Flightworthy active control landing gear for a supersonic aircraft

TABLE III MINIMUM MARGINS OF SAFETY Min. Mh Loading Part M. S. M. S.

-Material No. Description Condition Yield u1t. Pages :1004818 Support Assy., 321 CRES To determine Max. ext.

Orifice Plate Mlg max external pressure 1921 psi 6- 10 & internal Max. int.

pressure assy. pressure 2323 psi can withstand Tube Assys. 2 l-6-9 CRES :1004829- 101 Pressure(- 10 1) 1.09 0.98 11-14 - i02 & bending(- 104) - 103 Pressure(- 102) - 104 (- 103) 2.26 1.21 11-12 - 105 (- 107) - 106 (- 108) - 107 Pressure(-105) 2.15 1. 14 15-16 - 108 Pressure(- 106) 2.12 1. 11 16- 17 :1004819- 001 15-5 Ph CRES Adapter Pressure 9. 32 5. 62 19 Pressure & Moment (Bolt) 1. 51 20 1.21 20 (Lug) 10048 19- 00 7 Servovalve 606 l-T6 Wt. & G loads Large Large 23 Brackets Alum. Sheet (Tie-down bolts Wt. & G loads 27 Large Large (Bracket bolts) Wt. & G loads (-007) 1. 57(Tens. 28 (-008) .arge(Shea. 28 10048 19- 003 Accumulator 7075-T7351 Wt. & G loads Large Large 32 - 004 Clamps 2i Alum. bar (thru bolt) - 006 Brackets Wt. & G loads 0.39 33-34 (-003) 0.62 10048 19-003 Accumulator 7075-T7351 tit. & G loads 34-35 Alum. bar -004 Clamps & (-004) -006 Brackets Wt. & G loads Large 35-37 (bolts) Wt. & G loads Per 746(Shear) -oo Attachmt.

Nt. & G loads 0.49 39-41

-

It

TABLE IV I i 1: ELECTRONIC FUNCTIONS \ I Electronic Mechanization Mode Determination Flip-flop (U36A) is high or low as a function of scissors switch state Landing Mode Pre-Touchdown Analog switch U27A is closed to complete the pressure loop to maintain the strut at the design charging pressure.

Touchdown Flip-flop (U43A) is set when the wheel generator exceeds a threshold level Pre-set or derived from a sink rate sensor Sink Rate (Vs) Wing/gear acceleration is integrated by U3D Wing/Gear Velocity (V ) wg - VW,) is squared (Ull) Kinetic Energy (VS is multiplied by (Xm - Xs) (U19) Potential Energy F Comparison of Kinetic CZZparator (U13C) and Potential Energy FL1 Held Constant Sample and hold amplifier (U17) is switched to a hold state by U25A During Impact Servoloop Enabled Analog switch (U27A) opens the pressure loop and closes the force loop.

FL1 is squared by U18 Transition Velocity (V,) (V, - V,g) compared to Vt by comparator (U13A) Determination of Transition Rate Decay of FLC During Transition Ramp is generated by UlOD and U6F.

is determined by R45 and Cl3 Comparison of Fwg and F,m Comparator (~32C) FLC Maintained at zero and servoloop Analog switches (U24A and U34A) Disabled When / Fwg/ (Fmin FLC = Fmin Sign Fwg Sample and hold amplifier (Ul7) When /Fwg/ > F,n-, Take-Off Mode F LC =O Analog Switch (U24) Servoloop Enabled Control Law Implementation Passive networks associated with U26A, U26B, U26C & U26D A. Basic Loop Function The accelerometer signal is amplified by differential amplifiers UlA, It is then applied through the panel NOR.lJ!LAL/TEST switch, S6 UlB and U2A.

(A4), to amplifier U3A for inversion, permitting comparators U5A and U5B to determine that the level of the accelerometer signal is within limits of *3 g’s. Otherwise a failure signal is applied to NOR. gate U39 and the system reverts to a passive configuration. The accelerometer is also applied through the internal NORMAL/TEST switch, U6A to integrator U3B which produces the wing /gear velocity s ignal.

Under active control the accelerometer signal is subtracted from the limit force command signal in amplifier U2OC to produce the force error signal, which is applied to the compensation circuits of amplifiers U26A, U26B and U26C.

The compensated signal is applied through switch U27A to amplifier U28A and then a limiter which is composed of diodes CR12, CR13, CR14 and CR15.

The output of the limiter is applied to amplifier U28C, the output of which feeds the constant current amplifier U29 which, in turn, drives the servovalve. If a failure is detected then switch U34B removes excitation from U29 and therefore prevents the servovalve from being driven.

B. Take Off Mode Several functions begin when power is turned on, while the aircraft is at rest and are listed below. In the description of these functions, where sink rate (Vs) is involved, the sink rate is an input value rather than one supplied by a sink rate sensor.

The scissors switches are closed indicating that the struts are (1) partially compressed.

The “power-on” transient provides RESET of all circuit flip-flops (2) by means of Cl (A4), R.2 (A4), UlA (A4, and UlB (A41 and also energizes the isolation solenoid valves to expose the fluid in the landing gear shock strut pistons to the servovalves and the aircraft hydraulic system.

The tachometer output is low.

(3) The strut position voltage is greater than 0.2V as determined by (4) comparator U41B, indicating that the strut is partially compressed.

Analog switch U6A is in the NORMAL mode which inputs the F (5) w signal into U3B which is not integrating, making the V signal near zero, - wg since capacitor Cl is short-circuited by switch U6B.

is operative at a max- The kinetic energy calculation (Vs - V I2 (6) wg imum high level since V is near zero. The input sink rate voltage, V,, w is applied to multiplier Ull via analog switch U12A, while V is applied from w U3B via inverting amplifier U7A.

The potential energy is calculated by U19, performing the function (7) This value is smaller than (Vs - Vwgj2 so that K. E. >P. E.

(X - Xs) F m wg’ and IMPACT flip-flop is not set to IMPACT.

The sample/hold amplifier U17 is in the SAMPLE mode and F (8) w is near zero. The level of F is processed by the sample/hold amplifier w and supplied to the various circuit points for calculation, including the input to U26A via U6A, U34A and U2OC to close the force loop.

The output of the sample/hold amplifier is applied to TJ18 for cal- (9) culation of FLC , 2 (which is F 2 at this time), and near zero.

(10) For a preset sink rzt the values of FLC2 at the output of SAMPLE/ HOLD amplifier, U17 and V from U3B and U7A are small in comparison to ~!ziT the input sink rate voltage and therefore U13A does not set TRANSITION mode flip-flop U 14A.

(11) The strut pressure amplifiers, U2B, U2D, and U2C, are operative but the output voltage is not introduced by U27A to the servoloop since gate U25B has high level inputs from IMPACT flip-flop U14B and AIR.BORNE gate inverter U36A. That is, U27A is switched to the SERVO mode since the status is not AIRBORNE and lMPACT was not experienced.

(12) Analog switch U6B is not energized, leaving the switch closed around integrating capacitor Cl in amplifier U3B, making V near zero.

w (13) AIRBOR.NE flip-flop U5A (A4) is in the RESET state and the TAKE OFF lamp is energized.

(14) The force loop is closed. Analog switches U24.A and U34A are is fed back to the servo loop.

closed so that FLC=O and Fwg (15) The demodulated strut position signal is applied through U16A, the normal test switch, U12B, U20D, and U24B to U26A to close the position loop.

(16) In the passive test mode, the eight (8) input NOR gate, U39 has low-level inputs indicating the following: F is less than f 3 g’s (the output from U5A and U5B is low) wg Synchro has 400 Hz excitation, evidenced by a low output from excitation comparator U13B as controlled by amplifier U16C and rectifier amplifier U16~ Magnitude of servovalve spool feedback is less than 2 volts as determined by feedback comparators U38A and U38B when differ- ential signals from the demodulator are applied to them. The spool feedback is sensed by the LVDT excited by oscillator U32A and buffer amplifier U33.

Current is flowing in the coil of servovalve as detected by amplifier U35B detector U35D and amplifier U3lD. A low level input to NOR gate U39 is evidence of a satisfactory current level. inputs at pins 2, 5, 11 and 12 of U39 must be low to prevent failure of the test.

The remaining inputs are not used at this time.

C. Aircraft Take Off Several circuit changes accompany aircraft take-off as follows: The tachometer voltage increases with the ground speed of the (1) craft and returns to zero as the wheels spin down after lift-off.

The scissors switches change from closed to open as the aircraft (2) takes off.

The landing gear struts become fully extended upon take-off and the (3) output voltage from position amplifiers UlC, U16B, and U16A nears zero.

When the wheels spin down to near zero, the conditions in 1, 2, (4) and 3 above are used to establish an AIRBORNE signal.

When both channels have developed an AIRBORNE signal, U5A, (5) AIRBORNE flip-flop (A41 is set to the AIR.BORNE state and the LANDING MODE lamp is illuminated. The aircraft can now land under active control. If the LANDING MODE is not desired the pilot can remove power from the system which closes the solenoid valves and isolates the struts from the servovalves and the aircraft hydraulic system.

D. Flight The circuits continue to function during flight as long as power is applied.

If the TEST switch is closed the passive inputs of U39, pins 2,5,11, and 12 will be augmented by the following: The strut is tested for extension by comparator U41B. If the strut (1) position voltage is l.ess than 0.2 V, the #9 input to U3 9 will remain low and no failure will be indicated. This test can be conducted only between AIRBORNE and TOUCHDOWN.

A R.ESET signal is applied to all of the SET-RESET flip-flops (2) following the TEST interval. The AIR.BORNE flip-flop U5A (interchannel schematic) immediately returns to the SET state since the inputs remain high during flight.

E. Pre-Touchdown The servo loop is switched to a pressure control configuration by STRUT PR.ESSURE analog switch U27A which is activated by gate U25B.

The gate is enabled after AIRBORNE flip-flop U5A (A4) is set and prior to IMPACT flip-flop action. In this configuration the servo loop maintains the strut pres- sure at the pre-touchdown bias (charging) level. The solenoid valves are energized to permit control by the servovalves.

F. Landing This procedure begins with TOUCHDOWN and continues through ROLLOUT.

The controller accomplishes the following: The signal for TOUCHDOWN is derived from the tachometer signal (1) TOUCHDOWN flip-flop U43A is when it exceeds the input threshold of U44.

SET to register the event.

Integrator U3B begins to integrate the F signal from accelerometer (2) w amplifiers UlA, UlB, and U2A. Gates UlOA and U9A and amplifier U8A provide a high-level signal to U6B to remove the short from integrating cap- acitor Cl.

Values of F are integrated by U3B and applied to Ull via U7A (3) wg for calculation of kinetic energy (Vs - VwgJ2.

as V w The F signal is applied to U19 along with the strut position (4) w signal (Xm - Xs) from U16A for calculation of potential energy, (Xm - Xs) F w ’ The output of comparator U13C goes high when PE 2 KE and SETS (5) U14B IMPACT flip-fl6p to the IMPACT state. At this point, sample/hold ampli- fier U17 is switched to the HOLD state by the IMPACT state entering U25A.

At the IMPACT state, the servo loop is switched from the pressure (6) loop to the force loop by analog switch U27A and the loop acts to maintain F w by U24A closing the shorting switch across Cl.

This is accomplished by U6B via UlOA. The force loop compares F to the constant value of FLC provided w by sample/ hold amplifier U17 and a signal is applied to the servovalve to c maintain F equal to FLC.

P % i Amplifier U18 squares the now constant value of FLC and applies (7) wg 3.

the output to comparator U13A to determine when (FLCJ2 = (Vs - vw) which SETS flip-flop U14A TR.ANSITION flip-flop to the TRANSITION state.

The TRANSITION state is also used to enable UlOD and U8F to (8) develop a RAMP gate for U17 where the level of FLC is reduced at the rate of 445 kN/sec (100 klb/sec) or 4.902 V/set, by the discharge of R46 and C6.

The servo loop controls F ramp reference until F (9) to the FLC wg w&T At this point the F signal is less than -96 mV and com- reaches F mh.

wz Since U14A is now SET to the TRANSITION parator U32C output goes positive.

state, UlOB is enabled and ROLLOUT flip-flop U43B is SET to provide the ROLL- OUT interval.

(10) When ROLLOUT begins, the ramp gate to U17 is shut-off by UlOD The output of camparator U41C remains high so that UlOC is at and U8F.

a high level, permitting analog switches U24A and U34A to operate, making the servo loop reference FLC = 0 and disconnecting F from the servo loop.

wg This action leaves the servo loop connected as a position loop without a force signal feedback.

(11) The sample/hold amplifier U17 remains in the hold state even though the ramp gate is terminated and the calculations of force parameters are no longer needed.

(12) The output Fwg continues from +Fmin through zero and beyond where comparator U41C shifts to the low logic level causing sample/ -Fmin hold amplifier U17 to perform a short sample of F at near -F and hold min wg this value for reference to the servo force loop which has again become active F is greater than The servo loop now maintains F as F equal to min’ w w at a level of Fmin during this interval.

FLC (13) A return of Fwg to less than F min will cause the servo loop to revert to the position loop configuration and analog switches U24A and U34A = 0.

make FLC (14) Any further excursion of Fwg beyond *Fmin will cause the sample/ hold amplifier to sample, (as driven by U25A), and establish a servo loop refer- The force loop will then be closed to maintain F at ence of FLC = Fmin.

wg F min’ (15) The last portions of R.OLLOUT will find Fwg less than F and min the servo loop controlling position with FLC = 0 and F = 0 as a result of w analog switches U24A and U34A as driven by U9D.

can no longer be a signal to the servo (16) After ROLLOUT begins Fwg Otherwise, the system has position loop loop unless F exceeds F min.

-w activity only.

To take off again with the force loop closed the controller must be RESET by turning the power OFF and then ON again.

G. Control (Loop Compensation) Laws The compensation consists of a notch network and two lead-lag networks.

The notch network is implemented by means of the passive components associated with U26A. One lead-lag network is implemented by means of the passive com- ponents associated with U26B while the other lead-lag network is implemented by means of the passive components associated with U26C.

H. Description of Controller Tests 1. Continuous Tests. - Several tests are made continuously while the controllers are powered. These are as follows: is less than f 3 g’s (1) Fwg Synchro has 400 Hz excitation.

(2) I; f Magnitude of servovalve spool signal is less than 2 V, (LVDT test).

(3) i ,: Current is present in the EHSV coil.

(4) : 2. Pilot Initiated Tests. - Tests which can be made upon Push-to-Test command .-- only are: Strut partially contracted when not airborne (1) Strut position extended when airborne.

(2) Dynamic Test.

(3) All signals are applied to U39 for composition of the failure circuits into a single command source, as illustrated in Figure 7.

3. Detailed Description of test inputs for dynamic test. - The test inputs and relative timing are shown in Figure 8 . This test is performed by introducing a voltage representing 0. 1 g (-0. 3 volts), a voltage representing Vs (10 V) and a voltage representing Xm - Xs (0.5 VI into the system and establishing a test based upon this input. The following functions are performed.

U6A analog switch is set to TEST position and -0. 3 V is intro- (1) duced to integrator U3B as a F signal.

wg U6B is enabled to open the short around Cl and F is integrated (2) wg by U3B.

Ull performs the calculation for kinetic energy, (Vs - V 12.

(3) wg Analog switch U12A is switched to a reference value for Vs of (4) + 10. 0 volts for the calculation of kinetic energy.

U12B analog switch is switched from the strut position signal to (5) a reference of 0.5 V representing (Xm - Xs).

A+B+C+D+E+F+G A = Dynamic Test (Pin 4) B = Fwg Level (Pin 5) c = 400 Hz excitation to position synchro (Pin 2) D = Servovalve spool signal (Pin 11) E & F = Strut position (Pins 9 & 10) G = EHSV Current (Pin 12) Figure 7. Test NOR Gate For H to be high (non-failure); A, B, C, D, E, F, G must all be low.

High when 1.331 <ES 1.089 LVDT output A. 1 I I I 1 0.02 set * 10% after TRANSITION 22 ms 18 ms I 20 ms 1 A = Pin 13 of U22A B= Pin 12 of U22A JMPACT SIGNAL C = kn9ofU22B D= Pin 6 of U22C E = Pin 4 of U23B TEST = Pin 5 of U23B AIRBORNE = Pin 6 of U23B V= Pin 4 of u3l.A 4.7 set I PILOTS TEST INTERVAL I E.

0:K Test S Q",

I

U30A C8 !4-4.9 set- Reset V Sample Test

I

Set V Low for Test 0:K Figure 8. Test Input and Relative Timing U19 calculates the potential energy (X - Xs) F using (Xm -Xs) (6) wg = 0.5 v and F = -0.3 v.

wg U13C compares KE to PE. Using the values of F and (X (7) meXs) wg given, the time for U13C to reach IMPACT is 1.625 seconds. Flip-flop U14B is SET to indicate IMPACT when PE 2 KE.

The high levels of U45B and U14B enable U25A which sets (8) sample/hold amplifier U17 to hold the value of F = -0. 3 volts for calculation wg by U18 to develop (FLd2.

U13A compares Vs (+lOV) with (FL22 + Vwg from U7A and U18.

(9) U14A is SET when Vt is reached and the TRANSITION state is established.

Sample/hold U17 is switched to the RAMP mode by the output of U8F.

The time from IMPACT to TRANSITION is 2. 285 set for the test (10) voltages applied.

The value of F = -0. 3 V is equivalent to more than Fmin.

(11) wg Therefore, the servo loop input is FLc (rather than zero) since U24A is not switched.

U34A is switched so that F to the servo loop is zero.

(12) wg IMPACT flip-flop U14B has been SET so analog switch U27A is (13) closed and the servo loop is closed as when landing and F is greater than wg F min’ The servo loop has an input equivalent to FLc= -0.3 V which (14) begins to reduce at the rate of 4.902 V/set, (445 kN/sec) (100 000 lb/set), from the beginning of the TRANSITION period. The rate is determined by R46 and C6 as in normal operation.

The servo loop operates with F = 0 and the only feedback signal (15) wg is from U32B representing the demodulated feedback from the LVDT pickup.

The closed loop should present an output voltage at U32B of 1. 21 V (16) -3 representing a displacement of 3. 07 x m (0. 121 in). of the spool LVDT.

The circuit in Figure 9 is designed to perform the test for (17) dynamic performance of the system when the craft is airborne.

The coincidence of the LVDT voltage and the reference should occur a time of 0. 02 set 2 10% after the beginning of the TRANSITION mode.

Prior to the TRANSITION gate E. = 0 since Ein = -V diode and il = i2 (the diodes are identical) E. is expressed as 15 (l- eBt’rc) where r = R59+ R60 When TRANSITION goes positive capacitor C begins to charge at and c = C7.

l/C idt but since Eo((15 V then E. g it/C.

/ E. reaches 1.21 V at about 1.21/ 15 = 870 of full charge so that the slope is reasonably linear at this point.

The charge is to reach 1.21 V in 0.02 set, and C7=4.7 mfd, 5%, i = 284 mA.

Let r = 15V/O. 3mA = 50K Ohms l%, R60=10K Pot, choose R61=25K, 1% = 60.43 t volts or 1. 2086 volt in 0. 02 sec.

vC R59 il Airborne Test Circuit Figu re 9.

TRANSITION A failure can be indicated only during the TEST period which is 4.7 (18) This test is positive in nature since an OK test result seconds in duration.

must be achieved during the 4. 7-second test or a failure is annunciated at the end of the test period. In effect, the flip-flop is RESET by the leading edge of the TEST interval. A SET toggle must be received from the OK test circuits during the 4. 7 seconds or the still high state of the a output will report the failure, since the sample pulse completes the AND for a high level output.

time and voltage are linear for a linear sawtooth so that the In Figure 10 , same reference can be used for the time gate and voltage limits of the test.

13.67K 1. 331 Volts I Test Limits 240n 0.242 I’ 1.089 Volts 0.242 1lOOR *t= 4ms = f 10% (0.02 set) 60.43 1.33lV<E <l. 089V A 47K + Kc omparison Gate I I B - I I 18ms 22ms Test Circuit Figure 10.

A successful OK test pulse must be developed in the circuits of (19) paragraph (17) and can occur only after TRANSITION and IMPACT high levels are established. An intersect of the LVDT voltage must occur with 1.21 V f 10% at a time of 0.02 set f 10% or an OK pulse cannot be generated.

A successful dynamic test results in a TEST IN PROGRESS in- (20) dication without a PASSIVE FAILURE indication following. At the conclusion of the test the TEST IN PROGRESS lamp will be extinguished.

The AIRBOR.NE high level gate is AND’ed with the dynamic TEST (21) gate to permit this test only during flight.

A failure in either strut channel causes a signal to be applied through (22) UlB, U8D, Ql and Q2 of both channels to the solenoid valves of both channels thereby isolating the struts from the pressure source and causing reversion to a passive gear status.

X. SYSTEM SPECIFICATION A complete system specification for a flightworthy electrohydraulic active control landing gear system for a supersonic airplane has been included as Appendix B.

XI. CONCLUDING REMARKS The active control landing gear presented in this report is designed to be completely flightworthy and to meet all applicable military specifications.

It contains the ability to detect failures and upon such failure to cause reversion to a passive configuration.

Two potential problems which have not been addressed are: A. The effect of severe vibration at takeoff.

B. The possibility that on takeoff from an extremely uneven runway the strut may be depleted of fluid, causing the gas to fill the hydraulic chamber.

Under these conditions the static design pressure will be lost.

XII. APPENDICES The appendices are as follows: A Stress Analysis B System Specificat ion

APPENDIX A STRESS ANALYSIS

APPENDIX A STRESS ANALYSIS The following is the analysis of the Active Control Landing Gear, Dwg. No. 41004640. In all cases a conservative approach was taken.

In the case of the Orifice Plate Support Assy, under every condition imaginable, the AP (internal & external) was very low; therefore, the AP which the structure was capable of withstanding was calculated and is summarized in the analysis. Load analysis of attachment hardware is also included.

Table Of Contents Page Table of Minimum Margins of Safety Material Properties Design Criteria Weights of Parts 47 Support Assy (Dwg. No. 41004818) Tube Analysis (41004829-001, -002, -003, -004, -007, -008) (41004829-105) (41004829-106) Adapter (41004819-007) 62 Servo Valve Brackets (41004819-007, -008) Accumulator, Clamps & Brackets (Weights & Views) (Load Calculations) (NAS1228C132 Bolt) 79 (Lower Clamp 41004819-003) (Upper Clamp 41004819-004) (Bracket 41004819-006) Bracket, Shutoff Valve (41004819-009) Active Control Landing Gear Table of Minimum Margins of Safety LOADING MIN. M.S. MIN. M..S.

MATERIAL CONDITION YIELD PAGES PART NO. DESCRIPTION ULT 41004818 Support Assy, 321 CRES To determine Max. ext. 1.324x104KPa 49-53 orifice plate max. external pressure = (1921 psi) MLG. & internal pressure assy Max. int. 1.602x104KPa 49-53 can withstand pressure = (2323 psi) 41004829-101 Tube assys 21-6-9 CRES Pressure & bending (-101, -102 1.09 -103 -104) .98 54-57 pressure -104 (-102, -103, -105 -106 -107, -108) 2.26 1.21 54-56 -107 pressure 58-59 -108 (-105) 2.15 1.14 pressure 1.11 59-6 1 (-106) 2.12 41004819-001 Adapter 15-5 PH CRES 9.32 Pressure 5.62 pressure Ei ---- moment (bolt) 1.51 64 ---- 1.21 65 (lug) Servo valve 606 l-T6 alum Wt.& G loads: 41004819-007 sheet (tie-down -008 brackets 68-69 bolts) Large Large (bracket 69-75 Large Large bolts) ---- (-007) 1.57 (tens) --mm (-008) Large (shear) 7075-T7 351 Wt.& G Loads: 41004819-003 Accumulator (thru bolt) Large Large -004 clamps f --mm 81-82 -006 brackets (-003) 0.39 --mm 82-83 (-004) 0.62 s--m 83-86 (-006 bolts) Large 86-87 (-006 A/C load 2782 (tension) Per 86-87 reqts) 746 (shear) Attachment --es 88-90 (-106 fitting) .49 41004640 Landing Gear - Active Control Material Properties 21-6-9 Tubing per AMS5561 Tube Assys - 41004829 - CRES 9.79xlO'KPa F = .87* (min) TU (142 ksi) 8.273xlO'KPa = F (min) .91* T (120 ksi) Y E (elongation) - 20% (min) 1.792x108KPa = E .96* (26~10~ psi) = MIL-HDBK-5C, .286 P table 2.7.1.0 (b) 7.582x107KPa = G .96* (11x106 psi) I = . 3 !J Adapter - 41004819-001 - 15-5PH CRES per AMS5659 1.069x106KPa F (min) .92* TU = (155 ksi) 9.986x105KPa F .92* Ty = (145 ksi) 9.858x105Kpa F . go* cy = (143 ksi) MIL-HDBK-SC, table 2.6.7.0 (c) 668.7KPa F su = (97 ksi) l.965x108KPa = E . 98* (28.5~10~ psi) 7.721x107Kpa = G . 98* (11.2~10~ psi) = .27 lJ *39b°K (250°F) temp.

factors = . 283 P 41004640 Landing Gear - Active Control Material Properties (cant) Manifold - 41004819-002; Clamps - 41004819-003, 0004, ~005; -- Br.acket .- 41004819-006 (7075-T7351 AL ALY per QQ-A-250/12 4.619x105KPa = F (min) .82* TU (67 ksi) 3.93Oxlo'KPa F .85* (min) Ty = (57 ksi) 3.861xlO'KPa .87* (min) FCY = (56 ksi) 2.6020xlO'KPa MIL-HDBK-5C, F table 3.7.3.0 (b2) (min) .91* su = (38 ksi) 7.101x106KPa E = .92* (10.3x106 psi) 2.689x107KPa = G .92* (3.9x106 psi) = . 33 u = . 101 P Brackets - 41004819-007, -008 - 6061-T6 AL sheet per QQ-A-250/11 2.896x105KPa = F (min) .86* (42 ksi) TU 2.413x105KPa F .88* (min) Ty = ,MIL-HDBK-5C, (35 ksi) table 3.6.1.0 (b) 2.413Xlo5Kpa FC = (min) Y (35 ksi) 1.861xlO'KPa F (min) su = (27 ksi) *394'K (25O'F) temp. factors 41004640 Landins Gear - Active Control Material Properties (cant) Brackets - 41004819-007, -008 (cant) 6.825x107KPa = E .97* (9.9x106 psi) 2.620x107KPa MIL-HDBK-SC, table 3.6.1.0 (b) G = .97* (3.8~10~ psi) = . 33 1-I = .098 P - 41004818-001, -002 - 321 CRES bar per QQ-S-763, cond. A Support 5.171x105KPa F .87* TU = (75 ksi) F = 2.068xlO'KPa TY . 91* (30 ksi) E (elongation) = 40% y (red.area) = 50% 1.999x~08KPa = E .96* (29x106 psi) 8.618x107KPa = G .96* (12.5~10~ psi) = .3 IJ = .286 P MIL-HDBK-SC * table 2.7.1 .

0 (b) 1.861x105KPa F .95* cy = (27 ksi) 3.447x105KPa F .a4* su = (50 ksi) / factors temp.

"394'K (25O’F) 41004640 Landing Gear - Active Control Material Properties (COnt) Support - 41004818-003 - 321CRES sheet per MIL-S-6721, camp TI 6.894x105KPa = (100 ksi) (max) F . 87* TU 5.171x105KPa (75 ksi) (min) 2.068x105KPa = F .91* TY (30 ksi) l.861x105KPa .95* FCy = (27 ksi) MIL-HDBK-SC, table 2.7.1.0 (b) 3.44i'x105KPa . 84* FSu= (50 ksi) 1.999x108KPa = E .96* (29x106 psi) 8.618x107KPa G = . 96* (12.5~10~ psi) = .3 P = . 286 P Tube 41004818-004 - 321 CRES tube per MIL-T-8808, type 1 6.894x105KPa I F (MIL-T-8808) TU (100 ksi) (max) 5.171x105KPa = F . 87* TU (75 ksi) (min) MIL-HDBK-SC, 2.068xlO'KPa table 2.7.1.0 (b) F .91* Ty = (30 ksi) (min) 1.861x105KPa .95* FCy = (27 ksi) (min) *39b°K (25oOF) temp.

factors 41004640 Landing Gear - Active Control Material Properties (cant) Tube 41004818-004 (cant) 3.447xlO'KPa =: F .84* sU (50 ksi) (min) 1.9g9x108KPa = E -96" (29x106 psi) MIL-HDBK-5C, table 2.7.1.0 (b) 8.618x107KPa = G .96* (12.5~10~ psi) = .3 1J *394'K (25O'F) temp.

factors = . 286 P I Design Criteria System Limit Pressure = (2200 psi) System Proof Pressure = 1.5 x limit pressure System Burst Pressure = 2.5 x limit pressure G =1.5 G =2.1 G = 1.5 FWD UP INBD Data from aircraft manufacturer (Flight Ult. 'G' Loads) G = 1.5 G = 5.85 GOUTBD= 1.5 I AFT DOWN 41004640 - Landina Gear. Active Control Weights of Parts Weight No. Weight Per Side Description Req'd Part Number N (lb) N (lb) 19.17 19.17 1 Adapter (2) 41004819-001 (4.311 (4.311 34.43 34.43 1 Manifold (3) 41004819-002 (7.74) (7.74) 7.295 14.59 Clamp (4) 41004819-003 (1.64) (3.28) 15.66 31.31 41004819-101 2 Clamp (5) (7.04) (3.52) 4.048 8.095 2 Clamp (6) 41004819-005 (1.82) t.911 1.601 6.405 4 Bracket (7) 41004819-006 l-36) (1.44) . 667 1.334 2 Bracket (8) 41004819-007 (-15) l.30) 1.824 1.824 Bracket (9) 41004819-008 (-41) t.411 3.603 11.03 3 Pipe assy (10) 41004829-101 (2.43) t-81) 4.893 4.893 1 Pipe assy (11) 41004829-102 (1.10) (1.10) 5.738 5.738 1 Pipe assy (12) 41004829-103 (1.29) (1.29) 4.581 4.581 41004829-104 1 Pipe assy (13) (I.031 (1.03) 2.269 9.074 4 Pipe assy (14) 41004829-105 (2.04) t.511 7.072 7.072 1 Pipe assy (15) 41004829-106 (1.59) (1.59) 17.75 17.75 41004829-107 1 Pipe assy (16) (3.99) (3.99) 117.9 117.9 Servo valve (17) 23241830 (26.5) (26.5) 6.005 6.005 1 Relief valve(l8) 5130T-16TT-2000 (1.35) (1.35) 311 311 1 Bracket (19) 41004819-009 Lo71 co71 Valve, sol. oper.

25200 (HRT) shutoff, R.H. 22.24 22.24 (20) Valve, sol. oper. (5.001 (5.001 25450 (HRT) shutoff, L.H.

(22) Ill.2 444.8 4 Accumulator (empty) MS28797-7 (25.0 (100.00) (21) .979 3.914 4 Plug (23) AN814-12 l.221 (.88) 41004640 - Landing Gea,r, Active Control Weisl .s of Parts (cant) Weight No. iieight Per Side Req'd Description Part Number N (lb) N (lb) I.423 1.423 1 Plug (24) AN814-I6 t-32) l-32) 1.957 3.914 Plug (25) AN814-20 t.441 (.88) 1.423 5.693 4 Union (26) AN815-I2 (1.28) t.32) 2.135 8.540 4 Union (27) AN815-16 (1.92) t-48) 4.315 8.629 2 Tee (28) AN824-16 (1.94) l.97) 1.112 1.112 1 Check valve (29) 2C6510 (Crissair) t.25) t-25) 2.669 8.006 3 Reducer (30) AN919-26 (1.80) 1.60) 5.338 5.338 1 Pipe assy (31) 41004829-108 (1.20) (I.201 0534 .667 Bolt (38) 12 NAS1224Cl (:012) t-151 325 1.290 4 Bolt (39) NASl228Cl (:073) t-291 2.980 16 Bolt (40) NASI226C6 ,:EP t.67) 2.344 9.385 Bolt (41) NASl228CI32 t.527) (2.11) 0623 -267 Bolt (42) NASl35lC4HI2 (:014) (. 06) 165 .3.11 Bolt (43) NASl224C32 (:037) l.07) 0311 .I33 Nut (44) AN315C4R (1007) t.031 .0267 .0445 Bolt (48) NASl223Cl (-006) t.011 830.ON Total Per Side (186.59 lbs) (empty assy) 1042N Total Per Side (234.27 lbs) (accumulators full) 41004640 - Landing Gear, Active Control Orifice Plate, MLG 41004818 - Support Assy, 49.2 8 ..8 MIN MIN ,41004819-001 ADAPTER I I I 8 18-002 41004818-003 41004818-004 I N NN I I 558.8 -- 41004818-001

I

I I (B) DIMENSIONS IN MM I Figure A-l. Orifice Plate support Assembly Active Control 41004640 - Landing Gear, Orifice Plate, MLG (cant) 41004818 - Support Assy, the ap between chamber (A) and chamber Based on the System design, (B) (see FigA-1) is always very small unless the servo valve is either not working or has been shut off. If this is the case, any tension or compression loads on the Orifice Plate Support Assy are very small. Therefore, it has been decided to determine the max.

ap the assembly can withstand in either direction before yielding of the material (since material is 321 stainless, cond. A which has a yield allowable of only 40% of ultimate allowable).

Case 1 - Compression Load in Cylinder (pressure in (B)>pressure in (A) Determine max. ap in (B) over (A) 0.559M Length (cyl. wall & 6 tubes)=R=(22.00 in) r(rad. of gyr. -cyl.)=z 'dDp; = +\/(6.05)2 + (5.825)2 = 2.0996 10.47819 +l) = y/Pi = Tiqi202 +-~,5032 = .I9959 r(rad. of gyr. -tube)=q 110.22334 $(tube) = TE =K - CJ (for elastic buckling) where K (fixed gnds) = 4 cr (a_) 2 1.919x108KPa r E=(29x'fJ 1 L96)=(27s84xlo6 psi) 6.901x107Kpa 1~~t27.84~10~) 0 =4 = (10,010,515 psi)(main cyl) cr (10.47819)2 6.237x105KPa ~~(27.84~10~) = CJ cr = (90465 psi) (tubes) (l10.22334J2 for cyl & tubes Based on the above (J = F cr CY 41004640 - Landing Gear, Active Control 41004818 - Support Assy, Orifice Plate, MLG (cant) Case 1 - Compression Load in Cylinder (cant) 1.768x105KPa F cy(main cyl)=27000(.95)= (25650 psi) 1.i'68x105KPa F cy(tubes)=27000(.95)= (25650 psi) 6.05 + 5.825 00108M2 A(main cyl)=nDt=n 2 ('ogo)';I.67879 in2) -4M2 3.995x10 Attubes)= $(.6202-.

5032) = [ (-61917 in2) P (main tubes)=(25650)(I.67879)=~~~~~~1~b~ YIELD PYIELD(tubes)=(25650) (.61917)=:;~~~~1~b~ PYIELD(total)=43061 + I5882 =:;~~~~'~,~ I.324x104KPa AP (Pressure (B)>Pressure (A))== =(I921 psi) $(6.25)2 - Case 2 - Tension Load in Cylinder (Pressure in (A)>Pressure in (B) Cvlinder Wall & Tubes .00108M2 A(main cyl)= (1.67879 in2) -4M2 3.995x10 A(tubes)= t.61917 in2) l.882x105KPa Cyl)~30000(-91)~~27300 psi) F ty(main 41004640 - Landinq Gear, Active Control 41004818 - Support Assy, Orifice Plate, MLG (cant) Case 2 - Tension Load in Cylinder (cant) Cvlinder Wall '& Tubes 1.882x105KPa F ty(tubes)=30000(.91)= (27300 psi) P (main ~yl)=(27300)(1.6789)=~~~~~~'~~~ YIELD PYIELD(tubes)=(27300) (.61917)=:;~~~~1~b~ P YIELD(total)=45831 + 16903=:~~~~~1~bp 62734 1.623x104KPa ap(Pressure (A)>Pressure (B))= 2=(2354 psi) $(5.825) WELDS IN SHEAR (CYL. & TUBES) ND FLUSH 41004640 - Landing Gear, Active Control 41004818 - Suooort ASSY. Orifice Plate, MLG (cant) Case 2 - Tension Load in CyIinder (cant) Cvlinder Wall t Tubes t =. 090 C 063 min tf=.

K(weld factor)=.85 (MIL-HDBK-5C) 2.896x105KPa FS (main ~~1)=(500001 (.84)=t42000 psi) U 2.896x105KPa F (tubes)=(50000)(.84)=~42000 psi) sU -4#2 6.05 ; 5.825 (s7o7) ~,ogo)J.658x10 (main cyl)=?r AS (1.18690 in2) H -4M2 3.529x10 + T (.707)(.063)=( 546gg in2) AS (tubes)=6(n) .

H P (main cyl)=(42000)(.85) (1.1869)=~~~~:~1~b~ SHEAR P SHEAR(tubes)=(42000) (.85)(.54699)=~;~~~~1~b~ SHEAR(total.)=42372 + 19528=:~:~~~1~b~ P l.602x104KPa 61900 ap(Pressure (A)>Pressure (B)) = =(2323 psi) -$(5.825)2 Summary - Max Pressures Support Can Withstand Max Ext Pressure l.324x104~pa (Pressure in (B)>Pressure in (A)) (1921 Psi) Max Int Pressure 1.602x104KPa (Pressure in (A)>Pressure in (B)) (2323 psi) 41004640 - Landinq Gear, Active Control Tube Analysis 0254M (1.00 dia tube) O.D. .

min=(l.OO in) I.D. max=1.00-2(.052-.005)=.0230M t.906 in) 01905M t.75 dia tube) O.D. - min=(.750 in) 750-2(.039-.oo5)='"'732M I.D.max=.

t.682 in) 0159M c.625 dia tube) O.D. * min=(.625 in) 01445M I.D. =.625-2(.033-.005)=' max t.569 in) R max .5015 3765 .314 =-= 11.21 -c-z JO.67 =+= 11.07 .028 t .047 min Since these tubes are borderline between thinwall and thickwall, use thick wall analysis (Roark, 4th Ed., Table XIII, Case 35, P. 308).

Tube Assys 41004829-101, -102, -103, -104, -107, -108 '2 .906 01151M where - fax 2 -=i.453 in) (2 =p- b2:a2 1.00 .0127M ,bT,T = (.500 in) 1.517x104KPa P= (2200 psi) 6.949x104KPa 4532 = aR=2200 -2 500 -.453 . (10079 psi) b2+a2 (2200)(.5002+.4532) =1.541x105KPa =p- = 'H b2 a2 2 -500 (22359 psi) -.453 41004640 - Landing Gear, Active Control Tube Analysis (cant) Tube ASSYS 41004829-101, -102, -103, -104, -107, -408 (cant) -1.517x104KPa ar'-P' (-2200 psi) 8.465x104Kpa T=p- =(2200) b2:12 (12279 psi) 1.042x105Kpa o~(proof)=10079x1.5~ (15119 psi) 2.312X105Kpa uH(proof)=22359x1.5= (33539 psi) -2.275x104KPa ur(proof)=-2200x1.5= (-3300 psi) l.270x105KPa T(proof)=12279x1.5= (18419 psi) uE (equiv. stress theory)-.707 (a -o~)~+(o H-ur)2+(ur-uk) (Bruhn, P. Cl.91 J R 2.199x105KPa 15119-33539~2+(33539+3300~2+(-3300-,5119p = OE=- (31899 psi) (proof) 120000(.91) M.S.

1= )2] YIELD= 31899 - (pressure only) (.55) (120000) (.91) M.S.

l= (2.26( SHEAR= 18419 I YIELD l.737x105KPa uR(burst)=10079x2.5= (25198 psi) aH(burst)=22359x2.5=3~854x'o KPa (55898 psi) -3.792x104KPa cr(burst)=-2200x2.5= (-5500 psi) 41004640 - Landing Gear, Active Control Tube Analysis (cant) Tube ASSYS 47004829-707, -702, -103, -704, -107, -108 (cant) 2.116x105KPa r(burst)=12279x2.5= (30698 psi) 707 (25198-55898)2+(55898+5500)2+(-55~t)-~~,g~)2 =3-665x1o KPa uE=.

J- (53164 psi) 142000(.87) M.S.

-I= r1.321 ULT= 53164 (pressure only) (.55) (142000)(.87) M.S.

-l-p1 SHEAR= 30698 ULT Bendins of 41004829-101 Assume no support on tube assys from servo valve to Aircraft Pres- sure Return System. Let total weight be assumed at g shutoff valve and reacted in bending by above tube assy.

(very conservative).

5?.29N Total Wt=5.00+.48+1.03+.97+1.29+1.35+.81+.60~ (11.53 lb) SHUTOFF

/

TUB TUBE ASSYS VALVE -a- ---_ 51.29N 7," 1.53LB.)

49.53CM r) (SCALED) -gM4 4 6.664~10 I(tube)= $ 5o04 .

-.453

C 1 =(.016014 in4

41004640 - Landing Gear, Active Control Tube Analysis (cant) Bending of 41004829-101 (cant) . 0127M 'G' factor (vert. down)=5.85 g c=(.500 in) M=(11.53)(5.85)(19.62)='4g.5N-M (1323.3788 in/lb) (T = (1323.3788 in/lb)(.500) 2.849x105KPa =3302KPa T= (11.53)(5.85) = b .016014 (41319 psi) $(1.002-.9062) (479 psi) 3.891x105KPa 1.303x105KPa =41319+15119 = Total uR Total T=18419+479= (56438 psi) (proof) (18898 psi) (prooT) 4.586x105KPa =41319+25198= 2.149x105KPa (66517 psi) (burst) =30698+479= (31177 psi) (burstr 3.598x105KPa = (52194 psi) (proof & 5.85 g bending) ~20000(.91) -,= M.S.

YIELD= 52194 (proof pressure & bending) (.55) (?20000) (.91) M.S.

18898 SHEAR= YIELD 2 4.304x105KPa = (62433 psi) (burst & 5.85 g bending) 142000(.87) M.S.ULT= 62433 (burst pressure & bending) t.55) (142000) t.87) -,= M.S.SHEAR= 31177 ULT Active Control 41004640 - Landing Gear, -- Tube Analysis (cant) Bendina of 41004829-104 Assume all parts from servo valve to manifold have their C.G. at E of shutoff valve and are reacted in bending by above tube assy.

(very conservative).

10.14N Total Wt=.60+1.20+.48= (2.28 lb) M=(9.02)(5.85) (6.88) =41.02N-M (SCALED) (363.037 in/lb) I Since moment is less than -101 t and tube is same diameter & thickness, further analysis is unnecessary.

Tube Assv 41004829-105 1.517x104KPa a = . = 682 .00866# P= max 2 (-341 in) (2200 psi) 750 . 00953M = b %---= min t-375 in) .341 =7.244x104KPa uR (limit)=(2200) 2 2 .375 -.341 (10508 psi) 3752+.341 1.601x105KPa UH(limit)=(2200)' 2 =(23217 psi) .3752 -.341 -1.517x104KPa U, (limit)= (-22oo psi) 8.761x104KPa T(limit)=(2200) .

2 =(I2708 psi) .3752-.341 UH(proof)=23217x1.5=2.401x10 KPa (34826 PSI) 41004640 - Landing Gear, Active Control Tube Analysis (cant) Tube Assy 41004829-105 (cant) 2.275x104KPa ur(proof)=-2200x1.5= C-3300 psi) 1.314x105KPa r(proof)=12708x1.5= (19062 psi) U =.707 (15762-34826)2+(34826+3300)2+(-3jOO-,5762)2 =2-276x'o 'pa E (33013 psi) 1200001.91) -,= M.S.

YIELD' 33oLl3 (proof pressure only) (.55)(120000)(.9~) -,= M.S.

SHEAR= 19062 YIELD 1.811x105Kpa ~~(burst)=10508x2.5= (26770 psi) 4.002x105KPa aH(burst)=23217x2.5- (58043 psi) or(burst)=-2200x2.5=~~~~~~x~~i~pa 2.190x105KPa T(burst)=12708x2.5= (31770 psi) 3.793x105KPa uE=. 26270-58043)2+(58043+5500)2+(-5500-26270)2 = (55022 psi) 142000(.87) += M.S.

ULT= 55022 (burst pressure only) t (.55)(142000)(.87) -,= M.S.

SHEAR= 31770 ULT 41004640 - Landing Gear, Active Control Tube An‘alysis (cant) Tube Assy 41004829-106 1. 517xlo4KPa . 569 .

007231.1 P = (2200 psi) amax= 2 = f.2845 in) b = ++=.00794M min t.3125 in) OR (limit)= (2200) .3125 -.2845 .i""" 2 =7.344x104KPa 2 (10653 psi) .31252+.28452 = 1.620~10~~~a oH(limit)=2200 . 31252 -.28452 (23505 psi) -1.517xlo4KPa ar (limit) = (-2200 psi) 2 4 T(limit)=(2200) '2"' =8*861X10 KPa .3125 -.2845 2 (12853 psi) l.102x105KPa oR(proof)=10653x1.5= (15980 psi) 2.431x105KPa uH(proof)=23505x1.5= (35258 psi) -2.275x104KPa artproof)=-2200x1.5= (-3300 psi) l.329x105KPa T(proof)=12853x1.5= (19280 psi) aE= I 15980-35258)2+(35258+3300)2+(-3300-,5g80)2 =2-3o2x1o KPa (33387 psi) 41004640 - Landing Gear, Active Control Tube Analysis (cant) Tube Assy 41004829-106 (cant) 120000(.911 M.S.

YIELD= 33387 (proof pressure only) (.55) (120000) (.91) -1= M.S.

SHEAR= 19280 YIELD 1.836x105KPa uR(burst)=10653x2.5= (26633 psi) 4.051x105KPa uH(burst)=23505x2.5= (58763 psi) -3.792x104KPa ur(burst)=-2200x2.5= (-5500 psi) 2.215x105KPa -r(burst)=12853x2.5= (32133 psi) uE=.707 (26633-58763)2+(58763+5500)2+(-5500-26633)2 =3*836x1o KPa (55645 psi) 142000(.87) -1- , 22 M.S .ULT= 55645 -I (burst pressure only) (.55)(142000)(.87) -1= M.S.

SHEAR= 32133 ULT 41004640 - Landing Gear, Active Control Analysis - Adapter 41004819-001 (see FigA-7) Adapter is at top of MLG Support Assy, bolted to top and cut from square stock as an elbow. Cross section is square outside and circular inside, making it stronger than a simple cylinder.

Part will be analyzed as a cylinder using min.wall thickness as constant (very conservative). (Roark, 4th Ed., Table XIII, Case 35, P. 308).

*Min Top Wall =(2.75-.OlO)-(1.88+.010) +.00+~010)=;"~~~"r,, .

*Min Side Wall =(.70-.OIO) -~(l.oo+.olo)=;O~~~Or,, .

1.3125 .00492M *Min Threaded Fluid Conn =(2.75-.OlO)-(1.88+.010)-( 2 =(.19375 in) *See Fig. I . 01753M l.517x104KPa a=.6g-.185=a01283M (limit) (.505 in) "(.69 in) p=(2200 psi) a2 .505 2 1.750x104KPa cR=P - = =(2200) 2 2 b2-a2 .69 -.505 (2538 psi) b2+a2 6g2+.505 2 =5.016x104KPa - =(2200)' c*=P -69 2 -.5052 (7276 psi) b2-a 2 -l.517x104KPa cl = r (-2200 psi) 2 4 b2 2 =3.266x10 KPa =(2200) 2.69 T=p - -.505 b2-a2 .69 (4738 psi) 2.625x104KPa cR(proof)=2538x1.5= (3807 psi) 7.524x104KPa aH(proof)=7276xI.5= (10914 psi) -2.275x104KPa artproof)=-2200x1.5= (-3300 psi) 4.900x104KPa r(proof)=4738x1.5= (7107 psi) 41004640 - Landing Gear, Active Control ~ .~ Analysis - Adapter 41004819-001 (cant) 3807-10914)2+(10914+3300)2+(-~~00-3807)2 =8*485x1o KPa (12308 psi) '45000(.921 -,- M.S.

- YIELD= 12308 (.55)(145000)(.92) M.S.

SHEAR= 7107 YIELD 4.374x104KPa aQ(burst)=2538x2.5= (6345 psi) 1.254x105KPa uH(burst)=7276x2.5= (18190 psi) -3.792x104KPa ar(burst)=-2200x2.5= t-5500 psi) 8.166x104Kpa T(burst)=4738x2.5= (11845 psi) (6345-18190)2+(18190+5500)2+(-5500-6~~~)2 ='-4'4x1o KPa (20513 psi) '55000(.92) -,= M.S.

ULT= 20513 t-551 (155000) t-921 -'= 5 62 M.S. .

SHEAR= 11845 El ULT Loading on hold-down bolts = hydraulic end load + load due to motion of tubes & fittings between adapter & servovalve (which is tied to structure).

Assume entire load of tubes imposes moment on end of adapter, moment arm at C.G. of tubes (very conservative).

41004640 - Landing Gear, Active Control - Analysis - Adapter 41004819-001 (cant) ADAPTER 41004819-001 -2.54 DIA 24.9 (EST.]-E~

I-

DIMENSIONS IN CM Vertical Load = (2200 psi) ($) (1.00)2=~~~~~ lb) Wt(tube assys)=2(.8l)+(.48)+(.97)+(l.lO)+(7.35)=~~o~~Nlb) .

Moment = (5.52)(9.8)=~;~'~~~Min-lb) .

Assume moment reacted by 2 hold-down bolts: +(1728) + T Load/bolt = ' ~;:~~"]=~:~~"096 lb) '.294x'04N Min breaking strength of bolt = (2910 lb) (NAS1351C4H12) Assume burst pressure 6 max 'G' factor = 5.85 on Wt: +(1728x2.5)+; =:;;;; 12 lb) Max load/bolt = .

29'0 -,= (bolt) M.S.

BOLT= 1159.12 ULT 41004640 - Landing Gear, Active Control Analysis - Adapter 41004819-001 (cant) 6M 6(1'159.'2x.29) ub(on tab)= - =- (very conservative) bt2 (.50)(.25)2 6.35THICK u =4.449x105KPa b (64540 psi) WLT) 155000x.92 M.S.ULT= (lug) 64540 DIMENSIONS IN MM Landing Gear, Active Control 41004640 - Servo Valve Brackets 41004819-007, -008 Assume weight of all components from 41004819-001 Adapter to 41004819-002 Manifold are reacted at C.G. of 23241830 servo valve by the above brackets (very conservative).

179.3N Total Wt=26.5+11.53+2.28- (40.31 lb)"

--_

--

--

--

---

r -- -- -----_

r=-

DIMENSIONS IN CM Figure A-2. Servovalve Brackets W ='.5W W -1.5w OUTBD AFT W W =W(G W W =W (GFWD) FWD DOWN=W(GDOWNJ UP UP) INBD=W(GINBD) =?.5W =5.85W =2.'W =7.5w 1: Active Control I 41004640 - Landing Gear, Servo Valve Brackets 41004819-007, -008 (cant) Servo Valve Bolt Loads t 17.8 17.8 E R3fR4 + R4D - D D R2D D INBD DIMENSIONS IN MM =R (1) (CFD) WDOWN=R, +R2 +R3 +R4 -R; ; =R; =R; =.25WDOWN D D D D D D D D (CMI) W~~~~(2'50)=(R2 +R4 )(2.00)-(RI +R3 )(2'00) D D D D +R3 )-let KD=-(R2 +R4 )=(R, +R3 ) ‘-25WINBD’-(R2 +R4 )+(R, D D D D D D D D 2KD=1.25W INBD =--625 W =+.625 W -R +R &R +R KD=.625W INBD 13 INBD INBD 2D 4D D D from symmetry II 11 II II R4 =-.3125wINBD (2j R3 =+.3725WINBD R2 =-.3i25WINBD =+-3125WINBD R1 D D D D lCMF) WFWD(2.50)=(R3 +R4 )(3.00)-(RI +R2 )(3.00) D D D D +R4 =- )-let KL=R3 .83333wFWD=(~3 +R4 )-(R, +R2 (RI +R2 ) D D D D D D D D 2K;=.83333WFWD 41667WFwD & RI +R2 =-.4'667WFwD K;=.41667WFWD-R +R4 =.

D D 3D D 41004640 - Landing Gear, Active Control Servo Valve Brackets 41004819-007~ -008 (cant) Servo Valve Bolt Loads (c.ont) from symmetry III III 88, R3 =.20833W =.20833WFwD RI =-. 20833WFwD R2 =-.20833WFwD (3) FWD D D R4D D 1049N Combining (I), (2), (3) = (4) WDOWN=5.85(40.31)= (235.8135 lb) 269.0N W =w FwD=1.5(40.31)= INBD (60.465 lb) 290.2N =.25WDOWN+.3125WINBD-.20833W FWD=(65.25201 lb) R7 D 122.2N =.25WDOWN-.3125WINBD-.20833w FWD=(27.46139 lb) R2D (4) 402.3N =.25W DOWN+.3125W +.20833W INBD FWD=(90.44536 lb) R3D 234.2N =.25WDOWN-.3125WINBD+,20833W

I

FWD=(52.65474 lb) R4D 67.24~ =R2 =R =R4 (CFI) WINBD=Rl +R2 +R3 +R4 -R =~25WINBD=(15.1163 lb) (5) I I I I I '1 I 3I 67.24~ =R =R =R =.25W (CFF) WFWD=R1 +R2 +R3 +R4 -R FWD=(15."63 lb) 12 3 4 (6) F F F F F F F F Analysis of NAS1228ClW bolt (l/2-20 bolt) (4) =.25WDOWN+.3125WINBD+.20833WFwD Max tension load =R3 D 402.3N (90.44536 lb) R3D= u = 90.44536 =39OOKPa t .I599 (565.637 psi) "8000(.9'1 -,= LARGE M.S.

YIELD= 565.637 (5) (6) Max shear load =d- =$.25WINBD)2+(.25WFwD)2 =;;;I);;768 lb) R3 I F 41004640 - Landinq Gear, Active Control Servo Valve Brackets 41004819-007, -008 (cant) pervo Valve Bolt Loads (cont.), 991.8KPa =- 2.1 . 37'7'68 . 1486 = (743.861 psi) . 5~5*[“8000(.9’)1 -,- LARGE M.S.

YIELD= 143.861 -1 Bracket Bolt Loads 2F 2F 2F AD BD cD )R, (3.00)'R3 (3.00)=3.5O(FFF )+2.'O(F+FE )+.7O(FF (CMD 1-3 D D D D D D DD) D 3.0O(R -R3 )=7FA +4.2F +1.4F 'D D cD D BD 13 13 13 (CM )3.00(R2 -R4 )=7FA +4.2F +1.5F B C D2-4 D D D24 D24 D24 Ml 3=3.00(R1 -R3 ) M D D 'A =3.00(R2 -R4 ) FA = 7 D(~~u~~, P.Dl.14) where M2-4 D D D c ]=2[3.502+2.io2+.702i I=2[r2 +r2 +r2 AD BD D =34.30 *Ref. MIL-HDBK-5C, Para. 1.4.6.3 41004640 - Landing Gear, Active Control Servo Valve Brackets 41004819-007, -008 (cant) Bracket Bolt Loads (cant) (Mlw30r M2,4) (3.50) =F; = =F' =.30612 (RI -R3 );F; 34.30 D D FD D13 D13 D24 24 =.30612(R2 -R4 ) D D (Mle30r M2,4) (2.10) F; =F; = =.I8367 (R, -R3 );F; =F; 34.30 D D D13 D13 D24 D24 =. 18367(R2 -R4 ) 1 (7) D D (Ml,30r M2-4) t.70) =F' = F' =.06122 (R, -R 3 );F; =F; D 34.30 cD D D D D13 13 24 D24 =.06122(R2 -R4 ) D D (CFD )R, +R3 =2FA +2FB +2FC -cF; =F; =F; =F; =F; =F; l-3 D D D D D D13 D13 D13 D13 D13 D13 = ;(R, +R3 ) D D (8) =F; =F; =F; =F; =F; )R2 +R4 =2FA +2FB +2FC -F; ( CFD D D 2-4 D D D D24 D24 D24 D24 D24 D24 = ;tR2 +R4 1 D D !

) ) (RI +R3 ) (.50)=3.50(-FA +FF )+2.10(-FB +FE ( CMF F F l-3 D13 D13 D13 D13 +.70(-FC +FD ) D13 D13 ‘i [ g II i i 41004640 - Landing Gear, Active Control i Servo Valve Brackets 41004819-007, -008 (cant) Bracket Bolt Loads (cant) +F~ ( CMF ) )+2.10(-FB +FE 1 (RARE ) (.80)=3.50(-FA I 2-4 F F D24 D24 D24 D24 +.70(-F, +FD ) L D 24 D24 t.50) (R, +R3 ) (3.50)

\

111 111 F F = -FA =FF = . 05102 (RI +R3 ) 34.30 F F D13 D13 t.50) (RI +R3 ) (2.10) III F F -F =F = = . 03061 (RI +R3 ) E 34.30 BD F F D13 t-50) (RI ) (-70) +R3 111 111 F F = . 01020 (RI +R -F =F = 34.30 F 3F) cD DD 13 13 t.80) (R2 +R4 ) (3.50) (9) ,I, ,1, F F -F =F = = . 08163 (R2 +R4 ) A F 34.30 F F D24 D24 l.80) (R2 +R4 ) (2.10) 111 111 F F =F = -F = . 04898 (R2 +R4 ) B E 34.30 F F D24 D24 t.80) (R2 +R4 ) (-70) 111 111 F F -F =F = = . 01633 (R2 +R4 ) C D 34.30 F F

/

D24 D24 l-3 Combining (7), (8), (9) = (10) (for right side) II F =.47279R'1 -.13945R3 -.05102(R, +R =F; +FA +F;' A D D D F D13 D13 13 D13 (101 F 74.26~ AD = (16.69543 lb)

I

41004640 - Landing Gear, Active Control Servo Valve Brackets 41004819-007, -008 (cant) Bracket Bolt Loads (cant) F F' +F; +F;' =.35034R, -.017R3 -.03061(R, +R3 ) BD = BD D D F F 13 13 D13 D13 F 90.73N BD13= (20.39740 lb) =F; +FC +F;' =.22789R, -.10545R3 -.01020(R, +R3 ) FC D D F F D13 D?3 D13 D13 F 107.2N cD 13=(24.09937 lb) II 111 F F' +F +F =.22789R, +.?0545R3 +.0?02O(R, +R3 ) (10) D D DD =D D D F F D?3 D13 D?3 13 F '09.9N D D13=(24.71612 lb) FE =F; +F; +F;' =.35034R, -.017R3 +.03061(R, +R3 ) D D F F D13 D13 D13 D13 F 98.96N ED13= (22.24824 lb) =F; +F; +F;' =.47279R, -.13945R3 +.05102(R, +R3 ) FF D F F D D13 D13 D13 D13 F =87.98N FD 13 (19.78036 lb) 41004640 - Landing Gear, Active Control Servo Valve Brackets 41004819-007, -008 (cant) Bracket Bolt Loads (cant) 2-4 Combining (7), (8), (9) = (10) (for left side) I I, 111 F =.47279R2 -.13945R4 -.08163(R =FA +F +F A 2F+R4F) AD AD D D 24 D24 D24 24 F 14.11N A D24=(3.17288 lb) FB =F' +F +F =.35034R2 -.017R4 -.04898(R2 +R B D D F 4,) D24 D24 BD24 BD24 F =32.23N B D24 (7.24490 lb) I II 111 FC =FC +FC +FC =.22789R2 +.10545R4 -.01633(R' +R4 ) D D 2F F D24 D24 D24 D24 F =50.34N ’ (10) cD 24 (11.31692 lb) II IIt FD =F; +F +F =.22789R2 +.10545R4 +.01633(R2 +R4 ) D D D D D F F 24 D24 D24 D24 =54.73N FD D24 (12.30432 lb) FE =F; +F; +F;' =.35034R2 -.017R4 +.04898(R 2F+R4F) D D D D24 D24 D24 24 F 45.40N ED 24=(10.20649 lb) F =F; +F; +F;' =.47279R2 -.13945R4 +.08163(R I F 2F+R4 D D F D24 D24 D24 D24 F =36.07N F D24 (8.10865 lb) 41004640 - Landing Gear, Active Control Servo Valve Brackets 41004819-007, -008 (cant) Bracket Bolt Loads (c'ont) \ )R, +R3 =FA (CFF +FB +FC +FD =FB +FE -F +FF A l-3 F F F13 F13 F13 F13 F13 F13 Fi3 F13 =;(R, +R3 ) =F C =FD =FE =FF F F F13 F13 F13 F13 b(") )R2 +R4 =FA (IF +FB +FC F +FD +FE +FF -F AF = BF F2-4 F F F24 F24 F24 F24 F24 F24 24 24 =F =F =F =F =;(R2 +R4 ) C D E F F F F24 F24 F24 F24 ;(R, +R3 )=;(R2 +R4 )=::-;;;8 lb) I F F FF * =FF = c = D =FE (CF )FA =FB F F =;(R, +R3 1 [tens.or camp.]

II-3 I I I I13 I13 I13 I13 I13 13 (12) )FA =FB =FC =F =FE =F (CF =$(R2 +R4 1 [tens-or camp.]

D F '2-4 124 I I '24 '24 '24 '24 '24 I ;(R, +R3 )=&(R2 +R4 )=::';;;8 lb) I I II - Analysis of NAS1224ClW Bolt (l/4-28 bolt) 22.369N Max Tension Load = F~ = I13 (5.0388 lb) 9.544xlo5KPa 5.0388 Crt = -En= (138.429 psi) '18000(.91) M.S. -I= YIELD= 138.429 112.2N Max Shear Load (24.71612)2+(5.0388)2 =(25.2245 lb) Landing Gear, Active Control 41004640 - Servo Valve Brackets 41004819-007, -008 (cant) Analysis of NAS1224ClW Bolt (l/4-28 bolt) T= 25.2245 =533.4KPa I 0326 (773.758 psi) .55[“8O’JO(.91)1 -,= M.S.

773.75 YIELD= At bend of bracket, stress results from tension load and side load tending to bend bracket. The bending is distributed somewhat evenly along the length of the bracket, but the tension load is much less evenly distributed. Conservative analysis will be used.

) (1.25+.80.)

6 (R2 +R4 I I 8.074X104KPa = ub(bending stress)= z2 = bt (8.00) (.063)2 (11711.4 psi) F DD 13 6(24.71612) =2029KPa ut(tensile stress)= r= (8.00) l.063) Zbt (294.24 psi) u +u =8.277x?04KPa U TOTAL b t (12006 psi) (35000) (.88) M.S.

12006 YIELD= FA +FB +FC +FD +FE +FF R2 +R4F =413.5KPa F F F F F 30.2326 F F ;e = ‘c= bt bt (8.00) t.063) (59.985 psi) .

83(27000) C.88) -,= LARGE M.S.

YIELD= 59.985 Since the Up load on brackets is much smaller than down load.

buckling length of the bracket sheet metal is so small, it will not be necessary to analyze it for "up" load.

Active Control 41004640 - Landing Gear, Accumulator, Clamps L Brackets Composite Weight (assume reacted at C.G. of assy) .N LBS 4 Accumulators (empty) (MS28797-7) 444.8 100.00 14.59 2 Clamps (41004819-003) 3.28 7.04 2 Clamps (41004819-004) 31.31 8.095 1.82 2 Clamps (41004819-005) 6.405 1.44 4 Brackets (41004819-006) 2.11 4 Bolts (NAS1228C132) 9.385 2.980 16 Bolts (NAS1226C6) .67 2 Bolts (NAS1224C32) . 311 .07 4 Nuts (AN315C4R) . 133 . 03 2 Bolts (NAS1223Cl) . 0445 .Ol 7 Valve, Sol. Oper. Shutoff (25200 or 25450) 22.24 5.00 1 Bracket (41004819-009) . 311 .07 Oil for Accumulators (1600 in3x.0298 lb/in3) FJ12.1 47.68 Accumulator Subtotal 752.7 169.22 1 Manifold (41004819-002) 34.43 7.74 9.074 2.04 4 Pipe Assy (41004829-105) 4 Plug (AN8?4-12) 3.914 .88 1 Plug (AN814-16) 1.423 I 32 4 Union (AN815-12) 5.693 1.28 Total 807.2 181.48 w,= $69.22) =376.4N (84.61 lb w2= $169.22) +(181.48-169.22) 'FWDd- l&2 430.9N =(96.87 lb)

I

I ’ f .

W DOWN DIMENSIONS IN CM l&2 END VIEW 41004640 - Landing Gear, Active Controls Accumulator, Clamps L Brackets (cant) SIDE VIEW W W =2.1w, W =1.5w, W =1.5w, =5.85W, DOWN, INBD, FWD, up1 W =L5W2 W =5.8512 W =2.1w2 W =1.5w2 INBD, FWD2 DOWN2 UP2 w,=. 87344W2 =2202N 790.3N 564.5N W W W =(177.681 lb) FWD, =(126.915 lb) DOWN, (494.969 lb) up1 646.3~ 2521N =904.8N W W W FWD2=(145.305 lb) DOWN2 =(566.690 lb) (203.427 lb) up2 =564.5N W (126.915 lb) INBD, 646.3~ W lb) INBD2 =(145.305 Active Controls 41004640 - Landing Gear, Clamps & Brackets (cant) Accumulator, Load Calculations looseness of clamped Due to flexure of accumulators and reasonable 1 and area 2 can be treated separately.

the loadings at area assembly, (CFD) R;, =R' = f WDOWN 2 D4 I =R' fW DOWN, RD, D3= (22;275) )- (R" +R" ) ) (5.40)= (R; +R; (CM11 (WINBD +WINBD 2 2 4 Dl D3 I [ =R;; z-R; CR" 1 (2) -R; =.12067(wINBD +wINBD D4 12 3 1 loading must be treated as (CMF) Since this is a clamped assembly, instead of as an assembly.

reacting on individual parts =w FWD (5.40) MF 1 1 F= (5.40 i I 5.50)w~WDl I I - COMP. DIRECTLY A F=.98182WFWD INTO BRACKET '

T

3.7 w 11.38-8.44) (. 98182WF RD =( 11.38 1 1 "FWD, I II 7w R 5.7 287 R =.25365WFwD FWD~ I --- 1 Dl D3 (3) .7 281 :. R =.25365WFwD k28-:-$ 1 RD4= 7WFWD2 ) D2 2 DIMENSIONS IN CM Active Controls Landing Gear, 41004640 - Clamps f Brackets (cant) Accumulator, Load Calculations Combining (I), (2), (3) = (4) i t +.25365WFwD +WINBD -.12067 R =R; +R; +R;'=.5WDDWN

1 1 2 1

D1lll 1098N =(246.82770 lb) +.I2067 +R;'=.5WDOWN R =R; +RI1 +WINBD +.25365wFwD 1 2 3 2 D2 2 D2 2 2 1570N =(353.05040 lb) +.72817wFwD WINBD +WINBD -.I2067 =R; +'R; +R;'=.5WDOWN RD 1 2 3 3 1 1 3 3

C 1 '(4)

1366N =(307.05141 lb) +.72817WFwD =R' +Rn +R =.5w WINBD +WINBD DOWN +*12o67 RD 2

1 2 1

D4 D4 D4 C 1877N =(422.00053 lb) =RI +RI +RI +RI ( CFI)WINBD +WINBD 2 12 3 4 302.7N )= =.25(WINBD +WINBD R =RI =RI =RI (5) (68.055 lb) 1 2 I1 2 3 4 (CFF)WFWD +WFWD =RF +RF +RF +RF 1 2 12 3 4 302.7N =R =RF =R =.25(wFWD +wFWD (61 lb) )=(68.055 RF 1 2 4 1 F2 F3 Analysis of NAS1228C132 Bolt (I/2-20) +.72817WFwD +. 12067 +WINBD RD =.5WDOWN Max Tension Load = 23 2 4 2 1 1877N RD = lb) 4 (422.001 41004640 - Landing Gear, Active Controls Load Calculations (cant) Analvsis of NAS1228C132 Bolt (l/2-20) 422.001 =1.819x104KPa at= .I599 (2639 psi) 118000(.91) -1= LARGE M.S.

YIELD= 2639 Pp. (422.001) (10.90) .00271MM G(max defl)= 'AE =(.00106739 in) (.1486)(29x106) (5) (6) Max Shear Load= (68.055)2+(68.055)2 =;;;-;z,, lb) .

~= 96.2443 =4465KPa .

(647.674 psi) .55[118000(.91)1 -,= LARGE M.S.

YIELD= 647.674 Active Controls 41004640 - Landing Gear, Accumulator. ClamDs & Brackets (cant) Analvsis of Lower Clamr, (41004819-003) g SYM E BOLT E BOLT I 6.99-!= 7.47 I- I I I I 0. 0'1 02 0. 0'076 t ‘2.44 2 36THICK(-003);

I R . I R

2.79 t.051 THICK(-004) D4 D2 DIMENSIONS IN CM F2=.5W 1=;::::19379 lb) DOWN +.12067[WINBD +WINBD 1 2 +WINBD ]=.98182WFwD P4=. SW +.I2067 [WINBD DOWN2 1 2 =2041N (458.85715 lb) RD,

;!

F2 RD - F2

I

R D4 2.94 R D2 117.3N-M 140.2N-M =(1240.6816in-lb) 41004640 - Landing Gear, Active Controls Accumulator, Clamps & Brackets (cant) Analysis of Lower Clamp (.4!CJO4819-003) (cant) 140.2N-M M =2.94RD = MAX (1240.6816 in-lb) 2.3& bMIN = t.93 in) t MIN=(2.99-2.51)=;-~~c~n) .

6t1240.6816)

ad2L=

b bt2 (.93)(.48)2 2.395X105KPa 'b= (34741 psi) 57000(.85) M.S. -I= [0.391 YIELD= 34741 F4 458.85715 7087KPa == bt = l-93) t-48) =(1028 psi) [38000(.91)] (g) M.S.

YIELD= Analvsis of UDoer Clamp (41004819-004) Assume same loading conditions at -003 apply to -004.

140.2N-M M =2.94R = MAX D4 (1240.6816 in-lb) 2.74cM t MIN'(2-99-2.51)';:;;cyn) b MIN=(1.08 in) 6~ (J = _ 6tl240.6816) =2.062x105KPa b bt2 (l.08)(.48)2 (29916 psi) 41004640 - Landing Gear, Active Controls Accumulator, Clamps & Brackets (cant) Analysis of Upper Clamp (41004819-004) (cant) 57000(.851 += M.S.

YIELD= 29916 F4 458.85715 6101KPa T= bt = (1.081 t-481 =(885 psi) C38000 (.91)] (g) M.S.

YIELD= Analysis of Bracket (41004819-006) -0.81R 2PL 0.33 2PL / DIMENSIONS IN CM Since we don't know which will react first, loads in bolts "A" or we will use total loads in "A" first and then total loads in " B " , "B" next. (conservative).

41004640 - Landing Gear, Active Controls Accumulator, Clamps & Brackets (cant) Analysis of Bracket (41004819-006) (cant) Loading Condition # 1 =.98182wFwD =634-6N F 2 (142.66336 lb) MAX Reacted at bracket as: 634.6~ R (camp. load into bracket) F =FMAX=(142.66336 lb) Y (F MAX) (3.705) 2261N (shear load on bolts) RFH= 1.04 =(508.23822 lb) Loading Condition #2 ]+.72817WFwD =1877N +.12067[WIN,BD +wINBD =-5WDOWN 2 (422.00053 lb) 2 1 MAX 302.7N =.25[WFWD +WFWD ]= (68.055 lb) 302.7N R =.25[WINBD +wINBD I= (68.055 lb) I4 1 2 Reacted at bracket as: 1877N (tension load on bolts "A" =R = F (422.00053 lb) shear load on bolts "B") D4 D4 (shear load on F =R +(' bolts "A" 6 ,,Bw) l~~~~~~:~)RD4=l~~~"23806 lb) F4 F4 l-40+.040 (tens.or comp,load on bolts "B" +tl o5 =R FI -.010'RD4=~:~;r,,353 lb) shear load on bolts "A") I4 - It is obvious from the above information that loading condition #2 is the more severe of the two and will be used for anlysis.

41004640 - Landing Gear, Active Controls Accumulator, Clamps & Brackets (cant) Analysis of Bracket (41004819-006) (coll't) Bolts "A" (assuming only 1 bolt reacts) * f \ fF' (each bolt) (tension); F (shear); F (shear) D4 F4 I4 1 - -24 bolt F(422.001) = 1.657x104KPa 0.566CM2 at= As(tensile stress area)= -0878 (-0878 in2) '(2403. psi) (F Total Shear Load= i/m (362.238)2+(652.363)2 =;;;;N186 lb) =d .

- -24 bolt Tc= 746.186 =6.359x104KPa 0.519CM2 A(shear stress area)= -0809 t-0809 in2) (9224 psi) $ +ds = 24203 +dw =7-241x104KPa =P = MAX (10503 DSi) Bolts "B" (assuming only 1 bolt reacts) / (shear;; FD (shear); FD (tension) (each bolt) I4 4 4 Total Shear Load= (422.001)2+(362.238)2 =;;;;N14855 lb) .

556.149 = 4.740x104KPa T= -0809 (6875 psi) 41004640 - Landing Gear, Active Controls Accumulator, Clamps & Brackets (con-t) Analysis of Bracket (41004819-006 (cant) (J = T(652.363) =2.561x104KPa t -0878 (3715 psi) ut % 37215 +dz =6.190x104KPa (,) +T2 = -

J

uP = 2+ MAX (8979 psi)

TMAX= +dF = t\/(3715)2+4(6875)2 =4-g10x104Kpa

(7122 psi) M.S. (either bolts "A" or "B") Bolts "C!"

(direct tension loads) F +F =F 4FC =FC =.5FD c1 c2 c4 12 4 (tension due to moment F . 75F =2.25FC c, C caused by FF ) I 4 z 1.91CM -YzA .

‘(CM) .75F +2.25F =l.O6F c1 c2 F4 . 75F +2.25(& )=i.06~ c1 1 F4 1.06FF -4 FC = =.70667F 1.50 F4 F =.23556F c2 F4 41004640 - Landing Gear, Active Controls Accumulator, Clapms & Brackets (cant) Analysis of Bracket (41004819-006) (cant) 2.91 (max tension due to moment F =F = 1 =3.54878F 2 (l.22-.81)(F14) caused by FI at "A") c1 c2 I4 (total loads) F =.5F +.70667F +3.54878F (tension) cl D4 F4 I4 =(.5)(422.001)+(.70667)(362.238)+(3.54878)(652.363) 1.238x104N =(2782.076 lb) (max tension) =.5FD +.23556FF +3.54878F FC2 4 4 =4 =.(.5) (422.001)+(.23556)(362.238)+(3.54878)(652.363) 1.162x104N =*m (total loads) FC ,=F,-&.=\/ (362.238)2+(652.363)2 (shear) (assume only 1 3319N bolt takes load) =(746.1862 lb) Tmax shear) The above are load requirements for attaching to aircraft structure (assume 2 attachments located approximately as shown on P. 83).

Top of Fitting Use rectangular plate, 3 edges simply supported, 4th edge free, uniformly loaded. (Timoshenko, "Theory of Plates & Shells", P. 211-215).

7.11CM 9.4MM b,3-10CM b= 43571 a=(2.80) (1.22) a . tMIN=(.37 in) (Mx)MAX=.060qa2 (table 42) F +F c1 c2 = (2782.076)+(2611.422) =1.08gx104KPa '= ab (2.80)(1.22) (1578.8929 lb/in2 41004640 - Landing Gear, Active Controls Accumulator, Clamps & Brackets (cant) Analvsis of Bracket (41.004819-0'06) (cant) =(.060) ,1578.8929,,2.8)2=j;;;N;;(; in lb,in) l"x)MAX .

6 (Ma) MAX = 6t742.7112) =2.244X105Kpa 'b= t2 .372 (32551 psi) 57000(.85) M.S.

32551 YIELD= Back of Fitting 4 edges simply supported, load uniformly Use rectangular plate, distributed over rectangular area (Timoshenko, "Theory of Plates t Shells, P. 158-161).

2.54CM ,,7.37CM v=.33 (2.90 in) u=(l.OO in) b,7.11CM 5.08~~ 5=1.04 (2.80 in) v=(2.00 in) (conservative) 305CM b t - 96552 k= % =2.00 MIN=(.12 in) 2 =.

- =.35862 d=dE=2.236 a (p=1.481 (from table 27) x=2 -6964 (from table 26) p=.110 Y=.374 2902N p=F = I4 (652.363 lb) lr5 4a Sin a M= & ((2) ln( )+X-(p) 7Td (l+Y)+(y+Q) (1-V) X c J 71C1.04) 652.363 4t2.90) Sin 2-go )+2.6964-1.481)(1.33) M= 8'IT X n(2.236) +(.110+.374)(.67) J Landing Gear, Active Controls 41004640 - Accumulator, Clamps C Brackets (cant) - ~-~ __ Analysis of .Bracket (41004819-006) (cant) ~ _- . ~~ Back of Fitting (cant) Mx=(25.95670) (1.49111)+2.6964-l-481)(1.33)+.32428 346.8N-M/M M= x (77.96088 in-lb/in Id 4a Sin a M=& 1+x-m ( 1+v1- w-L+*/) ( 1-VI ((2)ln ( 7Td Y My=(25.95670) (.79905*2.6964-l-481)(1.33)-.32428 =:;;.;;;t{Min-lb,in)

C 1

.

=2.239x105KPa 6MX 6t77.96088) X u=t2= -12 (32483 psi) 57000(.85) M.S.

YIELD= 32483 Comnression Bucklina The only way for above bracket to be in compression is for RD to be instead of "down". 4 "UP " R +.I2067 (Ref. P. 78) =.5wup +.72817WFwD INBD2 D4 2 2 126.915+145.305 +(.72817)(145.305) =(.5)(203.427)+(.12067) C 1069N =(240.36903 lb) 1069N F =(240.369 lb) D4 41004640 - Landing Gear, Active Controls Accumulator, Clamps 6 Brackets (cant) Analysis of Bracket (41004819-006) (cant) Compression Buckling (cant) 675+.050 1048N F =68.055+ '; 05-.olo) (240.36W=(235 620 lb) .

.

F4 Less than loads on P-84 - Do not need to analyze shears.

'.40+-040)(240.369)=;;;;N874 lb) =68-055+ (1.05-.OIO .

FI (from P. 84) 1726N Total Compression Load = R +F =142.663+240.369=(383.032 lbl FV D4 Crippling allowable (Bruhn, Page C7.1) a+b l-25+1.40 =11 042 -= 2(.12) - 2t (from Fig. C7.3) Fee --j(-0.305CM - = -057 (one edge free) TYP 2.776x105KPa F ..=.05 57000)(.85)(10.3x106) = (40266 psi) 4.1CM2 Area= 2.80+2(1.25) (.12)= C (-636 in2) 383.032 =415OKPa U= C . 636 (602 psi) M.S.= F -I=p] 5.738N 41004640 - Landing Gear, Active Control (1.29 lb) Bracket, Shutoff Valve - 41004819-009 A/C (3.99 lb) (.48 1 22.55N (5.07 lb) 17.75N 2.135N 2.135N 16.06N 2.669N (3.99 lb) (-48 lb) 1 f-48 lb) (3.f 51 lb) (.60 lb) C 3.60'3N 'R3 (-8' 1 lb)

I

p-2.306M (90.80 in) z p-2.3799 (93.66 in) * 2.3831 (93.85 in)

I-

P (3.99) (36.00)(38.19)(74.19+36.00) ~A(74.19)+2MB(74.19+19.66)+jiC(19.66)= - 74.19 (-48) (72.19)(2.00)(74.19+72.19) _ (-48) (2.00)(17.66) (19.66+17.66) 19.66 74.19 _ (l-03)(7.22) (12.44) (19.66+12.44) _ (3.61) (13.08) (6.58).(19.66+6.58) 19.66 19.66 (.81)(16.61)(3.05) (19.66+3.05) _ (.60) '".":;';;" (19.66+.19) 19.66 .

l87.7MB=-8147.~531-136.73656-32.l8252-15l.O4847-4l4.6877-47.4OlO7 -2.24103=-8931.7504 -5.376N-M M~'(-47.58525 in-lb) 41004640 - Landina Gear. Active Control Bracket, Shutoff Valve - 41004819-009 (cant) 22.55N (5.07 lb) 17.75N 2.135N 2.135N 16.06N 2.669N (3.99 lb) (-48 lb) (.48 lb) (3.61 lb) (-60 lb) (0 in-lb) 5.376N-M 'd(O in-lb) 4.5818 3.603N h Ni (47.58525 in-lb) (1.03 lb) (-81 lb) 9.136N 11.281 11.28N 0.2172N (2.535 lb) (.04883 lb) (2.053891b) (2.535 lb) t tt t 0.05756 N 2.0788 1.918N 1.683N (-01294 lb) (.46706lb) (.43117 lb) t.37826 lb) t tt t 2.853N 8.612N 2.899N 10.68N (.641401b)(1.93611 lb) (-65174 lb) (2.40177 lb) t 1 tt 2.85313 5.374N 3.0448 (-64140 lb) (1.20823 lb) (-68434 lb) tt t 0.589N 2.643~ t-12566 lb) t.59420 lb) t t 10.77N 0.02581 (.00580 lb) (2.42041 lb) t 10.77N (2.42041 lb) t 6.340N 57.64~ 7.5041 %-(1.42543 lbs) R2-.(12.95758 lbs) R3-(l.68699 lbs) (check) R1+R2+R3=16.07=3.99+.48+5.O7+.48+l.O3+3.6l+.8l+.6O 57.65~ ,',max Wt on bracket = (12.96 lb) This Wt combined with the required "g" factors will determine max loading conditions to analyze bracket.

Landing Gear, Active Controls 41004640 - 41004819-009 (co.nt) Bracket, shutoff Valve - I I

(?I

i

I

7.62

u I

l--II n !F

. U. I"" -- I l ., W W INBD FWD I I I I 6.985 I W DOWN

I I

DIMENSIONS IN CM 86.47~ W =w AFT=1.5R = 2 (19.44 lb) FWD 337.3N W DOWN=5.85R = 2 (75.82 lb) 121.1N W UP=2'1R2.=(27.22 lb) 86.47N =W W =1-5R2=(19.44 lb) INBD OUTBD (moment at (A)-(A)) M~A)=2.39WDOWN+3.00WAFT=~~~~6~~Min-lb) .

160CM 3.353CM b=2.32-(2x.50)=(,.32 in) t=;.063 in) 6(239.53) _1.891x106KPa (274319 psi) (l.32)(.06312 ,111 - . ._.. -_-.--- ._---.-.. - ..__. - ___.__ - Landins Gear, Active Controls 41004640 Shutoff Valve - 41004819-009 (cant) Bracket, 6.589N-M =3.00WINBD =(3.00(l.9.44)=(58~32 in-lb) T(A) 3T(A) [,,.6 3=3(58.32) I+.6 S] =2.368x105KPa MAX= bt2 (1.32) L063)2 (34351 psi) This is not a structural part. For location purposes only. To make part structural, see below.

6 (239.53) 477CM

=;.188 in) (35000) (.88)(1.32) 508~~ To be structural, bracket must be either-' thick or must be t.20 in) designed to include side gussets capable of taking sufficient compression load.

APPENDM B SYSTEM SPECIFICATION - DESIGN SPECIFICATION, FLIGHTWORTHY ELECTRO-HYDRAULIC ACTIVE CONTROL LANDING GEAR SYSTEM FOR A SUPERSONIC AIRPLANE 1.0 SCOPE This document establishes the requirements and defines the design objectives for an,electro-hydraulic active control landing gear system for a supersonic aircraft, based on a modified landing gear.

2.0 APPLICABLE DOCUMENTS The following documents form a part of this specification to the extent that they are applicable: Contract NASl-15455 issued by NASA Langley Research Center, Hampton, Virginia MIL-STD-810C - Environmental Test Methods MIL-STD-461 - Electromagnetic Interference Characteristics MIL-E-5400 - Electrical Equipment, Airborne General Specification MIL-STD-454 - Standard General Requirements for Electronic Equipment MIL-STD-275 - Printed Wiring for Electronic Equipment QQ-A-325 - Aluminum Alloy Sheet MIL-S-19500/X - Semiconductor Devices, General Specification MIL-S-5541 - Chemical Film Finishes MIL-G-5514 - Packings, Installation and Gland Design Hydraulic, General Specification for MIL-I-6866 - Inspection, Penetrant Methods of MIL-S-8879 - Screw Threads, Controlled Radius Root with Increased Minor Diameter, General Specification for MIL-I-6868 - Inspection Process, Magnetic Particle MIL-H-27601 - Hydraulic Fluid, Petroleum Base High Temperature, Flight Vehicle MIL-R-83248 - Type, Class 1 Rubber, Fluorocarbon Elastomer High Temperature Fluid and Compression Set Resistant O-Ring MS33540 - Safety Wiring and Cotter Pinning, General Practices for MS33649 - Bosses, Fluid Connection Internal Straight Thread Other Publications For requirements not covered above, materials, processes, and standard products shall be selected in accordance with specifi- cations or standards from the sources indicated below and in the order of precedence shown: 1. Federal specifications and standards as listed in the Index of Federal Specifications, Standards and Handbooks published by the General Services Administration.

2. Military specifications and standards as listed in the Department of Defense Index of Specifications and Standards.

Industry Specifications and standards as listed in 3.

indexes published by recognized industrial associations but not limited to, the following without including, order of precedence: a. National Aerospace standards (NAS) as published by the National Aircraft Standards Committee of the Aerospace Industries Association.

b. Aerospace Materials standards (AMS), Aerospace .

Standards (AS), Aerospace Recommended Practices (ARP) and Aerospace Information Reports (AIR) published by the Society of Automotive Engineers.

3.0 REQUIREMENTS The active control landing gear system shall be designed in accordance with the requirements of this specification. The equipment shall meet the performance requirements when installed on the aircraft during conditions of landing, taxi, and take-off, and under the conditions of paragraph 3.1.2.1.

3.1 System Function and Definitions 3.1.1 System Function The ACLG shall operate as a closed loop to command the wing gear interface force to the level of the generated limit force.

3.1.2 Item Definition (FigureB-1) The active control main landing gear system (ACLG) ofthe supersonic airplane is a dual system, each set of which consists of a modified an accumulator, strut, a servovalve, an electronic controller, cockpit-mounted sink rate selector, and interface hardware tieing into existing hydraulic and electrical supply systems, and into existing aircraft mounted sensors, all of which are defined in 3.1.2.2.

System Conditions 3.1.2.1 The ACLG shall meet its performance requirements when the system parameters are as follows: Charging Gas GN2 7.375 x 10e4m3 (45 in3) Fully compressed gas volume Gas pressure, extended (70 OF) 1930 KPa. (280 PSIG) 0.0156m3 (952 in3) Oil volume (MIL-H-27601A) Strut Stroke 0.508m (20 ins. ) 121°C (250’F) Maximum oil temperature Minimum oil temperature -40 ‘C (-40’F) 1.512 x 105N (34,000 lbs) Design landing load Design maximum sink rate 3.048m/sec (10 ft/sec) Maximum landing load 2.335 x 105N (52,500 lbs) 2758KPa (400 psig) Tire pr.essure (unloaded) Touch down velocity 93.27m/sec (306 ft/sec) 3892N (875 lbs) Unsprung weight Wheel well temperature -54 to 121’C (-65’to 250’F) Gear position down at all times Associated Sensors 3.1.2.2 Wing/Gear Accelerometers Range - * 4.12 g’s Scale factor - 0.002 v/g @ 5vdc excitation Strut Hydraulic Pressure Transducer Range - O-l.72 x 104KPa. (o-2500 psi) Scale factor - 0.00232 mv/KPa (0.016 mv/psi) Strut Stroke Transducer Synchro Tne Range * 0.254m (* 10 ins. ) Scale Factor - 19.69 VRMS/m (0.5 VRMS/in.)

Scissors Switch open or closed (takeoff or landing mode indicator) Q-w Wheel Generator D. C. generator Tme Range O-2400 rpm (156 knots) Scale Factor - 19.6 mv/rpm 3.1.3 Interface Definition The landing gear system shall be capable of operation when system parameters are as listed below, when powered by the existing aircraft hydraulic and electrical systems, and when interconnected with cockpit controls and air- craft mounted sensors as listed in paragraph 3.1.2.2.

3.1.3.1 Hydraulic System The ACLG shall meet the requirements of this specification when used with an aircraft hydraulic system which is capable of supplying a pressure -4 m3/min. To -4 kPa (3350 psig) and a maximum flow to 8.194 x 10 of 2.31 x 10 provide the required transient flow the ACLG shall incorporate a 0. 0265m3 (7 gallon) accumulator for each of the two landing gear struts.

3.1.3.2 Electrical System The electrical power available on the aircraft consists of 28vdc, 26vrms. 400 Hz, and 115 vrms, 400 Hz in accordance with MIL-STD-704A.

3.7 COCKPIT CONTROLS AND INDICATORS The cockpit control panel shall include a power switch, a sink rate selector (to be used in lieu of a sink rate sensor), a test button and status indicator lamps.

STRUT REQUIREMENTS 3.2 The following sections define the configuration and requirements of the landing gear struts.

Strut Modifications 3.2.1 The struts shall be modified so that the hydraulic chambers at the lower end of the strut will connect to the control port of the servovalve thereby allowing the servovalve to port hydraulic fluid into or out of the strut as required to control the The basic gear structure shall remain wing gear interface force.

intact.

3.2.2 Materials Materials used shall be in accordance with the applicable military specifications and shall be compatible with those in the aircraft and landing gear.

3.2.3 Pressure The hydraulic pressure in the strut shall not exceed 2200 Relief valves shall be used to meet this requirement.

psi.

3.2.4 Installation Hardware Hardware such as supporting brackets for hydraulic iines shall be installed in the wheel well at locations agreed to by the aircraft manufacturer.

ACCUMULATORS 3.3 Accumulators shall be used for each strut to supply the necessary transient flow.

3.3.1 Volume The total volume of the accumulator for each strut shall be 0.0265m3 (7 gallons).

3.3.2 Location The accumulators shall be installed in the aircraft wheel well in a manner approved by the aircraft manufacturer.

3.4 SERVOVALVE A servovalve shall be used for each strut and shall meet the following requirements.

3.4.1 Performance The performance requirements are shown in the Appendix 3.4.2 Envelope The servovalve envelope is shown in HR drawing 23241510 3.5 CONTROLLER The controller shall be used as an integral part of the closed-loop servosystem which controls the aircraft wing/gear interface force during landing, taxi, and take-off. The controller shall accept sensor data, perform computations involving energy and momentum, effect prescribed control laws and provide an output current to the servovalve.

The controller shall also contain the electronic circuitry necessary to test the system for failures and cause a reversion to a passive gear configuration if a failure exists.

3.5.1 Requirements Input-Output Requirements 3.5.1.1 The input/output signal requirements are shown in block Primary inputs and outputs are those diagram form in FigureB-2.

Secondary inputs signals required to perform the control function.

and outputs are those signals used for testing and status indication FigureB-3' defines but which do not influence the control function.

the sign conventions and transducer polarities.

3.5.1.1.1 Primary Input Signals The primary input-output signals shall be those shown in TABLE B-I.

3.5.1.1.2 Secondary Input-Output Signals Secondary input-output signals shall be as specified in TABLE B-II and&III respectively.

Functional Controller Requirements 3.5.1.2 The controller shall have three basic functional requirements as shown in FigureB-4. These are: Operating mode determination, 1.

2.

Charging pressure regulationand limit force command determination, 'Control law implementation 3.

3.5.1.2.1 Operating Mode Determination The controller establishes the operating modes of the active control landing gear system upon application of electrical power and through its acceptance of condition states from sensors (3.1.2.2).

The controller's "state" is defined in terms of "enable function" and "modes".

3.5.1.2.1.1 Enable Functions The enable functions of the system (controller) are 'Controller Enable" "Servoloop Enable", and "Integrator Enable".

3.5.1.2.1.1.1 Controller Enable "Controller-Enable" is defined as that condition which allows the controller to perform calculations, This condition occurs when power is applied and the test confirms system integrity.

3.5.1.2.1.1.2 Servoloop Enable is defined as that condition which "Servoloop Enable" Under allows the controller to provide current to the servovalve coil.

occurs when the kinetic "Servoloop Enable" landing conditions, Under energy of the aircraft equals the work potential of the strut.

take-off or taxi conditions it occurs upon application of power.

3.5.1.2.1.1.3 Integrator Enable is defined as that condition which "Integrator Enable" which generates the wing-gear velocity from allows the integrator, It occurs at touchdown. This the wing-gear acceleration,to function.

enabling function is necessary in order to prevent integrator drift.

3.5.1.2.1.2 Modes The modes of the system (controller) are "Landing", "Take-off" and "Test".

3.511,2.1,2,1 Landing Mode -- The landing mode is selected by the controller when power is applied,fhe scissors switch is open and a successful test has been completed.

The controller shall then commence computation of kinetic energy and strut work potential in the manner shown in Figure B-5.

The controller shall not enable the servoloop until the kinetic energy is less than the work potential of the strut.

The landing mode encompasses several phases, each imposing a different functional demand on the controller.

These phases are: 1. passive phase 2. impact active control 3. transition 4. rollout and taxi 3.5.1.2.1.2.1.1 Passive Phase In the passive phase the controller shall: Close an auxilliary pressure loop to maintain the strut pres- 1.

sure at its charging value.

2. Compute kinetic energy of the aircraft, work potential of the strut and compare these values.

3. Sample and hold value of the wing-gear interface force (wing-gear acceleration).

3.5.1.2.1.2.1.2 Impact Active Control The impact active control phase shall commence when the energy comparison indicates that the work potential of the strut equals or ex- Upon such occurrence the controller ceeds the kinetic energy of the aircraft.

shall: 1. enable the servoloop 2.

discontinue energy computations 3. maintain a constant limit force 4. deliver a current to the servovalve in accordance with the control laws 5. calculate the velocity at which transition is to commence, in accordance with the relationships shown in Figure B-6, and compare this to the actual velocity to determine the starting point of transition.

3.5.1.2.1.2.1.3 Transition The transition to the rollout phase shall commence when the wing/gear interface velocity becomes equal to the transition velocity.

During transition the controller shall: 1. Linearly decrease the limit force command to a predetermined mjnimum force (F ).

min 2. Maintain active control about Fmm as long as F is greater w than Fmin or less than -F min.

Set the limit force command to zero and disable the force 3.

loop when F becomes less than Fmm.

w 3.5.1.2.1.2.1.4 RolloutandTaxi -- The rollout phase shall commence when the limit force During the rollout phase the command in the transition phase reaches zero.

controller shall maintain active control with the limit force command equal as long as F is greater than Fmm or F is less than -F For to Fmin min’ w wg the limit force command values of F less than Fmm or greater than -F min, w The controller shall remain in shall be set to zero and the force loop disabled.

the rollout mode until takeoff occurs or power is removed.

Takeoff Mode 3.5.1.2.1.2.2 The controller shall automatically select the takeoff mode of operation when: 1. Power has been applied.

2. The scissors switch is closed.

During takeoff the limit force command shall be zero and active control maintained about this limit force command.

Limit Force Command Computation 3.5.1.2.2 The limit force shall be the command to the servoloop and shall serve as the desired wing/gear interface force. The limit force command shall be generated according to the requirements of each mode and phase as described below.

Landing Mode 3.5.1.2.2.1 3.5.1.2.2.1.1 -PassivePhase Commencing with controller enablement and continuing until the servoloop is enabled the effective limit force command shall be zero.

Impact Active Control Phase 3.5.1.2.2.1.2 From the time the servoloop is enabled until the start of the transition phase the limit force command shall be constant, and equal to the value of the wing/gear interface force at the time the servoloop is enabled.

Transit ion 3.5.1.2.2.1.3 The limit force shall linearly decrease at a specified rate, from its value at the start of transition to Fmin at as shown in Figure B-7, which time FLC will be set to zero and the force loop disabled.

3.5.1.2.2.1.4 Rollout and Taxi During the rollout and taxi phase the limit force command shall be zero or f F min’ Limit Force Command - Takeoff Mode 3.5.1.2.2.2 In the takeoff mode of operation the limit force command shall be zero.

3.5.1.2.3 Control Laws The controller shall implement the control laws shown in Figure B-8 and the transfer functions of TABLE B-IV.

3.5.1.2.4 Pos it ion Loop The ACLG shall incorporate a low response position loop for the purpose of returning the’strut to its static position during the rollout phase of the landing.

3.5.1.2.5 Auxiliary Pressure Loop An auxiliary pressure loop shall be incorporated for the purpose of setting and maintaining the static hydraulic pressure of the gear.

Whenever the servoloop is enabled’the ausiliary pressure loop shall be disabled.

3.5.1.2.6 Dynamic R.equirements The loops of the ACLG shall meet the dynamic requirements of this section.

3.5.1.2.6.1 Force Loop The force loop shall meet the following requirements: Amplitude : Flat, to within 3 db, to TBD. Peaking shall not exceed 3 db .

Phase: Not to exceed 90” lag at TBD.

3.5.1.2.6.2 Posit ion Loop The position loop shall meet the following requirements when operating with the force loop closed: approximately 0.1 Hz, peaking not to Peaking frequency: exceed 5 db.

Phase: approximately 90° lag at the peaking frequency.

These parameters are not critical.

(See Schematic, Figure B-9) Design 3.5.2 The controller shall be designed to meet all the requirements of this specification and shall provide safe and reliable operation.

3.5.2.1 Computational and Control Law Implementation All computations and implementation of transfer functions shall be accomplished by means of analog circuitry in order to minimize the physical size of the controller and to optimize the controller response.

3.5.2.1.1 Limit Force Command The limit force command signal shall be the wing-gear interface force during the passive phase. This signal shall be applied to a sample and hold circuit, but since the servoloop is disabled during the passive phase the effective limit force command is zero. At the start of the impact active control phase, the input to the sample circuit shall be disconnected thus holding the limit force command signal constant at the last sampled value of the wing-gear interface force. During the transition phase the hold circuit shall decay in a linear manner to decrease the limit force command to a preset minimum force and for values of the wing-gear interface force less than F or greater ‘FmirI min then -Fmin the limit force command shall be set to zero.

3.5.2.1.2 Wing Gear Velocity Computation The wing gear velocity signal shall be generated by integration of the signal from the wing/gear accelerometer. Means shall be provided for enabling the integrator at the time the controller is enabled and disabling it when the controller is disabled.

3.5.2.1.3 Energy Computations 3.5.2.1.3.1 Kinetic Energy of the Aircraft The signal representing kinetic energy of the aircraft shall be computed by mathematically squaring the wing/gear interface velocity signal by means of an analog multiplier and attenu- ating it by one half as shown in Figure B-10.

3.5.2.1.3.2 Work Potential of the Strut The signal representing work potential of the strut shall be computed by subtracting the strut stroke signal from a constant signal representing maximum stroke and multiplying this signal by the wing/gear acceleration signal, using an analog multiplier as shown in Figure B-10.

Transition Velocity Computation 3.5.2.1.4 The transition velocity shall be computed by using an analog multiplier to mathematically square the signal representing the limit force command during the impact active control phase, and attenuating it by a constant which represents the reciprocal of the product of twice the aircraft mass per gear and the transition decay rate (R) as shown in Figure B-10.

3.5.2.2 Comparisons Comparisons for mode, phase determination and failure detection shall be accomplished by analog comparators driving analog switches.

3.5.2.3 Servovalve Driver The power stage of the controller shall be capable of supplying 250 ma into a 200 ohm load. 1t“s output impedance shall exceed 100 kilohms.

3.5.2.4 Controller Inputs (See Figure B-11) 3.5.2.4.1 Initial Sink Rate The controller shall be so designed that it can accept a signal from a sink rate sensor or a signal from a variable source representing sink rate which is set by a control on the front panel ofthe controller or a control inthe cockpit.

3.5.2.4.2 Servovalve Bias The servovalve bias command shall be generated by a Control which shall be capable of providing a DC voltage of 0 to f 10VDC.

3.5.2.4.3 Strut Position Command The strut position command as indicated in Figure B-8 shall be derived from a control which is capable of providing a DC voltage of 0 to f 10 VDC. This signal determines the static position of the strut, Strut Position Bias Signal 3.5.2.4.4 as indicated in Figure The strut position bias signal, It is required in order B-8, shall be variable from 0 to 2 10 VDC.

to produce a null signal when the strut is at its design static deflection.

3.5.2.4.5 Sensor Inputs The controller shall accept inputs from the following sensors: Wing/gear accelerometer Strut position sensor Strut hydraulic pressure transducer Characteristics of the signals shall be as described in 3.5.2.4.6 Wing/Gear Accelerometer Bias Signal The wing/gear accelerometer bias signal as indicated in Figure B-9, shall be variable from 0 to + 10 volts. It is required to produce a null signal when the accelerometer is mountedto output a signal equivalent to a positive one G-unit.

3.5.2.4.7 Top Panel Test Inputs Provisions shall be included for inserting test signals into the controller by means of top panel jacks as shown in Figure B-12 3.5.2.5 Front Panel Outputs (See Figure B-13) Buffered test outputs shall be available at front panel jacks as shown in Figure B-12.

3.5.2.6 Rear Panel The rear panel of the controller shall contain the connectors, suitably labeled, as shown in Figure B-14.

Servovalve Current 3.5.2.7 The controller shall provide an output current to the servovalve. The characteristics of the servovalve driver shall be as defined in paragraph 3.5.2.3.

3.5.2.8 Flight Safety The controller shall incorporate means for augmenting the safety of the aircraft as defined in the following paragraphs.

3.5.2.8.1 Built-in Test The controller shall incorporate means for determining the functional integrity of the system prior to landing or take-off.

The test shall consist of the following: 3.5.2.8.1.1 Pre-Land When a cockpit mounted test button is depressed a simulated acceleration (wing/gear force) and strut compression shall be applied to the controller. The servovalve stroke signal shall be compared to the required signal at a point in time when the limit force command is approximately half-way down the ramp during the transition phase. A difference in these signals,'in excess of a thresh- old value shall constitute a failure. Otherwise a green "Test in Progress" lamp in the cockpit shall be illuminated during the test and‘extinguished upon successful completion of the test.

3. 5. 2. 8. 1.2 Accelerometers The accelerometer shall be tested by monitoring the output to determine if its signal is in the range of 0 g's.

An incorrect signal shall constitute a failure.

3. 5.2. 8. 1. 3 Strut Stroke Transducer (Synchro) The synchro circuitry shall incorporate a self-test feature to detect failures. It shall be accomplished by applying a voltage across the unusedwinding and each of the usedwindings and determining if a.currentflows.

Auxiliary Pressure Loop 3.5.2.8.1.4 The auxiliary pressure loop shall be tested by A signal in excess of a threshold value monitoring the loop error.

which exists for a fixed amount of time shall constitute a failure.

3.5.2.8.2 Passive Reversion An indication of any failure, as described in paragraph 3.5.2.8.1 in either channel shall cause currents to be latched out from both servovalves which will cause the spool centering springs to null the valves andenergize a solenoid valve to prevent flow into or outofthe struts.

An amber light in the cockpit shall be illuminated to indicate that the gear is in a passive state.

3.5.2.9 Power Section The power section shall accept 115 V RMS 400 He, 26 V RMS, 400 He and 28 VDC power from the aircraft power system and shall generate the DC voltages necessary for controller operation as shown in Figure B-15.

3.5.2.10 External Electrical Connectors Connectors used for interconnection with the transducers, servovalves, and cockpit signals shall be arranged as shown in Figure B-16.

3.5.2.11 Parts Electronic parts used in the controller shall be of military specification quality.

3.5.2.12 Packaging The controller shall be contained within a single enclosure suitable for mounting in the electronics bay of the aircraft. Modular construction shall be employed to the extent necessary to provide easy access for maintenance purposes.

K

Fo wi

:ONTf?OuE R --.--.

D

L - -4

t ___. _ . ..- _.._ _^._^_.. __--_. .__---. -.---.-.-A.

FIGURE B-1. ACTIVE CONTROL LANDING GEAR. SYSTEM

_ . __. ~-.---

RIss=Rm&s *INITIAL SINK RATE -PRIMARY OUTPUTS ._-... *.__--..--A.

.BUILT-IN TEST l SERVOVALVE COMMANDS #W/G ACCELERATION CONTROLLER --LA---- l STRUT POSITION (2 CHANNELS) 0 SINK RATE (IF USED) dlYDB_A_~U: P-E I ; SECONDARY OUTXJlJs ES.ONT..EIAN& . ANALOG MONITOR

/p----b

. POINTS SECONDARY INPUTS

l EVE.NTf STATUS

FRONT PANEL TEST DISPLAY -VISUAL a DISCRETES INPUT$ . ANALOG FIGURE B-2. INPUT/OUTPUT SIGNAL SCHEMATIC WING-GEAR INTERFACE T- p +- W/G ACCELEROMDER -II-- AC DC Volts ------? Volts Demodulator I STR IJ_T POSITION I 1 SENSOR PRESSURE 0 PSIG= OV + PSIG= +V M (ins: VOLTS FULL EXT 0 0 FULL RET +O. 508 +10 (+20) FIGURE B-3. SIGN CONVENTIONS AND TRANSDUCER POLARITIES CHARGlNG PRESS- OPERATING CONTROL LAW I ELEC.

URE REGULATION MODE DETER- IMPLEMENTATION POWER COMMAND DETER- MINA TION -?

MINATION

I

l LANDING @CONTROL LAWS

’ 1

.ENERGY .IMPACT l POSITION DISCRETES COMPUTATIONS ACTIVE *FORCE l SINK RATE I @LIMIT FORCE CONTROL l TOUCHDOWN -D . SERVOVALVE INDICATION COMMAND l TRA NSITION COMMANDS rSERVOLOOP @ROLLOUT ENABLE aTAKEOFF .INTEGRATOR ROLLOUT ENABLE l TEST CONTROLLER

L - -- ----

FIGURE B-4. BASIC FUNCTIONAL REQUIREMENTS Strut Work Potential = H (igwg) (%MAX 3)

= (+)(+)

(vs -/~WGd~t)2

Kinetic Energy Active control shall be initiated when: Strut Work Potential > Kinetic Energy Or, in equation form: Condition For AC tive Control Initiation where, ZWG = Wing-Gear Acceleration, g’s dt = Integral of Wing-Gear Acceleration ‘WG

J-

m/set (inlsec) with respect to time, = Initial Airplane Sink Rate, m/set (in/set) vS Maximum Strut Stroke, m (in) X MAX = m (in) = Strut Stroke, xS Gravitational Acceleration = 9.804 m/set = (386 in/set ) g FIGURE B-5 ENERGY RELATIONSHIP S -~- The wing-gear interface velocity at which transition occurs shall be computed as follows: Transition starts when VT becomes equal to the wing-gear interfa:? velocity 'T = FLIP (touchdown sink rate minus XWG .clt 2w R J -r

J

where, = Transition Velocity, m/set (idsec) vT = N (lb) Limit Force Command during FLI impact = Airplane mass per main gear 1.388 x 104Kg (30,600 lbm) = Limit Force Transition Rate 4.448 x 105N (100,000 lb/set as defined in Figure VII.

= Gravitational Acceleration 9.804 m/sec2 (386 in/sec2) FIGURE B-6 TRANSITION VELOCITY COMPUTATION The limit force command from,the beginning of the transition phase to the beginning of the rollout phase shall be computed as follows:

IF Fwg

I I

- RT FLT = FLI wg I

I

FD then IF F < I-

wg I

= 0 and the force loop ‘-LT = Wing-gear interface force is open =wg where, = Preset limit force FD Limit Force Command During Transition, N (lb) FLT = Limit Force Command Prior to FLI = Transition N (lb) R = Limit Force Transition Rate,4.448x105N/sec (100,000 lb/set) T = Time set FIGURE B-7 LIMIT FORCE COMMAND DURING TRANSITION PHASE WING-GEAR ACCELEROMETER

-.--_ - -_- I -_-- - - -----.- - -

I

I CONTROLLER (2CHANNEL)

- -- -

I

S.V. 13IAS CMlJ

I

I I

I

-r I I I I I GAIN & StiAPlNC

1 I

SYSTEM

I

- SERVOVALVE COMhlAND I

I

I LOGIC AIRCRAFT AND

’ I

LANDING GEAR

,I

I I.

- --I

- - --- - - -

I

I

STRllT POS CMD I I --- -- .---- _- _. -..I STRUT SENSOR (DEMODULATED) FIGURE ~-8 CONTROL LAW FUNCTIONAL SCHEMATIC ’ BIAS (lg) Gl

r;7

N/G ACCELERATION! j- W/G ACCELERATION ICOMPUTATI~N L,M,T + z l~*ppiig

d

I

I

+ VELOCITY ’ TO

I

G3 SERVO VALVE I

I

I

INTEdRATOR ENABLE

I

I

I

I

J

I

I

STRUT + POSITION t

I

STRUT

I

COMMAND HYDRAULIC PRESSURE CONTELLER

L-

------ ------ J

--- POWERT FIGURE B-9. CONTROLLER FUNCTIONAL SCHEMATIC REPRESENT AN ANALOG MULTIPLIER

F7l

REPRESENTS AN ANALOG COMPARA TOR zw, Xmax 4.

- c COMPARISON OF KINETIC ENERGY AND WORK POTENTIAL OF THE STRUT.

vs-vwg COMPARISON OF WING-GEAR VELOCITY AND TRANSITION VELOCITY FIGURE B-10. COMPUTATIONS AND COMPARISONS

FROM

WIG

b

- ACCEL

i

L I

*-I--

+

I

6,

A--+ -

I

I

i + 1 Strut

Pressu e

A4 FROM

STRUT POSITION SENSOR

1 r

r

-..L.---._.- ._.-_- ---- .-._.-

sEFivoLcoP INTEG LIMIT STRUT ENABLE ENABLE POSITION FORCE CMD ACCEL FRONT PANEL TEST INPUTS --__ .--I FIGURE B-11. FRONT PANEL INPUTS FIGURE B-12 CONTROLLER FRONT PANEL SKETCH ENABLE - - -- -.. - --,--- --I SIGNAL ’ I 1 -3 -4-1 I COMPIJ

I

SV BLAS COM I I

--

SENSORS W/G ACCEL

c

-- STRUT POsITiON STRUT PRESS-HYD FORCE LIMIT SERVO- POS STRUT INTEG STRUT SV WIG W/G ENABLE POS CMD PRESS ACCEL ERROR “c”;” E;AT[E ERRoR YEL HYD FRONT PANEL OUTPUTS FIGURE B-13 FRONT PANEL OUTPUTS TRANSDUCER INPUT OUTPUTS INPUT POWER

l-------l

CONTROLLEI ENABLI STRUT STRUT WIG S.V.

PRESS.

POS.

ACCEL CMD HYD \ 400 Hz / FIGURE B-14 . CONTROLLER REAR PANEL SKETCH

I

I

l 115 VAC 400 HZ

(

l 26 VAC 400 HZ +, 15VDC l 28 VDC

I

POWER CONTROLLER

I

SECTION CIRCUITS +, 5 VDC

I

28VDC TRANSDUCERS

I

. ACCEL I1SVDC

I

l STRUT POS.

426VAC . PRESSURE

I

I

- - .~.I - . . - - - - . - . - - - - - - FIGURE B-15. POWER SECTION SCHEMATIC ACCELEROMETERS PRESSURE TRANSDUCERS

rJ

/I 28 VDC PWR RTN -- OUTPUT ---- SIG RTN CONNECTOR: TBD i CONNECTOR: TBD STRUT POSITION SENSOR SERVOVALVE CMD RTN i_..i CONNECTOR: TBD CONNECTOR: TBD ELECTRICAL CONNECTORS FIGURE B-16. EXTERNAL TABLE B-I PRIMARY INPUT SIGNAL SPECIFICATIONS SIGNAL TYPE SOURCE :POLARITY PARAMETER UNITS RANGE Accelerometer z4.12 g's Analog ling-Gear Acceleration See Fig.

g's 3.1-3 Meters l-0.508 m See Fig.

Synchro itrut Position (inches) Analog 10 to 20 in.) 3.1-3 Meters/set Front Panel Control or I-2.54 m/s -- :nitial Sink Rate @nches/sec > Analog Sink rate sensor 10to 100 in/ se 1 I itrut Pressure- Yolts Pressure Transducer 0-1.728$.x; o psI= o Analog Hydraulic (o-2500 psi) ,~~~:~,~ $ Touchdown indication O-2400 RPM Rotation volts Logic Wheel Generator = + volts volts Logic Aircraft scissors switch 0 or15VDC 15 VDC= Takeoff or Land Weight on gear.

TABLE B-I I SECONDARY INPUT SPECIFICATIONS SIGNAL TYPE SOURCE PARAMETER UNITS POLARITY SCALE FACTOR RANGE i-N = - Accel. 1,102~10-~ V/N Front Panel , 846x105N Limit Force lnalog

<ifs>

200,000 lbs) (4.902 x 10 -5 V/lb) Jack Command 2 MV/g @ 5 vdc Wing-Gear inalog See Fig. 3.3

g’s f5g

exe itat ion.

Acceleration l-2.54 m /set See Fig. 3.3 3.937 vlmlsec

Wing-Gear glT3%> Analog

Velocity 0 to 100 in/se 0.1 Volts/in/set) meters See Fig. 3.3 hnalog I-.508 m 19.68 v/m Strut Position gnches) 0 to 20 in) (0.5 volts/inch) -- -- Servoloop Volts Logic c*> Level Enable -- Logic Integrator Volts (*I Level Enable * Lo ic l= 0 tc 01 +2 Lo I TABLE B-III SECONDARY OUTPUT SPECIFICATIONS -.-.__--- SIGNAL FRONT PANEL --

r

SC JRCE RANGE , SCALE FACTOR PARAMETER UNITS rYPE ’

T

iack 'isus;splay -A Analog x Servovalve Command ma f 10 v l-176 ma/(V Force Error) X Analog f 10 g 1.8 V/g Wing-Gear g’s Acceleration 0 to ,508 m ‘9.843 v/m (0 to 20 in) (0.25 V/in) Analog X Strut Position 0 to ,508 m / (90. f;$j;/nm) (0 to 20 id i .

Analog X Strut Position Error Pa psi X Strut Pressure- Analog m Hydraulic 1.102x10 -5 V/N j. 896x105N :4.902 x 10-s V/lb) (200,000 lb) Analog X Force Error $s, 1.102~10-~ V/N 3. 896x105N 4.902 x lO-5 V/lb) (200,000 lb) X Limit Force Command (ifs) Analog -2.54 mlsec 13.937 vlmlsec to 100 in/se ~1’0.1 V/in/set) X Wing-Gear Velocity gE$ Analog -- -- Logic X 1 Servoloop Enable -- Level Logic -- -- Integrator Enable -- X Level -- -- r- X Takeoff Mode Visual -- -- X Landing Mode Visual TABLE B-IV CONTROL LAW TRANSFER FUNCTIONS SYMBOL I EF. FIGURE C-8 TRANSFER FUNCTION PARAMETER VALUES I = 1.0 v/v

G1 KWG KWG

(S2+2s2Wl S+W12)(TlS+l)(T3S+1)KA G2 = 0.0281 set T1 (S2+231 WlS+W12)(T2S+l)(T4S+1) = 0.0141 set T2 = 0.001 set T3 = 0.0001 set T4 = 251.2 rad/sec w1 = 0.1 s2 = 5.1 Sl KA = 176 ma/v nominal (variable from 50% to 200% of nominal) = 0.00098 V/V KF KF = 0.1 set TFS + 1 TF -1 = J,ap&xe Ooera tor -set XIII. REFERENCES 1. Ross, Irving; and Edson, Ralph: An Electronic Control for an Electrohydraulic Active Control Aircraft Landing Gear.

NASA Contractor Report 3 113, 1979.

1. Report No. 2. Government Accession No. 3. Recipient’s Catalog No.

NASA CR-3298 5. Report Date 4. Title and Subtitle Flightworthy Active Control Landing Gear System for a Supersonic Aircraft 7. Author(s) 6. Performing Orgamzation Report No.

Irving R.oss 10. Work Unit No.

9. Performing Organization Name and Address Hydraulic Research Textron Valencia, California 91355 L 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Contractor Report National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, LX 20546 15. Supplementary Notes Langley Technical Monitor: John R. McGehee Final Report 16. Abstract Hydraulic Research, under NASA Contract NASl- 15455, designed a flightworthy active control landing gear system for a supersonic aircraft, the purpose of which is to minimize aircraft loads during takeoff, impact, rollout and taxi .

The design consists of hydromechanical modifications to the existing gear and the development of a fail-safe electronic controller. Analytical results indicate that for an aircraft sink rate of 0.914 m/set (3 ft. /sec.) the system achieves a peak load reduction of 3 6% during landing impact.

7. Key Words (Suggested by Author(s)) 16. Distribution Statement Aircraft Landing Gear Unclassified - Unlimited Electronic Controls Active Controls Landing Loads Subject Category 05 22. Price’ 9. Security Classif. (of this report) 20. Security Classif. Iof this page) 21. No. of Pages Unclassified $7.25 Unclassified 139 * For sale by the National Technical Information Service, Springfield, Virginra 22161 NASA-Langley. 1980