been given t o the ease with which component parts of the structure can be inspected, main t ained and repaired. The fuselage and wings are designed to facilitate the removal and replacement of damaged sections as illustrated in the component breakdown drawing Fig. 20. The fuselage is constructed in six main sections. These com pr ise: - (a) The radome and probe section. (b) The nose electronics section. (c) The front fuselage section containing the cockpit, intakes, fuselage fuel tanks, armament and equipment. (d) The fuselage center section containing the speed brakes and equipment. (e) The r~ar fuselage section containing the engines and after • burners. (f) The tail cone section. Item (c) forms the largest fuselage section and is approximately 30, 5 feet long by 9. 75 feet wide by 7. 2 feet deep. The wing structure, which is continuous across t he fuselage, is divided into ten separate parts. These comprise:- (a) Left and right inner wings joined at the fuselage center line (b) Left and right inner wing trailing edges ( c) Left and right inner wing leading edges (d) Left and right outer wings (e) Left and right outer wing leading edges The elevators and ailerons are separately hinged to the trailing edge assemblies of the inner wings and outer wings respectively. The dimen·sions of the largest wing panel are approximately 14 feet by 31. 5 feet, NOTE; These dimensions will allow transport by road or rail. Bolted joints connect the fin to the wing structure, This application also applies to the attachment of the main and nose undercarriage units. 3 , z. 5 Interchangeability and Replaceability: Effective on the first prototype air- A V R 0 CANADA plaHc, the component parts· of all airplanes of the same model will be interchange able or replaceable in accordance with and to the extent required by the RCAF. 3,2,6 Finish: The finish of the airplane and parts will be in accordance with RCAF requirements as stipulated in CAP 479. 3 , 2, 7 Identification and Marking: The air plane and its components will be identi fied and otherwise marked in accordance with RCAF requirements. 3 , 2, 8 Extreme Temperature Operation: The air plane as a whole, including equipment, will be so constructed that it will function satisfactorily in any or all temperature conditions that will be encountered, A ground temperature range of:- -65 degrees Fahrenheit to +160 degrees Fahrenheit has been established as the range in which the airplane will operate, 3 ,2,9 Climatic Requirements: The airplane and its equipment will not be adversely affected by other climatic conditions incident to the temperature range stated in the preceding sub-paragraph, and will be capable of transfer from one climate to another without penalty of extensive modification and adjustment, 3,2, 10 Lubri ·· ation: Lubrication of the airplane will conform to the requirements of the RCAF. 3, 2. 11 Standard Parts: Standard parts used in the construction of the airplane and contractor furnished equipment as inc orporated in the various systems in stalled in the airplane will conform to RCAF requirements as laid down in CAP 479. 3. 2. 12 Crew: The crew will consist of the following:- One (l) Pilot and One (1) Airborne interception radar operator 3,2.13 Equipment lnstallations: Equipment to be carried in the airplane is listed below. 3• 2, 13.1 Armament: Guided m.issil e s (Hughes 'Falcon') or . . . . . . . . . 8 off Guided missiles (Douglas 'Sparrow 2 ') •••••••• , , 3 off Further information on this e quipment is contained in paragraph 3.18. 3 , 2.13.2 Communications and Navigation Equipment: This consists of the Hughes 36 AV R 0 C A N A D A MX.1179 integrated navigation - communication - inte rception s y stem, operating details of which ar e des c rib ed in the follo wing sub-paragraphs. lt is anticipated that this digital computing MX.1179 syste m will not be available for fitment to early CF-105 aircraft, and, as an inte rim measure, the two-man analog comput ing MG-3 system wi ll be fitt e d. The MG-3 system has not the capacity for memory storage and consequent automati c operation which is one of the features of the MX.1179 system, but it is the mos t advanced system likely to be available for fitment to early CF-105 air p lanes. 3.2.13.2,1 Operation: Due to the high speed of the airplane, it is consider e d essential to fit automatic inte rception and navigation equipment to the type spec ified in the preceding sub-paragraph. An additional function of the equipment is armament fire c ontrol. D e tails of this latte r procedure ar e outlined in paragraph 3. 17. 3, 2, 13, 2, 2 lnte rc eption navi gation is eff e ct e d by ref e rring the G. C. I, broadcast data to the airplane position as d e te rmined by an A. P. I., corrected for wind by the use of DME omni range. All of the .fore goin g data are fed, as pulses, through a single, high speed digital c omp uter which, through magnetically stored instruc tions processes all the relevant informa tion, and sends the appropriate instructions to the auto-pilot which c ontrols the flight path. A gyro-stabilized platform, which gives a verti c al and north reference , is used as a datum for the steering instruc tions. 3. 2. 13. 2. 3 After the interc eption has b e en complet e d automatically, the computer gives the necessary instructions to re a ch a m a rshalling area defined by its co ordinates in the memory of the c ompute r. From this area, the airplane is directed to the landing slot by a signal sent by the ground controller. This can be relayed eithe r through the pilot or by automatic m e ans. Landing will be accomplished by AILS or AGCA, the choice b e ing dependent on which system is first developed to a high state of perfection. NOTE: It is assume d that the AILS will m ee t this c riterion but w ill be dis p l a c e d, ultimately, by a supe rior AGCA system . 3, 2 . 13, 2. 4 Several r e dundanc ies occur in the na vig ation system, These will permit it to fun c tion with re a sonable efficiency in the event of a number of contin gencies. lf the ground data link is sev ered, the wind vector computed from the last DME omni rang e information is retained in the memory of the computer and applied, from the severage point onwards, to the API without change, The error involved by this method is often quite small, The path of the target may also be computed from GCI data stor e d m agneti cally in the event of failure of this link: reasonabl e accuracy is ensured providing that no evasive measur e s are attempted by the target, ADF is also used as an auxiliary source of information and may be compared with the data se c ured from the other sources to check their veracity. Alte rnatively, it may be used in the event of their failure, 37 AV R 0 CANADA 3, Z, 13, z. 5 All navigational information is displayed to the pilot in a convenient for m s o that he may take over any part of the system which becomes unserviceable, En emy interference with the ground data link can be avoided if the pilot changes the wave length. A range of approximately twenty frequencies is provided f or this pur pose. Any other remedial action, which could be taken by the pilot of a convention al a irplane, can also be applied by the pilot of this airplane during an emergency. For this purpose a manual over-ride of all controls is provided, guidance being ob tained from the information displayed to the pilot from various sources. 3, Z.13,2. 6 Under normal circumstances, the only flight functions expected from the pilot are:- (a) Taxying (b) Take-off until terrain clearance has been secured, and -{c) Stopping the engines after taxying to the ramp. Landing is accomplished automatically. 3, Z, 14 Equipment and Furnishings: Following is a list of equipment and furnish ings provided in the airplane: - Item Two turbo-jet engines and afterburners ••• Instruments ..• • .....•.. • , • • • • • • • • • Hydraulic and pneumatic equipment , .... , Electrical equipment • •. , •• •• •••• • • • • • Electronic equipment .. • •••• , • ••• • • • • Armament . • • . • • • . • •••• • • • • • • • • Furnishings . • . • • • .... " • • • • • • • Air conditioning and anti-icing equipment Auxiliary gear .• • ••. •• ..... • • • • • • • • • • 3, 3 Aerodynamics: Reference Para. 3.12 3.14 3.15 3.16 3,17 3,18 3,19 3,20 3.22 3.3.1 General: As the mission of this airplane requires the use of supersonic speeds, the basic c·onfiguration is so designed as to achieve the _best p~s_sible aero dynamic characteristics in transonic and supersonic flight. This condition can be obtained by using the maximum wing sweep and the minimum t/c. Detailed stud- ies have shown that with conventional planforms, it is practically impossible to find accom_modatio~ in a wing of less than 6% t/c for the necessary equi~":e_nt. In addition to this, the weight of the wing structure becomes excess~ve. ~tihzl.Ilg a delta.configuration however a 3. 5% t/c wing is perfectly practical with a 60 deg rees sweep at the 'leading edge. Furthermore, due to the favourable _disposition of m_aterial , the weight per square foot does not exceed that of a conventional unswept wmg. These considerations made the selection of the delta planform a mandatory 38 A V R 0 CANADA requirement. Turning to the tail unit, it has been found that, with the delta config uration, a horizontal tail becomes not only a superfluity but also an embarrassment. The reasons are as follows: - (a) If a tail is raised substantially above the wing chord line it becomes destabilizing at moderate angles of incidence, thus giving rise to longitudinal instabil ity. This, apart from being dangerous, restricts the airplane 's manoeuverability seriously. (b) If the tail is near the chord plane, these difficulties are overcome but it becomes impossible to secure the necessary ground angle for the low aspect ratio wing. The wake from the wing is also a serious problem in this case. Accordingly then, the supersonic airpl_ane is made tail-less. Troubles with long itudinal damping which have been associated with this configuration are avoided by using at least 60 degrees sweep. 3,3.1.1 Subsonic Profile Drag Synthesis: The drag used in the subsonic perform ance estimates was synthesized by adding up the drags for the various parts ioclud ing appropriate allowances for roughness and interference. From the result, the drag coefficient of the wetted area (or, to use the technical appellation 'aero dynamic cleanness 1) was computed. As a further criterion, the ratio of the total drag to the flat plate friction drag was also calculated. This is known as the 'cleanness ratio'. These two quantities were compared with the values achieved by several airplanes in service; the comparison forms a check on the validity of the original estimate. Following are details of the synthesis:- Note: The figures in parentheses in the following text refer to the list of ref- erences in sub-paragraph 3. 3.1. 5. 1. Body - including canopy and side inlets : Due to the jet location in the tail, the diameter-length ratio of the fuselage (d/ 1) is effectively reduced by approximate ly 40%. Then, based on the wetted area, the drag for a streamline body, with transition at the nose . for cruise flight conditi;ns, is found (1 & 2). An add itional drag increment of 25% is adde d for roughness and leakage. The drag is then referred to the gross wing area. Wing: It has been shown that subsonic drag is independent of the angle of sweep back (3 & 4) providing that the drag is based on the streamwise thickness chord (t/c) and thickness distribution. On tnis basis, the exposed wing drag is det er mined (1). An additional drag increment, due to flush rivets, is taken as . 001 (5 & 6). The wing drag is then referred to the gross wing area. !ertical Tail: Data for the vertical tail drag is found in the same way as t ha t for the wing drag. AV R 0 CANADA Interferenc e : The drag of wing + interference has been found from rocket model tests (7) on models similar in configuration, differences occurring prim arily in body cross-section area to wing gross area ratio. These, however, bracket the ratio r equired so that a reasonable estimate of the interference can be obtained. Miscellaneous: The 25% drag incr e ment already added to the body and the allowance for rivets on the wing and tail are only for normal conditions. It is thought that the armament section doors and the air conditioning system will give rise to more drag. The component drags stated ar e enumerated below together with their percentage of the total zero lift drag. Part CD 0 % Total Wing • 00389 44. 8 Body - including canopy and side air intakes . 0030 6 35. 1 Vertical tail • 00081 9. 3 Armame nt • 00015 1. 7 lnte rferenc e . 00079 9. I TOTAL: • 0087 100. 0 The aerodynamic cleanness and cleanne ss ratio of s everal jet airplanes - including the CF-105 - are given in th e follow ing list as a check on the low speed profile drag estima te. lt should be not e d that the smooth flat plate friction drag used to obtain the cl eanness ratio is varied in a c c ordanc e with the a veraged Reynolds number of each airplane. Airplane Aerodynamic Cleanne ss Cleanness Ratio CF-100 • 003 67 1. 50 F80 • 00299 1. 20 F86 • 0032 6 1. 30 Mete or 4 • 00380 1. 52 CF-105 . 0 0311 1. 41 !he a e rodynamic cle ann e ss r equired for the CF-105 is r e latively quite _g ood. This is not as difficult to attain as for the othe r airplanes due to the much higher Rey nolds numbers of the components. The true situation is mor e clearly illust_rated by the cleanness ratio where this is allow ed for; h ere the CF-105 lies approximately half way betw e en t~e F86 and the CF-100. Whil e the F80 sets a very high standard of surface finish, the 'F86 is not, by c omparison, remarkable in this respect. SECRET AV R 0 CANADA Accordingly, it is felt that a small improvement over t h e CF-100 is quite practical and does not represent any undue optimism as to what may be achieved by a cer tain amount of attention to surface finish. _ 3. 3, 1. 2 Supersonic Profile Drag Synthesis: (Refer to Fig. 21) The supersonic drag estimation was based on available experimental data and theoretical cal culations, from which a detailed analysis, much as in the subsonic case, was evolved. CF-105 wind tunnel results are not given due to the difficulty in cor rec ting profile drag for a ducted model configuration. However, on the basis of present corrections there appears to be good agreement with the estimates, F uselage: The fuselage drag was estimated on the basis of a parabolic body of equal frontal area. That this approximation is reasonable for ducted bodies hav ing near unity mass flow through their intakes is shown in Refs. 8 and 9. (When the flow is less than unity, the incr:ease in drag is handled through the spillage drag term, see paragraph 3. 3.1. 5). For the equivalent parabolic body, the m ax imum cross-sectional area was assumed at mid-length and the fineness ratio taken equal to 8. From Refs. 7, 10, 11, the drag coefficients of similar bodies based on their frontal area were determined and from these the CF-105 body drag coefficient was estimated. Base drag was subtracted due to the effective elimin ation of the base by the jet stream; and a further correction was applied to take into account the ratio of base area to maximum frontal area. The fuselage wave drag was also calculated theoretically using refs. 12 and 13. Adding the wave drag to the estimated skin friction drag (14) gives good agree ment with the value based on experimental data. Canopy: The drag of the canopy was estimated from data on 'V' -shaped and normal-front canopies on Free Flight Models (15 & 16) and then corrected for Fineness Ratio (taking into account the dorsal fairing which extends back to the vertical fin). Wing: Available wind-tunnel data (7) brackets the CF-105 maximum body cross sectional area to gross wing area ratio almost exactly and from this an estimate of wing plus interference drag was made. The model with the larger frontal to wing area ~xperienced a drag rise over the subsonic value and the model 'Yith the smaller ratio, a drag decrease. The CF-105 ratio being midway, it could then be assumed to have a constant drag. However, on the basis of theoretical calculations(l7) and Free Flight Tests (18), the supersonic wing plus interference drag was increased 38% over the sub sonic value for a wing of the CF-105 thickness to chord ratio, sweepback and aspect ratio. The estimate is therefore felt to be quite conservative. ~ertical Tail: Being of the same thickness and sweepback, the fin was treated in the same manner as the wing. "II -Q . Ila -,. :Ill 0 "II -.. ffl a :Ill .. Q " 0 ffl "II "II - " -ffl z ... CD MIN vs MACH NUMBER .020 .016 .012 Co M1N .0087 r .004 J. 67 SPILLAGE AT MILITARY RATING ABOVE TROPOPAUSE MISCELLANEOUS AND INTERFERENCE I VERTICAL TAIL WING CANOPY BASIC BODY o,4-------,.------""T"-----r------"""T"-----.,....----""'T------,------ A ~ J W U U U U M MACH NUMBER ,. < ,a 0 " .. z .. C, .. AV R 0 CANADA Miscellaneous: From Reference 19, it may be seen that a fair deg r ee of wavi ness and pitting has a negligible effect on the supersonic drag. The intP-rfer ence of the wing and fin is already covered in their respective sections . This extra drag then provides for rivets (20), armament doors, leakage and air conditioning system effects. 3. 3. 1. 3 Drag Efficiency: By means of tests on a • 03 scale model of the CF-105 in the Cornell Transonic Wind Tunnel (21), the drag efficiency "e" (Fig. 22) was deter m ined up to Mach 1. 23. References 22-25 were used to extrapolate the curve to higher supersonic speeds. 3. 3. 1. 4 Elevator Drag: Again by means of Wind Tunnel Tests on a scale model, the effect of elevator deflection on the minimum drag CDMIN/82 (Fig. 23) and the lift minimum drag CLCDMIN (Fig. 24) we re obtained up to a Mach number of 1. 23, References 26 - 28 were used to determine the elevator effect at higher Mach N um bers. 3,3,1. 5 Spillage Drag: This has been calculated using an estimate based on Ref e r ences 29 and 30. Due to the presence of a mass flow bypass, the spilL.1ge drag has been reduced to a relatively small value except at very high Mach Numbers and low engine r.p.m. The spillage drag correction is applied as a decrease in engine thrust rather than an increase in drag as shown in Fig. 21, since it arises from ·and depends on the mass flow which varies with altitude below the tropopause. 3,3,1.5.1 List of References: 1. Royal Aeronautical Society Data Sheets. 2. Meleus, H. 3. Maskell, E. C. 4, Whitcomb, R. T. 5• Young, Serry and Morris "Results of Drag Measurements of Fuselage Models in the High-Speed Tunnel of the DVL" MAP/VG R &: T 857. "High Speed Tunnel Tests on a Series of 14% Thick Wings of Varying Sweepback and Taper Ratio, 11 RAE Aero 2295. 11 An Investigation of the Effects of Sweep on the Characteristics of a High Aspect Ratio Wing in the Langley 8-Foot High-Speed Tunnel" NACA RM L6J0la. "Flight Tests on the Effect of Surface Finish on Wing Dragt1 ARC R &: M No. 2258. 43 A V R 0 CANADA q "' ·~ + V V) V 0 ~ "' I ~ !2 "' ... :::; II. t II. <( ~ ... > 0 10 <( l .. ... ... 0 z - ""= ~ ... 10 ~ ::, ~ z :c V ~ --t- 9 "! t- "! ., JI~ st FIG.23 EFFECT OF ELEVATOR DEFLECTION ON MINIMUM DRAG COEFFICIENT 0, "' M 0 45 AV RO CANADA I I I I I t -r---+- r--· - -- 0 <'i I r ~ I I c- >- - t I ~ r- r--, - t-----1 ~ 1------1- >- i---+ - , - u I ... - I !; r-- "' i--c------ ... H -i---, '° 'i ---r- i J: H - - ,) I V ' \ ~ c( I r-- 'i I Ii \ .,, - ~ - - > z I r--- ;; I J 0 " J 1 ... ~, V q ---- J /1 ' - . -=,- t I t r- I r---+ -t - I I t I ' "! >- ' >- 1 I I -Q) 00 00 00 ':,, ~ tO ':> ~ ', ' I 0 7 z 0 0 °'! ;; '"' 0 + u ... V FIG. 24 LIFT AT MINIMUM DRAG AV R 0 C A N A D A 6, Day and Schwarzbach 7. Schult, E. D. B. P ierpont and Braden 9, Nichols and Pendley 10, Hart and Katz 11 . Carlson, H, W. 12. 13, l4, Johnson, J.E. l5. Alexander, S. R. ! 6, Purser, P. E. "A Flight Investigation of the Effects of Surface Finish on Wing Profile Drag" Jour, of Aero. Sciences, April, 1946. "Comparison of Large-S cale Flight Measurements of Zero-Lift Drag at Mach Numbers from 0, 9 to 1, 6 of Two Wing-Body Combinations having similar 60 Triangular Wings with N. A. C. A, 65A003 Sections" N,A. C,A. RM L50I22. "Investigation at Transonic Speeds of a Forward Located Underslung Air Inlet on a Body of Revolu tion" N. A, C.A, RM L52Kl7, "Performance of Air Inlets at Transonic and Low Supersonic Speeds" N. A. C, A. RM 1--,52A07. " Flight Investigations at High-Subsonic, Transonic and Supersonic Speeds to Determine Zero-Lift Drag of Fin-Stabilized Bodies of Revolution having Fine ness Ratios of 12, 5, 8. 91 and 6, 04 and Varying Positions of Maxi.~um Diameter" N, A, C. A. RM L9I30. "Preliminary Investigation of the Effects of Body Contouring as Specified by the Transonic Area Rule on the Aerodynamic Characteristics of a D elta Wing Body Combination at Mach Numbers of 1. 41 and 2.01" N,A.C.A, RM L53G03, Theoretical Forebody Wave Drag R. A, E, Tech. Note Aero 1934. Theoretical After body Wave Drag R. A, E, Report Aero 2348. "Skin Friction at Supersonic Speeds" R, A, E. TM Aer-:J 130. "Effect of Windshield Shape of a Pilot's Canopy on the Drag of an N,A,C.A. RM-2 Drag Research Model in Flight at Transonic Speeds" N. A, C, A, RM L8E04. "Effect of a Pilot's Canopy on the Drag of an N. A, C, A, RM-2 Drag Research Model in Flight at Transonic and Supersonic Speeds" N, A, C,A, RM L7I22, 47 AV R 0 CANADA 17. Warren, C. H. 18, Lawrence, T. 19. Jackson, H. H. 20, Hopko, R. N. 21. 22, Heitmeyer and Smith 23, Smith and Heitmeyer 24• Smith and Heitmeyer 25• Hall and Heitmeyer "The Estimation of the Drag of an Aircraft at Supersonic Speeds" R. A. E, TM Aero 132. "Estimation of Swept Wing Drag at Supersonic Speeds - A Preliminary Survey" R. A. E. TN Aero 2223. "Flight Measurements of the Effects of Surface Condition on the Supersonic Drag of Fin-Stabilized Parabolic Bodies of Revolution" N. A, C, A, RM L52B26, "Preliminary Free-Flight Investigation of the Effects of Rivets and Lap Joints on the Drag of Bodies at Zero Lift at Supersonic Mach Numbers to2,l" N,A,C.A, RML52F09. Transonic Wind Tunnel Tests of a . 03 Scale Model of the A. V. Roe C-105 Aircraft in the Cornell Aeronautical Laboratory, Inc, 4 Foot Transonic Wind Tunnel - Report No. AA-891-W-l. "Lift, Drag and Pitching Moment of Low-Aspect Ra.io Wings at Subsonic and Supersonic Speeds - Plane Triangular Wing of Aspect Ratio 2 with N, A. C. A. 0003-63 Section" N, A. C. A. RM A50K24a. "Lift, ·Drag and Pitching Moment of Low-Aspect Ratio Wings at Subsonic and Supersonic Speeds - Plane Triangular Wing of Aspect Ratio 2 with N. A, C,A, 0005-63 Section" N. A. C. A. RM A50K21. "Lift, Drag and Pitching Moment of Low-Aspect Ratio Wings at Subsonic and Su:E£rsonic Speeds - Plane Triangular Wing of Aspect Ratio 2 with N,A,C.A, 0008- 63Section" N,A.C,A. RM A50K20. "Lift, Drag and Pitching Moment of Low-Aspect Ratio Wings at Subsonic and Supersonic Speeds - Twisted and Cambered Triangular Wing.of Aspect Ratio 2 with N.A. C,A. 0003-63 Thickness Dis tribution" N, A. C. A. RM A51E01. 48 AV R 0 CANADA 26. Boyd and Pfyl "Experimental Investigation of Aerodynamically Balanced Trailing-Edge Control -Surfaces on an Aspect Ratio 2 Triangular Wing at Subsonic and Supersonic Speeds" N. A. C.A. RM A52LO4. 27, Boyd, J, W. "Aerodynamic Characteristics of Two-25%-.Area Trailing-Edge Flaps on an Aspect Ratio 2 Wing at Subsonic and Supersonic Speeds" N. A, C. A, RM A52DOlc. 28. Mitcham, Crabill & Stevens "Flight Determination of the Drag and Long itudinal Stability and Control Characteristics of a Rocket-Powered Model of a 6 00 Delta-Wing Air plane from Mach Numbers of O. 75 to 1. 70 11 N.A.C.A. RM L511O4. 29, Fraenkel "The External Drag of Some Pitot-Type Intakes at Supersonic Speeds - Part I" R. A. E. Report Aero 2380. 30, Klein "The Calc ulation of Scoop Drag for a General Configuration in a Supersonic Stream" Douglas SM 13744. 3.3.1.6 Intake System: In order to increase total pressure recovery at high Mach Numbers, a 120, two-dimensional compression ramp has been used. A diverter-type boundary layer bleed of generous dimensions, which also houses the air-conditioning intake, has been installed to increase total pressure recovery at all Mach Numbers. The D-shaped intake with slightly rounded lips diffuses into a circular duct of sufficient size to keep the Mach Number below O. 5 under all conditions. Intake lips designed for high critical Mach Numbers have been employed to ensure satis factory operation at both Sea Level static and Supersonic cruising conditions. A mass flow bypass situated just ahead of the engine is used at Mach Numbers above 1,·5 in order to prevent instability of the shock system and also serves to decrease the spillage drag. 3, 3. 2 Stability and Control: 3, 3, 2. 1 Longitudinal Stability and Control: The airplane will have positive st_atic longitudinal stability for all speeds, loadings and power conditions'. . A push will be required to increase and a pull to decrease the speed for all conditions except when the Mach number is between 1 05 and 1. 15 up to 40,000 ft. and• 90 to 1. 10 for altitude over 40,000 ft. The s\atic margin (measure of static stability) increases 49 AV R 0 CANADA steadily with Mach number until it is approximately four times the low speed value at supersonic speeds. 3, 3, 2. 1. 1 Dynamic Longitudinal Stability: At all speeds and altitudes short period pitching oscillations will be damped. Long period (phugoid) oscillations will be damped by an electronic damper which is included as part of the auto-pilot. 3,3.2.1. 2 Longitudinal cont:i:ol is achieved by plain flap-type elevators which are power operated and fully irreversible. Aerodynamic balancing aids are not fitted to the surfaces, The effectiveness of the controls at:- (a) Sea level is shown in Fig, 16 {b) 40,000 feet is shown in Fig, 17 (c) 50,000 feet is shown in Fig. 18 1t will be seen from the graphs that nearly all potential manoeuverability - as limited by the maximum lift coefficient obtainable at a given Mach number - is realized by the elevators at a cost of 60,000 ft, lb. hinge moment required at each side, 3,3,2. 1.3 Take-off: It has been established that the main parameters involved in take-off characteristics from the control point of view are: - (a) The main landing gear position in relation to the c, g., and {b) The angle between the wing chord and the ground line (air- plane attitude - nose wheel on the ground). Suitable choice of these parameters has allowed the speed required to raise the nose to equal 90% of the take-off speed and has reduced to a minimum the stick movement required between the nose-up and take-off positions to trim in the air near ground level {equivalent to approximately 5 degrees elevator movement), In consequence, the undesirable aspects of 'stick pumping' at take-off and 'jumping off' the runway have been avoided. These are noticeable characteristics of some c~ntemporary delta wing airplanes, For the CF-105, however, a take-off technique similar to orthodox high-speed fighters is envisaged differing only in the rather large ground angle involved, Take-off speed is approximately 137 knots at a for wa rd c. g. of 2 7% m, a, c. 3, 3, 2, 1. 4 Landing: Landing will be accomplished at small incidences using the flying-in technique, A tail parachute, streamed either automatically or by manual control, will be used to achieve braking in order to keep the landing run within reasonable limits. Speed brakes are also used, 3• 3• 2• 1. 5 Balked Landing: Stability difficulties will not arise as long as _the speed is not allowed to fall below 140 knots. At speeds below 140 knots with full power on the airplane will have negative static longitudinal stability similar to that oc · . . curnng on most conventional airplanes. 50 A Y R 0 CANADA 3, 3.2,2 Center of Gravity Limits: Considerable difficulty arises, for an airplane operating through a large Mach number range, in fixing c. g. limits from a long itudinal stability point of view. This diffi c ulty is created by the following:- (a) The extensive aft travel of the aerodynamic center with Mach numbers, and (b) A forward travel with an increasing lift coefficient for mod- erate speeds (up to M =. 7), Aggravation of the situation is caused by the destabilizing effects of power at high lift coefficients, ln normal practice, it is customary - with orthodox subsonic air planes - to provide for a static margin of approximately 5% m. a. c. at low speeds and high lift coefficient. To achieve a similar provision with supersonic airplanes it would mean, in the case of the CF-105, an increase of its supersonic static margin by 40%. At high altitudes and speeds ( because of the low value of aero dynamic damping), elevator effectiveness in executing pull-outs and turns is almost dir ectly proportionate to the static margin. In effect, this means that the pen- alties at high speeds and altitudes - which are the operational requirements for this airplane - would b e either a 40% reduction in pull-out and turning performance or an even higher percentage increase in hinge moments required to maintain the per formance specified in the flight envelope graphs. (Figs. 16 - 18). Additionally, it would mean a higher overall drag due to higher elevator deflection and consequent greater loss of height in high 'g' turns, The whole performance, therefore, as an efficient interceptor airplane, would be affected. lt w'ould appear then, that to aim at the traditional low speed static margin is an unacceptable solution when deal ing with supersonic airplanes. Consequently the aft c. g, limit on this airplane is fixed at 31% m.a.c. which gives:- (a) A static margin of 4% at low speeds for a CL range of O to • 3, and (b) A static margin of 2% at low speeds for a CL range of. 3 to 1. 1. In the following table data are given on static margins at low speeds of other delta airplanes to indicate that other designers have reached similar conclusions. TABLE 16 - COMPARlSON OF NEUTRAL POlNT, AFT C.G. AND STATIC MARGlN FOR DELTA WING AlRPLANES Static Margin Static Margin Aircraft Neutral Point Aft C. G. Limit 0 M=l.S ALTERNATOR ZONE 1 SECRET EJECTOR PUMP OPERATES ONLY WHEN U/C IS DOWN ZONE2 FLIGHT (M:.25 -1.50) AIR COOLING & VENTILATION APPROX. AIR FLOWS ZONE 1 ¼ LB./SEC. TO 2 LB./SEC. ZONE 2 1/4 LB./SEC. TO 3 LB./SEC. ZONE 2 L AIR 4 APPRO)( ENGINE INTAKE ALTEI AIR t APPROX FIG. 29 ENGINE AIRFLOW AND COOL ·+---------.. __ _ ------ ----- ~ ZONE2 50) ILATION TO 2 LB./SEC. TO 3 LB./SEC. ----➔--------- L ENGINE INTAKE STATIC AIR COOLING & VENTILATION APPROX. AIR FLOWS ZONE 1 1 LB./SEC. ZONE 2 3 LB./SEC. ALTERNATOR ZONE 1 FLIGHT (M >1.50) AIR COOLING & VENTILATION APPROX. AIR FLOWS ZONE 1 35 LB./SEC. ZONE 2 3 LB./SEC. G.29 ENGINE AIRFLOW AND COOLING ZONE 2 70 A y R 0 CANADA engine a ir bleed pressurization, simultaneously from the remainder of the l e ft wing tanks and the aft fuselage tank through the left hand flow proportioning unit to maintain the fuel level in the left hand collector tank. The flow proportioning unit ensures that the individual rate of flow from each tributary tank is directly pro portional to it~ capacity and that all tanks empty simultaneously, thereby minimiz ing the e ffect of change of fuel load on aircraft c, g. position. To ensure continuous transfe r of fuel under normal manoeuvre conditions, the inlets to the transfer pipes are located in extreme corners of the tributary tanks and each inlet is protected by a valve which passes fuel only and not air. Fuel is transferred from the right wing tanks and the forward fuselage tank through the right hand flow proportioning unit to the right collector tank in the same manner described above for the left hand trans fer system. ln the event of failure of a booster pump, a by-pass pipe from the collector tank to the discharge side of the pump automatically enables the pressure ized fue l to supply the appropriate engine at a reduced power. ln the event of lateral fuel unbalance between left and right hand systems due to unequal engine power settings or failure of an engine, it is possible for the pilot to: - (a) Direct fuel from the left hand system to either or both,engines. (b) Direct fuel from the right hand system to either or both engine s . 3. 12. 9. 2 Pumps: lt is proposed to drive each booster pump mechanica~ly by · · ·d • ea box Mechanical drive was shafting from the appropriate s1 e accessories g r • · · · t of weight space and reliability chosen as having distinct advantages 1n respec , over electrical, hydraulic, or air turbine drives. ( ) The number O f cells and their respective capacities 3, 12. 9. 3 Tanks Internal : are as follows: - 12 integral wing tanks 2 bladder type fuselage tanks Total internal fue 1 15,675 lbs. 4, 515 lbs. 20, 190 lbs. d t 1· g a single tank of 3, 750 lbs. 3, 12. 9. 4 Tanks (Droppable): lt is propose O 5 _in f th fuse la e The tank will capacity ( at 7 • 5 lbs. /lmp. Gall.), from ~he _under_ s1~e o abl: in fli ~t~ Fuel is • be quickly detachable on the ground and will be J~t tison 1 t \oth right and t . d ·zation simultaneous Y o ransfe rred by engine air blee pressuri . d to transfer before left int e rnal system collector tanks, Drop tank fuel 15 arr;ng~ omplished· _ internal fuel and automatic follow-on of internal fuel trans er is ace • (a) (b) By making the degree of pressurization higher for drop tanks than for internal tanks, / al.r" valves to the transfer pipe inlets By fitting "fuel no in the drop tank. 71 AY R 0 SECRET CANADA LEGEND ENGINE AND AFTERBURNER FEED•·····C:::===::::o RJEL TRANSFER ················ ············ ·····-tc:::===::::1 TANK N0. 1 TANK NO. 2 1. 2. 3. 4. s. 6. 7, 8. 313 GALS. (2,347 LB.) 289 GALS. (2,168 LB.) LEGEND RJEL-NO-AIR AND CHECK VALVE SHUT-OFF VALVE R.OW PROPORTIONING UNIT RJEL BOOSTER PUMP ENGINE RJEL CONNECTION AFTERBURNER FUEL CONNECTION SERVO OPERATED SHUT-OFF VALVE NON RETURN VALVE 3.0" DIA. NOTE: FUEL VOLUMES AND WEIGHTS QUOTED ARE USABLE QUANTITIES TANK NO. 5 L. H. COLLECTOR TANK 150 GALS. (1,125 LB.) FIG. 30 FUEL SYSTEM - DIAGRAMMATIC OF FUEL FEED J TANK NO. 2 289 GALS. (2,168 LB.) NOTE: ; AND WEIGHTS QUOTED ;ABLE QUANTITIES L SYSTEM - DIAGRAMMATIC OF FUEL FEED AND TRANSFER 72 A y R 0 CANADA TANK NO. 3 R. H. TA TANK NO. 1 LEGEND 1. AIR PRESSURE REGULATOR (DIFFERENTIAL) 2. PRESSURE RELIEF VAL VE (DIFFERENTIAL) 3. AIR PRESSURE REGULATOR (ABSOLUTE) 4. LIQUID FUa TRAP S. AUTOMATIC RESTRICTOR 6. NEGATIVE "G" AND LOW LEVEL AIR ADMISSION VALVE 7. AIR RELEASE VALVE B. ,CONNECTION TO AIR CONDITIONING DUCT (DOWNSTREAM OF RAM AIR HEAT EXCHANGER) 9. PRESSURE RELIEF VALVE (ABSOLUTE) SECRET FIG. 31 FUEL SYSTEM - DIAGRAMMATIC OF FUEL PRESSUII TANK NO. 3 L. H. SYSTEM - DIAGRAMMATIC OF FUEL PRESSURIZ:ATION TANK NO. 7 L. H. TANK NO. 8 R. H. TANK NO. 8 L. H. 73 AYR 0 IICR ET CANADA LEGEND REFUEUING SYSTEM DEFIJELLING SYSTEM LEGEND 1. FUEL NO AIR VALVE 2• SHUT-OFF VALVE 3• RJEL PROPORTIONING UNIT RJEL BOOSTER PUMP 4. s. ENGINE RJEL CONNECTION 6• AFTERBURNER RJEL CONNECTION 7• SERVO-OPERATED SHUT-OFF VALVE S. PRESSURE RERJELLI NG V Al VE 9• NON RETURN VALVE 3.0" DIA. lO. AUER CAP· GRAVITY REFUELLING ll. PUMP. GRAVITY· REFUELLING l2. NON RETURN VALVE· GRAVITY REFUELLING COLLECTOR TANK TANK NO. 5 R. H. COLLECTOR TANK TANK NO. 5 L. H. FIG. 32 FUEL SYSTEM - DIAGRAMMATIC OF REFUELLING At ~NK NO. 2 COLLECTOR TANK TANK NO. 5 L. H. , SYSTEM - DI AGRAM MATIC OF REFUELLING AND DEFUELLING 74 A y R 0 CANADA 3, 12. 9, 5 Vent System: Wing tanks are pressurized to an absolute value and do not vent to atmo sphe r e . Fuselage tanks are pressurized to a differential value and outwardly vent through a relief valve to atmosphere. There is no inwards relief of fuselage tanks from atmosphere because of pressurization. The collector tanks incorpor ate valves which vent only accumulated air to atmosphere, subject to the collector t ank pressure being greater than the pressure required to prevent fuel boiling, 3, 12, 9. 6 Piping and Fittings: Refer to Figs. 30, 31 and 32, which show the proposed layout, 3. 12. 9. 7 Valves: Refer to Figs. 30, 31 and 32, which show the main valves required in the fu el system. 3. 12, 9. 8 Strainers and Filters: An adequate number of these will be incorporated in the system. 3, 12, 9. 9 Quantity Gauges and Flowmeters: Quantity gauges are of the capacitance type with fuel contents presented to the pilot on two indicators, one for each of the left hand and right hand fuel systems. Flowmeters will not be fitted. 3, 12, 9. 10 Drainage Provisions: These will be dealt with adequately in the design of th~ s ystem. 3, 12, 9, 11 Fuel Vapour Inertion: Whether an inertion system will be incorporated depends on RCAF requirements in this respect, 3.12, 9, 12 Fuel Evaporation Control: The degree of pressurization for all fuel tanks i s sufficient not only to transfer fuel, but to prevent fuel from boiling. 3. 12. 9. 13 Refuelling Provisions: Ground refuelling will be carried out from a single point pressure refuelling adaptor. Flow valves will be provided in the tanks to automatically shut-off the fuel when the tank is filled to capacity. Over-riding control i s supplied to stop refuelling at any desired partial fuel load, partial refuel ling being effected through the flow proportioning units in order to achieve the correct aircraft c. g. balance by filling each tank to the same proportion of its total capacity. Provision for in-flight refuelling is not incorporated in this aircraft. 3, 12 , 9. 14 Defuelling Provisions: Ground defuelling may be carried out from the same single point adaptor as is used for refuelling. 3· 12 , 10 Water Injection System: This is not fitted on this airplane. 3, 12, 11 Propulsion System Controls: The design of this system is in abeyance until the engine manufacturer has decided upon the type of control to be used for the infinit;ly variable afterburner nozzle, 75 AV R 0 C A N A D A 3, 12. 12 Start i ng System: Two systems are under consideration: (a) A ground supplied air turbine starter (b) A self contained air-fuel starter The final de cision as to which system will be used has not yet been made, 3, 12. 13 P ropeller: Not applicable. 3, 12. 14 Rocket Propulsion System: This is not required on this airplane, 3, 13 Auxiliary Powerplant: This is not required on this airplane, 3, 14 Instruments and Navigational Equipment: 3. 14. 1 Instruments: The instruments which will be fitted in the cockpits comprise the following: - (a) The normal engine and aircraft services instruments. (b) Such flight instruments as are required to satisfactorily fly and.land the airplane in the event of failure of the MX. 11 79 or MG - 3 electronic equipment. 3, 14, 2 Navigational Equipment: This equipment is included in the MG-3 and MX. 1179 electronic systems as described elsewhere in this brochure . 3. 14. 3 Installation: The installa tion of instruments and equipment will comply with RCAF requirements and specifications. 3. 15 Hydraulic, Emergency Air and Pneumatic Systems: 3. 15, 1 Hydraulic and Emergency Air Systems: 3. 15, 1, 1 Hydraulic System: The aircraft flying controls are operated by two separate hydraulic systems, and the utility and armament circuits by a third system, All systems have separate reservoirs and ground charging points. The systems are designed generally in accordance with MIL-H-5440 requirements. 8tandard hydraulic fluid, MIL-0-5606 and standard AN type fittings are used. Pressure lines are of one-eighth hard stainless steel and return lines are aluminum alloy. T he flying control systems and armament operate at a rated pressure of 4ooo P, s, i. and the utility system at a rated pressure of 3000 p. s. i. All pumps are driven from engine driven gearboxes. 76 A y R 0 CANADA 3. [5. I. I. I Flying Control System: 3, [5. I. I. I. 1 No. 1 Flying Control System: (Refer to Fig. 33} The No, 1 flying control system comprises a power circuit consisting o·f two ·variable delivery hydraulic pumps, each mounted on one pad of two dual pad gearboxes, one of which is driven by each engine. The two pumps are supplied by a single airless type pressurized reservoir, the fluid passing through a fuel heat exchanger, located in the suction line to maintain a maximum system temperature of 250°F. Delivery from the pumps passes through check valves and filters into a common pressure line in whic h is located an accumulator and a relief valve. 3, 15, 1, 1. 1. 2 No, 2 Flying Control and Speed Brake System: (Refer to Fig. 33) The No. 2 flying control system has an identical power circuit to the No, 1 system, the pumps being located on the other pads of the dual gearboxes, one pump being driven from each engine. The No, 2 system supplies pressure to the other half of the dual valves and actuators to give half the total hinge moment required, In the event of failure of the No. 1 system pressure, the No. 2 system will be un affected and will operate the controls normally and give up to half the rated output. In addition to the surface control units, the No. 2 system supplies pressure to the speed brake circuit, which consists of two actuators and servo valves which are controlled by a cockpit lever and give infinite positioning range. 3, 15, I. 1, 2 Utility Circuit: The utility power circuit comprises a single variable volume pump, pressurized reservoir and relief valve. The services actuated are the landing gear, armament, brakes and stee.ring. An automatic depressurizing valve permits the pump pressure to by-pass to the reservoir when there is no demand from any service. A reducing valve maintains the maximum pressure in the landing gear circuit at 3, 000 p. s. i. 3.15. 1. I. 2. 1 Landing Gear Circuit: The landing gear circuit actuates the two main landing gear doors, main landing gear legs and nose landing gear leg. A solenoid operated open center 4-way control valve directs fluid to the up or down circuit as selected from the cockpit lever. A II down" selection directs the fluid to the main gear door up lod:s and door jacks. On the final down travel of the two '.11ain door jacks, mechanical valves are opened. These valves are cross-coupled in such a manner that both must be open before pressure fluid is admitted to the main gear up locks and extension portion of the actuators. The main legs are locked down by separate down locks, On the final movement of the main gear, mechanic ally sequenced valves, which block the retraction side of each door actu ator are closed, 11 Up" selection directs fluid from the solenoid control valve to the main ge ar down locks and retraction portions of the main nose actuator. On ~etraction of the legs, the mechanical sequence valves are opened admitting fluid into the retraction portion of the main door actuators. 3• IS, 1, 1. 2. 2 Armament Circuit: In the case of the "Falcon" missile installation, the armament circuit consists of eight missile launching actuators energized by two self-displacing accumulators. Each actuator has a servo control valve integral with 77 &YRO CANADA LEGEND 1. PUMP 9. ELEVATOR ACTUATOR 2. GROUND CONNECTION - SUCTION 10. AILERON ACTUATOR 3. GROUND CONNECTION - PRESSURE 11. RUDDER ACTUATOR 4. FILTER- H.P. 12. ELEVATOR PARALLEL SERVO S. RESERVOIR 13. AILERON PARALLEL SERVO 6. HEAT EXCHANGER 14. DAMPER SERVO 7. ACCUMULATOR lS. SPEED BREAK ACTUATOR 8. PRESSURE REDUONG VALVE 16. PRESSURE RELIEF VALVE ill CHECK VALVE r--, r--, I I I I 1 1 11 9i,-.~ r ---~-, I I I I .... _ _ J r-1-·- -· ·---· •-t.•7 I y• L __ _J L H. ENGINE GEARBOX ' t I L _____ J R. H. ENGINE GEARBOX FLYING CONTROLS NO. 1 SYSTEM FIG. 33 SECRET NOTE THERE IS ONE L. H. AND ONE R. H. GEARBOX THESE ARE SHOWN TWICE TO INDICATE SEGREGATION Of NO. 1 AND NO. 2 SYSTEMS HYDR AULICS SCHEMATIC FLYING CONTRCl NOTE THERE IS ONE L. H. AND ONE R. H. GEARBOX THESE ARE SHOWN TWICE TO INDICATE SEGREGATION OF NO. 1 AND NO. 2 SYSTEMS V'DRAULICS SCHEMATIC FLYING CONTROLS SYSTEM LEGEND PRESSURE LINES SUCTION LINES c==::J RETURN LINES L __ .J r.!-•-· _•...,!__•~1 L_\ ____ J \_ R. H. ENGINE GEARBOX FLYING CONTROLS NO. 2 SYSTEM 78 AV R 0 CANADA it, operated by programming electro-mechanical input. The armament system will be coupled to the main utility system through quick disconnect fittings. 3, 15, 1, 1. 2. 3 _Brake System: Two main power brake control valves are actuated through a linkage system to the pilot's foot pedals. Pressure to the brake valves is fed through a reducing valve to maintain the maximum pressure at 1500 p. s. i. An accumulator supplies fluid for sufficient brake operations for a single stop in the event of failure of the main pressure source, Anti-skid units are located at each wheel. 3.15. 1. 1. 2. 4 Steering and Anti-Shimmy Circuit: The steering cylinder is attached to the nose leg and is spring-centered, acting also as a shimmy damper. Steering is effected through a mechanical linkage, and a control valve and follow-up mechanism on the leg. A solenoid-operated stop valve is energized and supplies pressure to the control valve. This admits fluid to the desired side of the steering cylinder. A relief valve is fitted ;vhich admits fluid from one side of the cylinder to the other when a pressure build-up is caused through excessive turning loads. 3, 15. 1. 2 Emergency Air System: The emergency air system comprises an emergency air bottle and manually operated valve actuated by the cockpit under carriage lever, Air enters the main down lines system through a shuttle valve, and fluid from the retraction side of the circuit is directed to the return lines through a jettison valve, 3, 15, 2 Pneumatic System: 3, 15, 2. 1 Description and Components: A low pressure pneumatic system will be fitted, operating at a maximum pressure of 200 p. s, i. In flight, air will be_ supplied from both main engine compressors. On the ground, the syS t em will operate from an external source of compressed air. Following is a list of th e services operated by the system and their function: (a) Air Conditioning system to: (b) (c) (1) Control cockpit temperature and pressure, (2) Supply cooling air to electronic equipment. (3) Control armament bay temperature, (4) Cool miscellaneous equipment. An al . r turbine motor is fitted to each engine Engine starting: for starting. Fuel tank pressurization: by bleed air. Fuel tank pressure is maintained 80 AY R 0 C A N A D A (d) Miscellaneous: In addition to the aforementioned main systems air will be supplied for various items such as anti "G" valves, canopy seal, etc. The design of ducting and associated fittings will, in general, follow ARDCM 80-1. A schematic o f the system is given in Fig. 35. 3. 15. 2. 2 General Functioning: Air supply for the system is from both engines through non-return valves, so that it may be operated from either or both engines. Compressed ai r from an external source can be connected to the system through a quick disconnect. Air to the various systems is controlled by means of shut-off, pressure reducing and flow control valves, depending on the requirement of the system. The tem perature of air required for air conditioning is reduced by means of an air-to-air he at exchanger and an expansion turbine and proportioned to the various compartments as required. Additional cooling is supplied by evaporating water. 3. 16 Electrical: 3, 16. l De s c ription: All electrical power, except that from the batteries, is generated by t wo 40 k. v. a., 400 cycle alternators. Each alternator is driven by an engine mounted c onstant-speed drive, This arrangement ensures that the alternator power output is of a constant frequency. The composite system supplies various voltages (i. e. 28. 5v., 115v., 200v. and higher) for supply to the items of radar equipment. Direct current is obtained from rectifiers connected to the 26v. a. c. bus, The battery is connected in parallel with the 28v. d, c. bus for initial power requirements . A circuit diagram of the system is shown in Fig. 36. ELECTRICAL SERVICES A LIGHTING Taxi B Downwar d Ident Navi gation Front Cockpi t Rear Cockpit Equipment B ay Eme r gency Cockpit Floodlight INSTRUMENTS Turn and Bank Artifici al Horizon Outside Ai r Temp Ice Wa rni n g, Airframe Heated Pressure Heads Lighting C FLYING CONTROLS Hydraulic Pressure .Flight Control Warning Light Speed Brake Actuation Elevator, Rudder and Aileron,Trim Landing Gear Actuation and Warning Landing Gear Indication Nose Wheel Steering Hydraulic Pressure Warning Light 81 AYR 0 CANADA AIR TURBINE STARTER NON-RETURN VAL VE ,, __ ..,. ----. I I I I I I I I I I I \ QUICK DISCONNECT AIR TURBINE STARTER .... ◄. GROUND SUPPLY SECRET __________ ,. r----------- I I I I I : MAIN ENGINE COMPRESSOR : I I ~------- ; --- ----- I -----J t NON-RETURN VALVE ♦ NON-RETURN VALVE ----------~. r----------- 1 I I I I I I MAIN ENGINE COMPRESSOR I I : ~-- I --~--------- I _______ ., AIR INLET VALVE FUEL TANKS PRESSURIZATION CON' AIR CONDITIONING EQUIPMENT MISC. EQUIPI (ANTI "G" VALVE, CAN• FIG. 35 LOW PRESSURE PNEUMATIC SYSTEM SCHEMA -------"' I I I I RESSOR I I I I _______ J NON-RETURN VALVE ~ON-RETURN VALVE -------i I I I ESSOR I I I I I _______ ., r AIR INLET VALVE 'Ill,.. - FUEL TANKS PRESSURIZATION C ONTROL AIR CONDITIONING EQUIPMENT MISC. EQUIPMENT ~ , ' , MISC. EQUIPMENT (ANTI "G" VALVE, CANOPY SEAL ETC.) 35 LOW PRESSURE PNEUMATIC SYSTEM SCHEMATIC -,. ' ' COCKPITS ' ' ARMAMENT BAY ,,. ,. T 0 FUEL TANKS ELECTRONIC EQUIPMENT 82 LEGEND ( 1) 40 k. v. a. - 115/200v., 400 Cycle a. c. Alternator. (2) Alternator Control Panel. (3) Alternator Transformer Relays. (4) Selenium Rectifiers and Transformers. (5) 24 v. 9-26 A. H. Battery. (6) External Supply Socket - 400 Cycle. (7) 28. 5 v. t 1 v. d, c. Bus Bar. ( 8) Alternator Switches. (9) Electronic Transformer Relay. (10) Electrical Monitor Load Relay. (11) Alternator Regulators. av10 CANADA -"r----r--____J__I I \ I r-------- 5! , _ _____!_.I 0 Ii=------=== -- L __ __l-G ~ LH I CD ©- I I I 1_+ 1, I llltrl, cb !. c-D I I -{i) I I 11 CD ll H. 1-11 1--7-0 cb \ I I I I I L_ I I FIG. 36 ELECTRICS CIRCUIT DIAGRAM =--~==-==-=---==-=========::::y~~-----=-----=--=t_[! ~! .--- {D 9) 1 1_-i----~-, _L-L._._.J -0 -0 -0 --0 ,-----+-I , --1 ,----l-- ~t-__ _J r I ElECTltONIC LOAD VOLTAGE 2 S FREQUENCY • '2 ~ MISC. AND CYCLIC LOADS I • . I -----. ,---4-+-J_ cb . I..__L.ft 1--'-1=-~-=--=--=-TIH[~ ________ ._: I FIG. 36 ELECTRICS CIRCUIT DIAGRAM i IJ AYR 0 C A N A D A D ENGINE INSTR UMENTB Starting and Ignition Engine Relight Low Pressure Tachometer Oil Pressure Warning Light Inlet Oil Temp Turbine Entry Temp Bear ing Temp Engine Anti-Icing Engine De -Icing Engi n e Ice Detection Engine Afterburner and Controls Fir e Detection Fire Extinguisher E FUEL SYSTEM Fue l Crossfeed Control Fuel Contents Indication Low Level Fuel Warning L ight and F, P. U. By-pass Low Pressure Fuel Cocks Exte rnal Fuel Tank Jettison Ground Refuelling and Defuelling F AlR CONDITIONING Air Conditioning System C ockpits Air Conditioning System Armament Bay Cabin Pressure Warning Cockpit Heating G ARMAMENT Missiles H ELECTRONICS Radio and Radar 1 ENGINE Hydraulic Pressure Warning Light Afterburner Emergency Switch Hydraulic Emergency Switch Fuel Control Emergency Warning Light Primary Pump Warning Light Engine Fuel Control Selector J MISCELLANEOUS Windscreen De-lcing and De-Misting Tail Parachute Actuation 3· 16, 2 E l ectrical Power Supply: The electrical load to be supplied by the alte r nators is as follows : - Supply Electronics ..•••• , • • • • • Electrics including fuel pumps ••• Armament and Fire Control Missiles . . . . . . . . . . .. Dema nd 5 k, v. a. 16. 54 k, v, a. 7. 5 k. v. a, 7, 2 k, v. a, In the event of failure of an alternator the remaining one will carry partial load for any lengt h of time. The normal load will be approximately 36. 5 k, v. a, A control panel fo h · • • b h • ·t b k r eac alternator is provided in the cockpit. The ranc circut rea ers are located in the nose wheel well and aft of station 495. Good 1 · d k • • • 1 vo t age regulation is obtained by using compoun feedbac • Prov1s1on 1s a so made in the power circuit for over-voltage protection, feeder protection, reverse current protection, faulty alternator disconnection and load equalization, The emer 1 1 • gency systems are ·the main load on the battery. An externa supp Y 1s provided fo r the a. c. system only. Ground power is obtained by supplying air to 84 AV R 0 CANADA the cooling system of the a ltern ator co ntro l e quipme nt while r e ce iving power from external a lternators. Danger of alternator and equipment overheating i s obviat ed by the circulation of cooling air supplied by the external pneumat ic syst em. 3, 16. 3 Electrical Power Conversion: Some d, c, power is required and is obtained from selenium rectifiers connected to the 28, 5v. a, c, bus, These units are cooled by the air conditioning system. 3. 16. 4 Equipment - Installation: The alternators are installed on the nose bullet of the engine. Removal of the alternators can be accomplished through the intake tunnel. The alt erna tor control panels and rectifier units are l ocated aft of s tation 485, Access is gained to these latter components by an access door on the under side of the fuselage, The battery is located in the nose wheel well giving ease of access for inspectio n and servicing. The alternator switches are lo cated in, and are accessible from, the cockpit. 3, 16, 5 Wiring: Wiring will be fitted in accordance with Spee, MIL-W - 7 139A. 3, 16, 6 Bonding: All flying controls, moving surfaces, electrical pa n el s a nd junction boxes will be bonded to the main structure in accordance with the instructi ons issued in Spee. MIL-B-5087. 3, 16, 7 Controls: Trip free circuit breakers will be used on d, c, cir cuits and some a, c, circuits located in +160oF or lower ambient zones. Fuses wi ll be used on other a, c, circuits. Current limiters will be used in the power distr ibution and branch feeder circuits in ambients above +160°F. 3, 16. 8 Lighting : Lighting will be provided for the following:- (a) Cockpit and instrumentation ( c) Navigation, and ( (d) Downward identification b) Taxi 3,!6,9 Starting and Ignition: (Refer to para, 3. 12. 12). 3 16 10 R t le supplies the 200/115v. a, c , , ' • eceptacles: An external power recep ac . . 400 cyc le 3 phase supply only. Special AN connectors will be ~sed on all ~:~:tion boxes and panels, Provision will be made for fuel line grounding a ::1d a c will be fitted to g round the fuselage structure. 3 6 • t refer to s ub-paragraph • 1 • 11 Indicators: For details of the electrical ind1ca ors 3• 16, 1. - Warning lamps are provided for : - (a) Ice warning (b) Fire detection, and ( c) Fuel pressure warning 85 AV R 0 CANADA 3, 16. 12 Electri cal Drives: Not used in this airplane. 3, 16. 13 Re lays: Relays will be used in the following circuits: (a) High power {b) Alternator control, and (c) Battery 3, 16, 14 Booster Coil: Not used in this airplane. 3, 16. 15 Radio Filters: These items will be incorporated as required. 3. 17 Electronics: 3,l7, 1 Description: Provision is made in the CF-105 to accommodate an integrated interceptor system based on a digital computer, such as the Hughes MX. 1179 equip ment. Such an integrated system would include facilities for automatic navigation , interception, attack, return to base and automatic landing. Becaus e t he time phasing of the CF-105 is earlier than the anticipated availability of the di gital computing system, the electronic installations have been designed to accompli sh essentially the same objective using sub-systems based on analog ~o-~put ing devices, such as the Hughes E-9 and MG-3 system. Considerable flex- 1 ~1ltty in the mode of operation is provided for. The equipment specified for the aircraft will enable operation under : ( a) (b) ( c) Close control, defined as continuous transmission of commands to the interceptor via data link from GCI to vector the interceptor into the most favourable position for attacking the target. Modified close control, defined as intermittent transmission of target position and course vector to the interceptor by GCI, the interception problem being taken care of in the interceptor, Broadcast control, defined as the transmission t o all interceptors in an operating area of target position and course vector. 3,17. Z L ocation of Electronic Equipment: (Refer to Figs. 1, 19, 37 and 38) . electro nic equipment in the CF-105, apart from cockpit operating components , i s located • f . The m our principal regions: (a) Nose Compartment. (b) Electronic Equipment Bay (immediately forward of the Armament Bay). (c) Armament Bay. (d) Dorsal Compartment. 86 AV I 0 CANADA In addition, accelerometers are located in an area very close to the ce nter of gravity, and ant ennas requiring special locations are housed in the most favourable a r eas available. These are described in detail in the paragraphs which follow. The electronic equipment in the aircraft is divided into the following major groups: Telecommunications, F i re Control, Navigation ( other than radio aids}, and Flight Control System. 3.17, 3 Communication Equipment : 3. 17, 3. 1 UHF Communication: An AN/ ARC-34 UHF Transmitter-Receiver is pro vided for voice communication, air-to-air and air-to-ground. The transmitter receiver is located in the electronic equipment bay. Antennas are provided at the top of the vertical fin and in the bottom of tl).e electronic equipment bay in orde r t o provide coverage in all important directions. These antennas will be shared with the UHF data link receiver and it is intended to provide diversity operation so th at both antennas are used continuously. It is anticipated that this diversity operation will be secured by means of ferrite attenuators alternately keyed by means of a high fr equency switching device, the frequency of commutation being h i gh enough t o avoid deg r adation of the received intelligence. AN/ARA-25 UHF homing equipment will be provided in conjunction with AN/ARC-34. This device is discussed further under Ele ctronic Countermeasures, sub-para. 3. 17. 6. :-17, 3. 2 UHF D"ata Link: Either AN/ARR-39 or AN/ARR-43 UHF data link receiv '.ng and de-coding equipment for reception of interception contr o l and automatic GCA information will be accommodated in the radar nose. As discussed under UHF Communication, the two UHF antennas will be used for both communication and data ~ink with commutation between them to provide for continuous contact regardless of intercepto r attitude. 3, 17 , 3. 3 Interphone: AN/ AIC-10 interphone system is provided, giving both crew members access to the UHF communication transmitter and selection of any o r a ll :oice r eceiver outputs. An external interphone connection point is provided, tieing tnwith the interphone function of the AN/AIC-10. 3.17 4 Na • t · . • viga ion Equipment: 3• 1:· 4· 1 L ow F requency Radio Compass : AN/ARN-6 low frequ~ncy r~dio compass equipment is p rovided, the receiver being located in the electronic equipment bay. Both co ckpits have control facilities tuning being done remotely by means of AN/ ARA-19 . . ' A fl h f ·te loop being develope d by . remote tuning servo equipment. us erri • . B~ndix Radio for the AN/ ARN-6 from the commercial flush radio compass loop, Will be installed in the underside of the electronic equipment bay. A large flu sh panel over the loop will accommodate the sense antenna in close proximit y to the receiver, the r e sulting reduction in sense a ntenna lead capacity being employe d to offset the low effective height of this type of antenna. 87 l VI 0 CANADA 3,17,4.2 TACAN: AN/ARN-21 TACAN equipment will be accommodated as an alternative t o the Doppler ground track and ground speed device described be low. The TACAN equipment will be housed in the electronic equipment bay and will use the same antennas as the L-band air-to-ground IFF. These are dealt with in detail under l dent ification, sub-para. 3. 1 7. 5. 3.17. 4. 3 Doppler: The Doppler ground track and ground speed detection equip- ment being developed by the Canadian Defence Research Board will be accommodated in the ele ctronic equipment bay as an alternative to TACAN. A radome sandwich panel in the underside of this bay will be designed for optimum transmission at the frequency of the equipment. The Doppler device will feed data to the navigation and interception computer equipment described later. 3.17.4,4 Radio Altimeter: AN/APN-71 Flareout Radar Altimeter will be installed on the ventral center line of the aircraft, providing altitude information to the cock pit indicator and also to the flight control system. 3.17.4,5 Compass: A J-2 type of flux valve compass is provided. In line with the objective of integrating the gyro requirements for the various electronic systems, it is expected that the directional gyro of this compass will be integrated with and stabilized by the vertical gyro which also serves the fire control system, flight control sy stem and Doppler radar antenna stabilization. 3, 17, 4. 6 Navigation and Interception Computers: The airplane will carry the navigation and interception computer being developed to fill the requirements of the RCAF fo r a dead-reckoning device to fulfil the following functions: (a) Navigation to intercept targets. (b) Return to base course computation. (c) Computation of distance to destination. (d) Collision or pursuit course navigation to a desired point which may be a moving target. ( e) Computation of time to go on the basis of fuel remaining. (f) Indication of ground position, This computer will be capable of drawing information from TA~AN, data link or . Doppler sources and will also be designed to continue computation by dead-re c koning, W~en data inputs are interrupted, on the basis of the last data received. o_utputs W1ll be included to enable the courses computed by the navigation interception com puter to be executed by the flight control system automatically. 3, 1 7 5 Id t· f' • en 1 1cation Equipment: 3, 17. 5. 1 Air-to-Ground IFF: The CF-105 will carry AN/ APX-19 or AN/ APX-25 88 A Y R 0 CANADA Mk X IFF transponder equipment in the electronic equipment bay along with flush antennas in the dorsal and ventral surfaces of the fuselage. The transponder will be commutated between the two antennas. The same antennas will also serve the TACAN system when it is carried. In the case of TACAN, the received signal indi cation (11 Flag Alarm") will be used to switch between antennas in case of signal failure, to provide continuous contact regardless of aircraft attitude. 3.17. 5. 2 Air-to-Air IFF: AN/ APX-26 air-to-air IFF interrogator-responser and AN/APX-27 air-to-air IFF transponder, or their equivalents, will be installed. The AN/APX-26 group will be installed in the radar nose and will be integrated with the radar antenna and waveguide. The transponder antenna, located in the vertical fin, will be connected by waveguide to the equipment group in the dorsal electronic compartment. 3.17. 6 Electronic Countermeasures: 3, 17. 6.1 UHF Homer: As noted under UHF Communication, AN/ARA-25 automatic UHF direction finding equipme::it will be installed in conjunction with AN/ARC-34 to enable the interceptor to home on sources of UHF communication or data link jam mi:ig, In addition, this homing system will be available far use as a navigation and approach aid. The UHF homer is located on the nose wheel door. :• 17, 6. 2 Radar Homer : AN/ ARD-10 broadband radar homer equipment will be '.nstalled in the radar nose to provide facilities for homing on sources of radar illumination or jamming. 3• 17, 7 Fire Control System: The fire control system of the CF-105 is a version of tile Hughes E-9 and MG-3 family. 3• 17, 7.1 Radar Sub-System: The radar sub-system performs automatic search, automatic trac;king, ground mapping and interrogation and reception of radar ground beacons. During automatic tracking the radar sub-system supplies the computing sub-system with the necessary information for computing an attack course againS t the target on which the radar is locked. When using semi-active X-band missiles the radar illuminates the target throughout the missiles' flight so that reflected RF energy is available for missile guidance. 3. 17 7 2 l d l t • • • Computer Sub-System: The computer sub-system inc u es an e ec ron1c analog computer capable of generating lead-collision missile attacks and either lead :::suit or lead-collision rocket attacks. It also schedules the preparatory operations ich rnust be performed on the missiles before they are launched, and continuously :o~putes time remaining until missile firing. Signals are supplied at specific inter as to the missile auxiliaries sub-system to control the sequence of the pre-launch 0 Peratio • d f II ake 11 ns required to set up the missiles. The computer prov1 es or sn operation to enable the interceptor to follow another aircraft accurately and safely rega dl · • d ~ ess of visibility. In this non-attack mode of operation the computer prov1 es steerin . . 8 s1gnals for a lead-pursuit course. 89 AV R 0 CANADA 3.17. 7. 3 Controls and Power Requir e ments : Controls and cockpit equipment are generally s imilar to the E-9 fire control system while the central power supply and missile auxiliary configuration similar to the MG-3 ar e used. The multi-purpose analog computer and 250 kilowatt peak power tunable radar common to both systems are employed. To increase the AI range capabilities, accommodation is provided for a 35 inch diameter antenna swept volume. With conventional gimballing a 30 inch diameter dish can thus be accommodated, while with an antenna gimballed in the vi cinity of its center of gravity, dish sizes approaching the limits of the swept volume may be carried. The radar and computer components of the system are accommodated in racks in the radar nose, with a plenum chamber in the center of the nose compartment sup plying adequate cooling air for all operation environments. The high voltage DC requirements of all parts of the fire control system are sup plied by a multi-output rotating generator driven by a fixed frequency AC motor from the aircraft 400 cycle power supply. Regulators are provided for the various voltage outputs, primarily to limit _the effect of transients and to isolate the major portions of the system from each other with regard to the effects of such transients. 3• 17, 7, 4 Missile Auxiliaries: The missile auxiliaries sub-system consists of central computing equipment and individual channels for the setting up of proper par ameter s in the missiles. With GAR-1 missiles the functions of the missile auxil iaries sub-system in clude : (1) Applicat ion of external power to critical circuits within the missile. (2) Transfer to internal missile power. (3) Se rvoing of the missile antenna into alignment with the radar antenna. (4) Servoing of the radar transmitter and the missile receiver into frequency correspondence. (5) Servoing of the missile range gate into time coincidence with the radar range gate. (6) Computing optimum missile firing range. (7) Computing the missile guidance and control parameters in accord ance with aerodynamic conditions. (8) Blowing of missile attenuator fuses to set the computed para meters into the missiles. Equivalent o , allJcil' . peration for 'Sparrow 2' missiles would be provided by suitable missile 1aries co f , , , n orm1ng to the requirements of the Sparrow. AYR 0 CANADA 3.17, 7. 5_ Ai r Da_ta Computer: The fire control system includes an air data com puter which p rovides, to a high degree of accuracy, the following quantities: (1) Mach number. (2) True airspeed. (3) Dynamic pressure. (4) Missile jump angle, (5) True angle of attack. ( 6) Air density times speed of sound. The air data computer receives inputs from a probe projecting from the apex of fue radome , carrying the pitot-static head, relative wind direction sensors and radome de-ici fl 'd d' • • • • • 1 . ng m 1spenser. This computer 1s provided primarily for weapon aunchmg co t t· • • • . mpu a 10n requirements but also permits weight and space economy by mtegrating • • 1 • . s1m1 ar requirements for the flight control system and certain flight mstruments . 3. 1 7 7 6 L t · f ( 1 • · · oca 10n o Fire Control Equipment: The radar, computers mu ti-purpose attack comp t · d • • • • • ) d t u er, air ata computer and m1ss1le auxiliaries computers an sys em ~~m~uter test units are contained in the radar nose compartment. The missile aux- iliaries a s · d • • h t b soc1ate with the individual missile channels are housed in t e armamen _ay adj_acent to the missiles while power supply equipment and weapon salvo selec- tion units d · • 1 • • an power supply components are accommodated 1n the e ectromc equip- ment bay a d th • 1 h' t· • d . n e general equipment bay just forward of 1t. Weapon aunc ing im1ng evices i 1 d' • 1 • th ' nc u 1ng the intervalometer and weapon extension contro s, are 1n e armament b ay. With the ob · · • h th • Jective of achieving economy of equipment and weig t, e gyro require- ~ents for the fire control system will be integrated as far as possible wi th th ose or the fl' h • 1 t· t th ig t control system the gyros being in a favourable location re a 1ve o e cente f . ' • t t r O gr a v i ty of the aircraft in the dorsal electromc compar men· 3.l77? Q ·1t' k'tf us · · · ptical Sighting: An optical sight is provided in the pi o s coc pt or l e at low altitudes where airborne radar tracking is ineffective due to grou nd cutter. ~-17. 8 F light Control System: The flight control system of the CF-105 will be a eve lopme t f • 1 t Th. 5 • s a rate system n o the Minneapolis-Honeywell E-10 autopt o • 1 1 SECRET 91 l YR 0 CANADA using DC signals and operates through h i gh performance hydraulic servos. The flight control system will provide the following functions: (a) Stabilization and damping of the aircraft. (b) Flight in response to manual commands of course, altitude or Mach number. (c) Automatic flight response to commands of th_e navigation and interception computer. (d) Automatic attacks under the control of the fire control system, (e) Automatic landing under the control of AGCA commands trans mitted by data link, the closing phase of the landing being con- trolled by the flareout altimeter. Manual commands will be inserted by means of stick force sensing using the air craft's conventional flight controls, The rate gyros and accelerometers of the flight control system will be located close to the C. G. in the dorsal electronic c~mpartment, the remainder of the electronic portions being in the radar nose. Air_ data and vertical gyro reference requirements of the flight control system will be mtegrated as far as possible with those of the fire control system and the Doppler antenna stabilization. 3• 18 Armament: 3.I S. l Air- t A" • • • • co . 0 - ir Guided Missiles: (Refer to Fig. 37). The CF-105 armament ns1sts of f H • d . our ughes 'Falcon' radar seeker guided missiles and four infra-re equwalents c · d • • 1 b lnf a rne in two rows of four in the armament bay. The rear four wil e ra-Red . · 1 • • • • lo . m i s si es and the forward four will be radar seeker missiles. Missile wenng is b • Y means of a hydraulically operated parallel link mechamsm. ~::rs. underne ath each missile are mechanically linked to the lowering mechanism will open to · • • 1 h the full permit extension and will close again as the missi e approac es bay d Y extended position. Aerodynamic considerations dictate that the armament oors be closed with the missiles in extended launching position. In order to a . . • • • mis . 1 void homing and aerodynamic interference between missiles, inboard siles 5 o i e angled out 1° 40' from the airplane center line and outboard mis-s1 es w·11 b fr om th · 1 f f" • Al • • 1 angled 4 o do fe airp ane center line at the time o 1r1ng. so, missi es are wn rom the airplane datum to ensure clean separation from fuselage, The • missiles o 1 1 to th . a re stowed 3 nose down relative to the fuselage datum and paral e e airc r aft o 1 durin 1 center line, Rotation to 4 nose down will take place gradual Y the rn~ o_wer ing. Transition to 1 ° 40' or 5° in azimuth will take place as soon as iss1le f" ms are clear of the doors. 92 A Y R 0 CANADA ~------------ ----A.''-6"-- - - - - ---- ------J RADAR UNIT HINGED ELECTRONIC CRATE GYRO UNIT SECTION THROUGH "FALCON GAR I" CONTROL ACTUATOR 8 FALCON GUIDED MISSILES l'-8" · :·.:·:·:'·: ·.>·.·.:-::_:·:: ·_ ........... . HYDRAUI LAUNCH FIG. 37 ARMAMENT AND ELECTRONIC EQUIPMENT INS'.1',t t a ,, CONTROL ACTUATOR •N GUIDED MISSILES HYDRAULICALLY OPERA TED LAUNCHING MECHANISM ARMAMENT DATA • RANGE 5,000-25,000 FT. • MAX. VELOCITY 2,000 FT. PER SECOND GREATER THAN VELOCITY OF LAUNCHING AIRCRAFT • WEIGHT (PRE-LAUNCH) 128 LB. • LAUNCHING ALTITUDE LIMITATIONS 10,000 - 60,000 FT. • TEMPERATURE LIMITATIONS (PRE-LAUNCH) o •F-130.F AND ELECTRONIC EQUIPMENT INS'.l'ALLATIONS - FALCON 93 A y R 0 CANADA To permit serv1c1ng of the missiles and missile auxiliaries, provisions are included for partial lowering of the launchers, the doors being retained fully open. The hydraulic system will consist of a hydraulic jack with an uplock operating each missile launcher extension gear and hydraulic accumulators sufficient to extend all eight m i s siles in one pass. Retraction power is supplied by hydraulic pumps, The extension of each missile is regulated by having the jack follow a programmed input and this input is such that accelerations of the missile during lowering are not excessive. The programmed input is accomplished by a cam in a programming box, whi ch will drive a position error hydraulic servo valve on each lowering jack. Missiles may be fired in a salvo of eight, a salvo of either four radar seekers or four infra - red seekers or a mixed salvo of two radar seekers and two infra-red seekers. In launching mixed salvos the rear missiles will be fired first, their launche r s being left extended while the front missiles are lowered and fired. All empty launchers will then be retracted simultaneously. In the case of an attack with either four radar seeker or four infra-red missiles the appropriate row will be lowered and fired and the empty launchers retracted immediately. An alternative armament of three Douglas 'Sparrow 2' fully active homing missiles can be a ccommodated, stowed in a similar manner to 'Falcon' missiles - refer to Fig. 38. 3.18. 2 Rockets: A further alternative armament installation can be arranged along eithe r of the following lines for conditions which might be unsuitable for missiles, such a s low altitude attacks: (a) (b) An expendable rocket package containing fifteen 2-inch rockets can be fitted in place of each Falcon missile giving a total of 120 rockets. With the launching gear strengthened to allow for increased loads there is sufficient space to accommodate 8 launchers each con taining 25 rockets, or a total of two-hundred 2-inch rockets. 3. 19 Furnishings and Equipment: 3. 19. 1 Accommodations for Personnel: 3. 19 . 1. 1 Crew Seats: The pilot and radar operator will each be provided with an upward ejection seat for ejection under eme~gency conditions. Each seat will be equipped with a safety belt and shoulder strap assembly. • 11 b d at each crew station for 3. 19. 1. 2 Anti 'g' Suit Provisions: Provision w1 e ma e 94 AV R 0 CANADA RADAR UNIT PRO.XIMITY FUZE 12'-11.5" UMBILICAL UNIT CONTROL ACTUATOR SECTION THROUGH "SPARROW 2 ,, MISSILE HINGED ELECTRONIC CRATE a C ; 3 SPARROW 2 GUIDED MISSILES FIG.38 HYDRAULICALLY OPERATED LAUNCHING. MECHANISM ARMAMENT AND ELECTRONIC EQUIPMENT INSTA! l N A D A it'ERTER PRO_XIMITY NIT FUZE NIT SECTION THROUGH "SPARROW 2" MISSILE HYDRAULICALLY OPERATED 2 GUIDED MISSILES LAUNCHING MECHANISM FIG. 38 ARMAMENT AND ELECTRONIC EQUIPMENT INSTALLATIONS • SPARROV OL HYDRAULIC ·oR POWER SUPPL y CAL UPPLY MISSILE HYDRAULICALLY OPERATED LAUNCHING MECHANISM ARMAMENT DATA • RANGE 5,000-33,000 FT. • MAX. VELOCITY 1,700 FT. PER SECOND GREATER t:HAN VELOCITY OF LAUNCHING AIRCRAFT • WEIGHT' (PRE-LAUNCH) 355 LB. • LAUNCHING ALTITUDE LIMITATIONS 10,000-50,000 FT. • TEMPERATURE LIMITATIONS (PRE-LAUNCH) -75°F TO +150°F (FOR ROCKET MOTOR) i, ELECTRONIC EQUIPMENT INSTALLATIONS • SPARROW 2 AV R 0 C A N A D A the use of a service supplied anti 'g' suit The • t ·rul . • air O 1 ate these suits ·11 b supplied by the low pressure pneumatic system. wi e 3.19. 1. 3 Partial Pressure Suit Provision· Provis·o ·11 b . . . • i n w1 e made at each crew station for the use of a service supplied partial pressure suit. 3. 19. 2 Miscellaneous Equipment: Miscellaneous equipment will be fitted as required by the RCAF. 3. 19. 3 Furnishings: A protective anti-skid coating will be applied to the cockpit floor in areas subjected to wear. 3. 19. 4 Erne rgency Equipment: 3. 19. 4. 1 First Aid Equipment: Provision is made in both cockpits for the instal lation of two service supplied shell dressings in a position accessible to the respec tive crew member. 3. 19. 4. 2 Crash Equipment: A crow bar will be provided in each cockpit in a posi tion accessible to the respective crew member. 3.19. 4. 3 Fire Extinguisher : A hand fire extinguisher will be provided in each cockpit in a position accessible to the respective crew member. 3. 19. 5 Oxygen Equipment: A liquid oxygen system designed to the requirements of D. 0. R., A. F. H. Q., D. N. D., will be installed in the airplane. The system will provide the crew with the necessary oxygen for breathing and also provide for the inflation of partial pressure suits. 3.19. 6 Emergency Rescue Equipment: Such equipment as required by the RCAF wi 11 be fi tte d. 3. 20 Air Conditioning and Anti-Icing Equipment: 3. 20. 1 Air Conditioning: Refer to sub-para. 3. 1 5. 2 and Fig. 35, 3. 20. 2 Anti-icing: 3. 20. 2. 1 Anti-icing of ·Non-transparent Areas: The radome and other critical instrument reference areas and the intake ducts will be fitted with anti-icing means. 3. 20. 2. 2 Anti-icing, Defrosting and Defogging of Transparent Areas: The are~s to be protected are the windscreen and the canopy windows. The exact method which will be used is still being investigated. 3. 21 Photographic Equipment: This equipment will not be fitted to this airplane. 96

CF-105: twin engine supersonic all-weather fighter

Archive

AVRO Museum

Name/Title

Entry/Object ID

Category

Web Links and URLs

General Notes

Note