(2)  
Remove access panels 9512 and 9514 (Ref Chapter 12, Access Doors and Panels).

C.  Adjust Rudder Power Control Unit
(1)  
Install rigging pin R-5.

(2)  
Check location of rudder trailing edge with respect to rudder index tab (Fig. 501). Place a straightedge against side of rudder at trailing edge. If both sides of trailing edge fall within width of rudder index tab, tighten both checknuts on main input (vernier control) rod  (Fig. 502). Combined total number of visible threads on sleeve and rod end shall not exceed 27 turns.

(3)  
If either side of rudder trailing edge is outside index tab, but less than 0.50 inch from nearest side of tab, proceed as follows:

(a)  
Loosen both checknuts on main input (vernier control) rod.

(b)  
Turn sleeve to obtain small amounts of right or left rudder until both sides of trailing edge fall within width of index tab.

May 01/03 27-21-91 Page 501
BOEING PROPRIETARY - Copyright . - Unpublished Work - See title page for details.

G27258 G27285
Rudder Travel and Index Plate  501 
27-21-91Page 502   Figure 501 BOEING PROPRIETARY - Copyright . - Unpublished Work - See title page for details.  Nov 01/74 

526  Rudder Power Control Unit Adjustment 
Aug 15/69  Figure 502  27-21-91 
Page 503 
BOEING PROPRIETARY - Copyright . - Unpublished Work - See title page for details. 

(c)  Tighten both checknuts, and check that combined total number of visible threads on sleeve and rod end does not exceed 27 turns.
(4)  
If either side of rudder trailing edge is more than 0.50 inch from nearest side of index tab, proceed as follows:

(a)  
Position flight control switches A and B to OFF.

(b)  
Loosen both checknuts on vernier control rod, and remove clevis bolt from forward end of rod assembly.

CAUTION:  RUDDER POWER MUST BE OFF BEFORE DISCONNECTING VERNIER CONTROL ROD FROM ACTUATOR INPUT CRANK.

(c)  
Adjust by turning sleeve and rod end together relative to tube and rod end relative to sleeve an equal number of turns in the same direction, keeping number of threads exposed approximately equal in both parts. One-half turn is equivalent to approximately one-half inch of rudder movement at trailing edge. Lengthen assembly for right rudder, and shorten for left rudder.

(d)  
Position rod end to actuator input crank and install clevis bolt.

(e)  
Position flight control switches A and B to ON.

(f)  
Turn vernier control rod sleeve until both sides of rudder trailing edge fall within width of rudder index tab.

(g)  
Tighten both checknuts on vernier control rod. Combined total number of visible threads on sleeve and rod end shall not exceed 27 turns.

(5)  
Remove rigging pin R-5.

(6)  
Test rudder power control unit per paragraph 2.

(7)  
Replace access panels.

500 
27-21-91 Page 504  BOEING PROPRIETARY - Copyright . - Unpublished Work - See title page for details.  Feb 15/73 

2.  Rudder Power Control Unit Test
A.  Prepare for Test
(1)  
Disconnect pushrod to nose wheel steering (Fig. 503).

NOTE:  Nose gear steering need not be disconnected if airplane is jacked so that nose gear is fully extended.

(2)  
Provide rudder systems A and B hydraulic power (AMM 27-21-0).

(3)  
Check that there is no leakage at power control unit hydraulic connectors.

(4)  
Position rudder trim knob to neutral.

(5)  
Check that rudder is at neutral.

B.  Test yaw damper operation. (Airplanes not incorporating SB 27A1206)
(1)  
Check that yaw damper circuit breakers on P6 are closed.

(2)  
Place autopilot system select switch to B position and yaw damper engage switch to ON.

(3)  
Move yaw damper test switch, located on center instrument panel, to the right. Observe yaw damper indicator needle moves to the right. Observe rudder moves approximately 2 to 4 degrees to the right.

CAUTION: DO NOT ACTUATE TEST SWITCH FOR MORE THAN 10 SECONDS.

(4)  
Move yaw damper test switch to the left. Observe yaw damper indicator needle moves to the left. Observe rudder moves approximately 2 to 4 degrees to the left.

CAUTION: DO NOT ACTUATE TEST SWITCH FOR MORE THAN 10 SECONDS.

(5)  
Disengage yaw damper.

B1.  Perform Yaw Damper Coupler Installation Test (AMM 22-12-01/401) (Airplanes incorporating  SB 27A1206).
　
中國航空網 www.k6050.com
航空翻譯 www.aviation.cn
本文鏈接地址：737 AMM 飛機維護手冊 飛行操縱 FLIGHT CONTROLS 2(37)