STEMME S12 Maintenance Manual

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MAINTENANCE MANUAL STEMME S12

This aircraft may only be maintained in accordance with the instructions given in this manual.

CONTACT

STEMME AG

Flugplatzstrasse F2, Nr. 6-7

D-15344 Strausberg

Phone: +49.3341.3612-0

E-mail: [email protected]

Web: www.stemme.com

Reproduction or disclosure to third parties of this document any part there of is not permitted, except with the prior and express written permission of STEMME AG.

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0.2 LIST OF EFFECTIVE PAGES

This list is only valid for the serial no. specied on title page. The list contains all amendments

of the maintenance manual, eective until nal approval of this serial no. amendments added later must be recorded.

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ANNEX A

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0.3 CONTENTS

MANUAL CONTENTS

Cover Sheet. ..............................................................................................................0-1

0.1 Record of Amendments. ............................................................................... 0-3

0.2 List of Eective Pages . ................................................................................. 0-5

0.3 Contents. ...................................................................................................0-11

1. General. ...................................................................................................... 1-1

1.1 remarks on maintenance. .............................................................................. 1-1

1.2 Conversion Table. .......................................................................................... 1-3

1.3 Abbreviations. ..............................................................................................1-4

1.4 Safety symbols. ............................................................................................1-6

2. Description. ................................................................................................. 2-1

2.1 Technical Data. ............................................................................................2-2

2.2 Three View Drawing. .....................................................................................2-7

3. Description of Assemblies. .............................................................................. 3-1

3.1 Airframe, Primary and Secondary Structure. ..................................................... 3-1

3.1.1 Wing. .......................................................................................................... 3-1

3.1.2 Fuselage. .....................................................................................................3-2

3.1.3 Tail Unit. ......................................................................................................3-2

3.2 Cockpit. ......................................................................................................3-2

3.2.1 General. ......................................................................................................3-2

3.2.2 Control Elements and Instruments. .................................................................3-3

3.2.3 Controls at the Instrument Panel. ....................................................................3-5

3.2.4 Instrument Panel. .........................................................................................3-5

3.3 Flight Control System. ....................................................................................3-9

3.3.1 Longitudinal Control......................................................................................3-9

3.3.2 Longitudinal Trim (-). ....................................................................................3-9

3.3.3 Wing Flap Control System. ..............................................................................3-9

3.3.4 Lateral Control. ............................................................................................3-9

3.3.5 Control of Air Brakes. ................................................................................... 3-10

3.3.6 Directional Control. ..................................................................................... 3-10

3.4 Power-Plant. ...............................................................................................3-11

3.4.1 Engine. .......................................................................................................3-11

3.4.2 Lubrication System. ......................................................................................3-12

3.4.3 Cooling System. .......................................................................................... 3-13

3.4.4 Air Induction System. ................................................................................... 3-14

3.4.5 Engine Exhaust incl. Turbocharger and Attachment. ......................................... 3-14

3.4.6 Fuel System. ............................................................................................... 3-15

3.4.6.1 Fuel System Related Controls and Indication System. ....................................... 3-15

3.4.6.2 Fuel System Design. ..................................................................................... 3-16

3.4.6.3 Fuel System of the Fuselage. ......................................................................... 3-16

3.4.6.4 Electrical Layout and Operation of the Fuel Pumps. ..........................................3-17

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3.4.7 Engine Controls and Instrumentation. ............................................................ 3-19

3.4.8 Fire Protection. ........................................................................................... 3-19

3.4.9 Engine Cowlings. ........................................................................................ 3-19

3.4.10 Propeller. ................................................................................................... 3-20

3.4.10.1 Design of Variable Pitch Propeller. ..................................................................3-21

3.4.10.2 Propeller Blades. ..........................................................................................3-21

3.4.10.3 Propeller Blade Angle Control Mechanism. .......................................................3-21

3.4.10.4 Propeller Weight Balance. ............................................................................ 3-24

3.4.10.5 Expansion Element and Control System. ......................................................... 3-25

3.4.10.6 Identication of Propeller System and Assembly. .............................................3-26

3.4.11 Power Transmission. .................................................................................... 3-26

3.4.12 Forward Propeller Reduction Gearbox. ...........................................................3-26

3.4.13 Operation Mechanism of the Propeller Folding System. ..................................... 3-27

3.4.13.1 Nose-cone. ................................................................................................ 3-27

3.4.13.2 Propeller Positioning. ................................................................................... 3-27

3.4.13.3 Propeller Brake. ..........................................................................................3-28

3.4.14 FLUIDS. .....................................................................................................3-28

3.4.14.1 Fuel. .........................................................................................................3-28

3.4.14.2 Coolant Fluid. ............................................................................................. 3-30

3.4.14.3 Lubricants. ................................................................................................ 3-31

3.5 Landing Gear.............................................................................................. 3-33

3.5.1 Main Landing Gear. ..................................................................................... 3-33

3.5.2 Tail Wheel. .................................................................................................3-34

3.6 Flight Control Instruments, Pitot Static System. ............................................... 3-35

3.7 Electrical System. ........................................................................................ 3-37

3.7.1 General System Architecture. ........................................................................ 3-37

3.7.2 DC Generation (ATA 24-30-00). ..................................................................... 3-39

3.7.2.1 Main System DC Power Generation. ...............................................................3-39

3.7.2.2 Engine System DC Power Generation. ............................................................. 3-40

3.7.2.3 Over Voltage Protection. .............................................................................. 3-41

3.7.3 DC Indication (ATA 24-37-00)........................................................................ 3-41

3.7.3.1 Description. ............................................................................................... 3-41

3.7.4 Parts Specication. ..................................................................................... 3-41

3.7.5 Operation. ................................................................................................. 3-41

3.7.6 DC Power Distribution System (ATA 24-60-00). ................................................ 3-42

3.7.6.1 Wire Specications. ..................................................................................... 3-42

3.7.6.2 Design of Electrical Connections. .................................................................. 3-42

3.7.6.3 Protection Devices. ..................................................................................... 3-43

3.7.6.4 Fuse and CB Specications: Main System. ....................................................... 3-44

3.7.6.5 Fuse Specications: Engine System. ...............................................................3-45

3.7.7 Switching Devices. ...................................................................................... 3-45

3.7.7.1 Manual Switches. ........................................................................................ 3-46

3.7.7.2 Electromechanical Relays. ............................................................................ 3-46

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3.7.7.3 Semiconductor Switch Box. .......................................................................... 3-47

3.7.8 Main System Distribution Box........................................................................ 3-47

3.7.9 Engine System Distribution Box . ................................................................... 3-47

3.7.10 TCU Board. ................................................................................................ 3-47

3.7.11 Instrument Panel Bus System. ....................................................................... 3-49

3.7.12 Engine Starting System (ATA 80-00-00). ......................................................... 3-50

3.7.12.1 Description. ...............................................................................................3-50

3.7.12.2 Parts Specication. ..................................................................................... 3-50

3.7.12.3 Operation. .................................................................................................3-50

3.7.13 Engine Ignition System (ATA 74-00-00). .......................................................... 3-51

3.7.13.1 Description. ............................................................................................... 3-51

3.7.13.2 Parts Specication. ..................................................................................... 3-51

3.7.13.3 Operation. ................................................................................................. 3-51

3.7.14 Engine RPM Indicating System (ATA 77-14-00). ................................................ 3-52

3.7.14.1 Description. ............................................................................................... 3-52

3.7.14.2 Parts Specication. ..................................................................................... 3-52

3.7.14.3 Operation. ................................................................................................. 3-52

3.7.15 Cylinder Head Temperature Indicating System (ATA 77-22-00). .......................... 3-52

3.7.15.1 Description. ............................................................................................... 3-52

3.7.15.2 Parts Specication. ..................................................................................... 3-52

3.7.15.3 Operation. ................................................................................................. 3-52

3.7.16 Oil Pressure Indicating System (ATA 79-31-00). ................................................ 3-53

3.7.16.1 Description. ............................................................................................... 3-53

3.7.16.2 Parts Specication. ..................................................................................... 3-53

3.7.16.3 Operation. ................................................................................................. 3-53

3.7.17 Oil Temperature Indicating System (ATA 79-33-00). .......................................... 3-53

3.7.17.1 Description. ...............................................................................................3-53

3.7.17.2 Parts Specication. ..................................................................................... 3-53

3.7.17.3 Operation. ................................................................................................. 3-54

3.7.18 Fuel Distribution System (ATA 28-20-00). ........................................................ 3-54

3.7.18.1 Description. ............................................................................................... 3-54

3.7.18.2 Parts Specication. .....................................................................................3-54

3.7.18.3 Operation. .................................................................................................3-54

3.7.18.4 Transfer Pump Control and Tank Selection. ..................................................... 3-55

3.7.19 Fuel Quantity Indication (ATA 28-40-00).........................................................3-55

3.7.19.1 Description. ...............................................................................................3-55

3.7.19.2 Parts Specication. ..................................................................................... 3-55

3.7.19.3 Operation. .................................................................................................3-55

3.7.20 Fuel Pressure Indicating (ATA 73-32-00). ......................................................... 3-56

3.7.20.1 Description. ...............................................................................................3-56

3.7.20.2 Parts Specication. .....................................................................................3-56

3.7.20.3 Operation. .................................................................................................3-56

3.7.21 Fire Detection System (ATA 26-12-00). ............................................................ 3-56

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3.7.21.1 Description. ...............................................................................................3-56

3.7.21.2 Parts Specication. .....................................................................................3-56

3.8 COM and NAV Equipment. ............................................................................ 3-57

3.9 Oxygen System. .......................................................................................... 3-59

3.10 Wing Folding Device. ................................................................................... 3-61

4. Airworthiness Limitations Section. ..................................................................4-1

4.1 General. ......................................................................................................4-2

4.2 Maintenance Limitations. ..............................................................................4-2

4.2.1 Paint Finish. .................................................................................................4-2

4.3 Component Replacement and Overhaul Limitations. ..........................................4-3

4.4 Structural Limitations. .................................................................................. 4-5

5. Time Limits / Maintenance Checks. ..................................................................5-1

5.1 Overhaul and Replacement Schedule. ..............................................................5-2

5.2 Pre-Flight Inspections. ...................................................................................5-7

5.2.1 Rubber Hose and Clamp Integrity. ...................................................................5-7

5.3 Periodic Inspections. ...................................................................................... 5-7

5.3.1 Inspection Intervals. ..................................................................................... 5-8

5.3.2 General Remarks on Periodic Inspections. ....................................................... 5-10

5.3.3 Unscheduled Maintenance. ...........................................................................5-11

5.3.4 Special Conditions and Cautionary Notice. ..................................................... 5-12

5.3.5 General Remarks on Maintenance. ................................................................ 5-12

5.3.6 Inspection Groups and Maintenance Criteria. .................................................. 5-14

5.3.7 Other Particulars of Maintenance. ................................................................. 5-17

5.4 Check List for Periodic Inspections. ................................................................ 5-19

5.4.1 General. .................................................................................................... 5-19

5.4.2 Wings and Fuel System Components in the Wing Center Section. ....................... 5-20

5.4.3 Forward Fuselage. ....................................................................................... 5-21

5.4.4 Cockpit. .................................................................................................... 5-22

5.4.5 Center Fuselage (except for fairings). ............................................................ 5-23

5.4.6 Tail Boom. ..................................................................................................5-24

5.4.7 Empennage. ............................................................................................... 5-24

5.4.8 Fuel System Components in the Fuselage. .......................................................5-25

5.4.9 Engine and Engine Mountings. ...................................................................... 5-26

5.4.10 Lubrication System. ..................................................................................... 5-28

5.4.11 Cooling System (Liquid Cooling, Ram Air Cooling). ........................................... 5-29

5.4.12 Air Induction System. ................................................................................... 5-29

5.4.13 Engine Controls & Monitoring. ....................................................................... 5-29

5.4.14 Center Fuselage Fairing, Engine Cowlings and Firewall. ..................................... 5-30

5.4.15 Propeller. ................................................................................................... 5-31

5.4.16 Drive Shaft with Propeller Reduction Gear. ...................................................... 5-32

5.4.17 Main Landing Gear. ..................................................................................... 5-33

5.4.18 Tail Wheel. ................................................................................................. 5-34

5.4.19 Flight Instrumentation and Pitot / Static Systems. ............................................5-34

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5.4.20 Electric System (except for engine and TCU). .................................................. 5-35

5.4.21 COM and NAV Equipment. ............................................................................5-35

5.4.22 Completion Tasks. ....................................................................................... 5-36

5.4.23 Optional Equipment. ................................................................................... 5-37

5.5 Special Inspections. ..................................................................................... 5-39

5.5.1 Inspection Following a Heavy Landing or a Wing Tip Landing. ............................5-39

5.5.2 Inspection Following an Impact to the Rotating Propeller. ................................. 5-39

5.5.3 Inspection following Exceptional Events with Wing Folding Device. .....................5-40

6. Maintenance Instructions, Tolerances, Adjustment Data for the Aircraft................6-1

6.1 General Remarks. .........................................................................................6-1

6.2 Towing on ground, Jack Points and Lifting. .......................................................6-1

6.3 Determination of Empty Weight and Corresponding

Center of Gravity; Weight Limits. .....................................................................6-3

6.4 Flight Control System. ................................................................................... 6-8

6.4.1 Deection of Control Surfaces, Control System Friction, and Control Forces. ......... 6-8

6.4.1.1 Control Surface Deections. .......................................................................... 6-8

6.4.1.2 Friction in Control System. ............................................................................. 6-9

6.4.1.3 Pilot Forces. ................................................................................................ 6-9

6.4.2 Weights and Moments of Control Surfaces...................................................... 6-10

6.4.3 Free Play in Flight Control System. ................................................................. 6-10

6.5 Lubrication. ................................................................................................6-11

6.5.1 General Remarks. ........................................................................................6-11

6.5.2 Lubrication Plan. ........................................................................................ 6-12

6.6 Surface of Composite Structures. .................................................................. 6-13

6.7 Drainage and Ventilation Holes. .................................................................... 6-15

6.8 Tightening Torques of Screwed Joints. ............................................................ 6-15

7. Maintenance Instructions, Tolerances and Adjustment

Data for Assemblies / Equipment. .................................................................... 7-1

7.1 Airframe. ..................................................................................................... 7-1

7.1.1 Wing. .......................................................................................................... 7-1

7.1.1.1 Cracks and Structural Damage. ...................................................................... 7-1

7.1.1.2 Drainage and Ventilation Holes. ...................................................................... 7-1

7.1.1.3 Wing Fittings. ............................................................................................... 7-1

7.1.1.4 Flaps and Ailerons. .......................................................................................7-2

7.1.2 Fuselage. .....................................................................................................7-3

7.1.2.1 Cracks and Structural Damage. ......................................................................7-3

7.1.2.2 Drainage Holes. ............................................................................................7-3

7.1.2.3 Nose-cone. ..................................................................................................7-3

7.1.2.4 Center Fuselage. ...........................................................................................7-4

7.1.3 Empennage. .................................................................................................7-4

7.1.3.1 General Condition. ........................................................................................7-4

7.1.3.2 Vertical Tail. .................................................................................................7-5

7.1.3.3 Horizontal Tail. .............................................................................................7-6

7.1.3.4 Weights and Static Hinge Moments of Control Surfaces. ..................................... 7-6

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7.2 Cockpit. ......................................................................................................7-7

7.2.1 Canopy. ....................................................................................................... 7-7

7.2.1.1 Condition and Function. ................................................................................7-7

7.2.1.2 Test of Canopy Emergency Jettison. .................................................................7-7

7.2.1.3 Re-Installation of the Canopy. ........................................................................7-8

7.2.2 Equipment and Systems. ................................................................................7-9

7.2.2.1 Seat Belts. ...................................................................................................7-9

7.2.2.2 Instruments and Switches. .............................................................................7-9

7.2.2.3 Air Hoses. ....................................................................................................7-9

7.2.2.4 Seats. .........................................................................................................7-9

7.3 Controls. ................................................................................................... 7-10

7.3.1 Controls in Fuselage. ................................................................................... 7-10

7.3.2 Controls in the Wing. .................................................................................... 7-11

7.3.3 Controls in Tail Cone /Vertical Tail. .................................................................. 7-11

7.3.4 Deection of Control Surfaces, Control System Friction, Control Forces. ...............7-12

7.3.5 Free Play of Control System Bearings...............................................................7-12

7.4 Power-Plant. ...............................................................................................7-13

7.4.1 Engine. .......................................................................................................7-13

7.4.1.1 General. .....................................................................................................7-13

7.4.1.2 Removal and Installation of The Engine. ......................................................... 7-14

7.4.1.3 Cleaning of the Engine. ................................................................................ 7-15

7.4.1.4 Visual Check of Engine. ................................................................................ 7-15

7.4.1.5 Leakage Check. .......................................................................................... 7-15

7.4.1.6 Inspection of External Engine Components. .................................................... 7-15

7.4.1.7 Check of Waste-Gate. ................................................................................. 7-15

7.4.1.8 Check of Gearbox. ...................................................................................... 7-16

7.4.1.9 Inspection of the Magnetic Plug. ................................................................... 7-16

7.4.1.10 Examination of Drive Gear Set....................................................................... 7-16

7.4.1.11 Check of the Carburetor. .............................................................................. 7-16

7.4.1.12 Check of Engine Wiring. ................................................................................7-17

7.4.1.13 Check of V-Belt Tension. ...............................................................................7-17

7.4.1.14 Replacement of Spark Plugs. .........................................................................7-17

7.4.1.15 Checking of Spark Plug Connector. .................................................................7-17

7.4.1.16 Check of Compression. .................................................................................7-17

7.4.1.17 Test Run of Engine. ...................................................................................... 7-18

7.4.1.18 Inspection of Engine Mountings. ................................................................... 7-18

7.4.2 Lubrication System. ..................................................................................... 7-18

7.4.2.1 Visual and Leakage Check of Oil System. ........................................................ 7-18

7.4.2.2 Oil Level Check. .......................................................................................... 7-19

7.4.2.3 Oil Change. ................................................................................................ 7-19

7.4.2.4 Oil lter - Replacement and Inspection. ......................................................... 7-19

7.4.3 Cooling System. .......................................................................................... 7-19

7.4.3.1 Visual and Leakage Check of the System. ....................................................... 7-19

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7.4.3.2 Flushing of Cooling System. .......................................................................... 7-20

7.4.3.3 Coolant Replacement. ................................................................................. 7-20

7.4.3.4 Ram Air Cooling. .......................................................................................... 7-21

7.4.4 Air Induction System. ....................................................................................7-21

7.4.4.1 Intercooler and Air box. .................................................................................7-21

7.4.4.2 Air lter. .....................................................................................................7-21

7.4.5 Exhaust System. ..........................................................................................7-21

7.4.5.1 Turbocharger / Muer Mountings. ..................................................................7-21

7.4.5.2 Exhaust Headers. ........................................................................................ 7-22

7.4.5.3 Heat Insulation. .......................................................................................... 7-22

7.4.6 Fuel System. ............................................................................................... 7-23

7.4.6.1 Fuel System in the Wing. .............................................................................. 7-24

7.4.6.2 Fuel System in the Fuselage. ......................................................................... 7-25

7.4.7 Engine Controls / Monitoring. ........................................................................7-26

7.4.8 Fire Protection. ........................................................................................... 7-28

7.4.8.1 Cowling-Fire Protective Painting. .................................................................. 7-28

7.4.8.2 Fire Protective Sleeves. .................................................................................7-29

7.4.8.3 Firewall. .................................................................................................... 7-29

7.4.9 Cowlings.................................................................................................... 7-29

7.4.10 Propeller. ................................................................................................... 7-30

7.4.10.1 General. .................................................................................................... 7-30

7.4.10.2 Adjustment and Inspection Results, Tolerances. ................................................7-31

7.4.10.3 Visual Inspection of the Propeller Unit. ............................................................7-31

7.4.10.4 Removal of the Propeller Unit. ...................................................................... 7-32

7.4.10.5 Installation of the Propeller Unit. .................................................................. 7-33

7.4.10.6 Spring Tension of the Propeller Folding Mechanism. .......................................... 7-34

7.4.10.7 Checking of Pitch Control and Adjusting the Variable Pitch Propeller. ................. 7-35

7.4.10.8 General Aspects of the Electrical Circuit. ........................................................7-36

7.4.10.9 Adjustment of the Electrical Switch for the

Take-O Setting Indicator and Stop Switch. .................................................... 7-36

7.4.10.10 Dynamic Balancing. .................................................................................... 7-37

7.4.11 Drive Shaft System. ..................................................................................... 7-37

7.4.12 Propeller Gearbox, Mounting and Support. ...................................................... 7-38

7.4.13 Propeller Operation. ....................................................................................7-40

7.5 Landing Gear.............................................................................................. 7-41

7.5.1 Main Landing Gear. ..................................................................................... 7-41

7.5.1.1 General. .................................................................................................... 7-41

7.5.1.2 Functional Test of the Retractable Landing Gear. ............................................. 7-42

7.5.1.3 Functional Check of Emergency Undercarriage Extension. ................................ 7-44

7.5.1.4 Tires, Condition and Ination Pressure. ..........................................................7-45

7.5.1.5 Adjustment and Bleeding of the Wheel Brake System. ...................................... 7-45

7.5.1.6 Replacement of Brake Lining. ....................................................................... 7-46

7.5.1.7 Removal and Installation of Landing Gear Legs. ..............................................7-46

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7.5.2 Tail Wheel. ................................................................................................. 7-46

7.6 Flight Control Instruments and Pitot and Static Pressure System. ....................... 7-47

7.6.1 Calibration of Stall Warning System (Optional). ...............................................7-48

7.6.2 Maintenance on the Static Pressure System. ................................................... 7-48

7.7 Electrical System. ........................................................................................ 7-49

7.7.1 General. .................................................................................................... 7-49

7.7.2 Batteries. ..................................................................................................7-49

7.7.3 Grounding. ................................................................................................ 7-50

7.7.4 E-Boxes. .................................................................................................... 7-50

7.8 Communication and Navigation Equipment. ...................................................7-51

7.9 Optional - Oxygen Equipment. ......................................................................7-51

8. List of Placards and their Positions. .................................................................8-1

8.1 Placards in the Cockpit, Front. ........................................................................ 8-1

8.2 Placards in the Cockpit, Rear. ........................................................................ 8-6

8.3 Placards on Fuselage. ....................................................................................8-7

9. Equipment. ..................................................................................................9-1

9.1 Minimum Equipment List. ..............................................................................9-1

9.2 Supplementary Equipment. ............................................................................9-3

9.3 Additional Equipment and Systems. .................................................................9-3

9.3.1 Additional Equipment. ...................................................................................9-3

9.3.2 Optional Systems. ........................................................................................ 9-5

10. List of Special Tools. .....................................................................................10-1

11. List of Maintenance Documents for Parts Being Approved

Independently from the Aircraft. .................................................................... 11-1

Annex A: Supplementary Instructions for Maintenance and Care,

Maintenance Instructions. .............................................................................. A-1

Annex B: Service Bulletins, Airworthiness Directives.........................................................B-1

Annex C: Documents (Inspection and Operation Reports). ...............................................C-1

Annex D: Maintenance and Inspection Forms. ................................................................D-1

Annex E: Maintenance Manual (Line Maintenance) for ROTAX Engine Type 914 Series. .......... E-1

Annex F: Technical Drawings. ...................................................................................... F-1

a) Flight Controls: Pitch Control and Trim. ......................................................................F-3

b) Flight Controls: Lateral Control (Wing, Flap, Aileron). ................................................... F-5

c) Flight Controls: Airbrake Control................................................................................F-7

d) Flight Controls: Rudder Control. ................................................................................F-9

e) Propulsion System. .................................................................................................F-13

f) S12 Main System - DC Generation. ............................................................................F-15

g) S12 Main System - Landing Gear. .............................................................................. F-17

h) S12 Main System - Electrical Trim. ............................................................................ F-19

i) S12 Main System - Autopilot. ..................................................................................... F-21

j) S12 Main System - Fire Warning. ............................................................................... F-23

k) S12 Main System - Lighting. .................................................................................... F-25

l) Transponder and Trac Display. ............................................................................... F-27

m) S12 Main System - Communication Radio. ............................................................... F-29

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n) S12 Engine System - DC Generation. .........................................................................F-31

o) S12 Engine System - Engine Monitoring and Control. .................................................. F-33

p) S12 Engine System - Fuel Pumps and Indication. ........................................................ F-35

q) S12 Engine System - Propeller Control. ...................................................................... F-37

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MAINTENANCE MANUAL STEMME S12

1. GENERAL

1.1 REMARKS ON MAINTENANCE

The legal owner or operator of the STEMME S12 is obliged to ensure that the maintenance of the aircraft follows the instructions of this manual in accordance with applicable national laws and regulations. Among others, these include:

• Scheduled maintenance,

• Adjustments,

• Replacement of uids and lubricants,

• Replacement of parts after expiration of their service life,

• Minor repairs.

Any maintenance work must be documented (aircraft logbook).

The manufacturer must be informed immediately in case of any change of ownership. The message must be conrmed by the manufacturer, so that all information concerning airworthiness (AD´s, SB´s) can be given to the legal owner.

For maintenance work the following documents are relevant:

1. This Maintenance Manual for the motorglider STEMME S12.

2. STEMME - Flight Manual for the motorglider STEMME S12.

3. ROTAX - Maintenance Manual (Line Maintenance) for ROTAX Engine Types ROTAX 912 and 914 Series.

4. ROTAX - Maintenance Manual (Heavy Maintenance) for ROTAX Engine Types ROTAX 912 and 914 Series.

5. ROTAX - Operating Manual ROTAX 914 F.

6. ROTAX - Installation Manual ROTAX 914 F.

7. ROTAX - Main Overhaul Manual ROTAX 914 F.

8. ROTAX - Spare Parts List ROTAX 914 F.

9. STEMME Document Number A26-11AM-M: Technical Specication of the ROTAX 914 F2/S1.

10. Maintenance instructions for the L’Hotellier quick-disconnects in ight control system.

11. Manufacturer’s documents referring to the equipment listed in the equipment list of the corresponding S/N.

12. SB´s published by STEMME, ROTAX and manufacturer of other equipment installed.

13. Maintenance Instructions from STEMME.

14. Service Information from ROTAX.

15. Aircraft Logbook.

The amount and kind of maintenance work depend on the aircraft utilization, the climate, aireld conditions, storage facilities and other factors, irrespective of the periodic checks. E. g., in sandy environs it might be necessary to clean all lters before every commencement of operation; on

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the other hand in coastal or in rainy regions it is important to take more care of the preservation of the aircraft.

The instructions in this manual are valid under normal conditions and use. Use only spare parts from the manufacturer or according to the manufacturer’s requirements.

NOTICE

Materials required and recommended procedures for minor repairs on composite materials are indicated in the repair guide Small Repairs of Composite Materials in Annex A of the S12 Aircraft Maintenance Manual.

In case of any incident endangering airworthiness, the manufacturer must be informed immediately. Maintenance work must be performed by qualied personnel.

NOTICE

This Maintenance Manual does not include instructions for assembly, daily inspection and Pre-ight inspection, which are provided in section 4 Normal Operating Procedures of the S12 Aircraft Flight Manual. To perform these procedures, the S12 Aircraft Flight Manual must be available to the maintenance personnel.

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1.2 CONVERSION TABLE

For the conversion of technical data the following factors have been used:

Data Factor 1 cubic in 16.387 cm 1 dr 1.772 g 100 fpm 0.508 m/s 1 ft 0.3048 m 1 ft lb 1.356 Nm 1 hp 0.7457 kW 1 Imp.gal. 4.546 l 1 in 25.4 mm 1 inHg 33.86 hPa 1 kgm 86.8 lb in 1 kgmm 0.001 kgm 1 kgmm 0.007 lb ft 1 kt 1.852 kph 1 lb 0.4536 kg 1 lbf 4.45 N 1 lbf ft 135.6 Ncm 1 mph 1.609 kph 1 sqft 0.0929 m 1 oz. 28.349 g 1 psi 0.06895 bar 1 US gal. 3.785 l

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1.3 ABBREVIATIONS

The following abbreviations are being used for clarity:

ACL Anti-Collision Light AD’s Airworthiness Directives AGL Above Ground Level AMM Aircraft Maintenance Manual ANSI American National Standards Institute AOA Angle Of Attack ASAP As Soon As Possible ASI Airspeed Indicator AUW All-Up-Weight CB Circuit Breaker CCT Cylinder Coolant Temperature CD Constant Depression CFRP Carbon Fiber Reinforced Plastic CG Center Of Gravity CHT Cylinder Head Temperature DCDI Dual Capacitive Discharge Ignition EFIS Electronic Flight Information System EMC Electromagnetic Compatibility EQL Equipment List FBO Fixed-Base Operator FOD Foreign Object Damage FRP Fiber Reinforced Plastic GEN Generator GFRP Glass Fiber Reinforced Plastic GPS Global Positioning System IMC Instrument Meteorological Conditions IPC Illustrated Parts Catalog KIAS Knots Indicated Airspeed kt Knots (equivalent to KIAS) LBA Luftfahrtbundesamt -

German Civil Aviation Authority LH Left Hand MAP Manifold Pressure MCP Maximum Continuous Power MEL Minimum Equipment list MMEL Master Minimum Equipment list MNLP Maximum Weight of Non Lifting Parts MTOW Maximum Take-O Weight NLP Non Lifting Part

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OAT Outside Air Temperature OVP Over Voltage Protection PIC Pilot In Command PPC Propeller Pitch Control PTT Push To Talk R/C Rate Of Climb RH Right Hand RP Reference Plane RPM Revolutions Per Minute RTD Resistance Temperature Detector RWY Runway SB’s Service Bulletins SI Service Information SL Service Letter STC Supplemental Type Certicate TBO Time Between Overhaul TBR Time Between Replacement TC Type Certicate TCDS Type Certicate Data Sheet TCU Turbocharger Control Unit TE Total Energy T/O Take-O TSO Technical Standard Order TTL Transistor-Transistor Logic

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1.4 SAFETY SYMBOLS

The following safety instructions are based loosely on the ANSI Z535.

Symbols Z535.3 Criteria for Safety Symbols Signs and Labels Z535.4 Product Safety Signs and Labels Product Safety

WARNING

CAUTION

NOTICE

Z535.6 Product Safety Information in Product Manuals, Instructions, and Other Collateral Materials.

WARNING An operating procedure or technique that may result in personal injury or loss of life if not followed.

CAUTION An operating procedure or technique that may result in damage to equipment if not followed.

NOTICE An operating procedure or technique needing special emphasis.

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MAINTENANCE MANUAL STEMME S12

2. DESCRIPTION

The STEMME S12 is a twin-seat, self-launching motorglider, constructed mostly from carbon ber and is aerodynamically optimized for high performance. The two seats are arranged side-by-side (forward of the wing) and the S12 is equipped with dual controls.

The wing is shoulder mounted and consists of a center section with aps and Schempp-Hirth air brakes, two outer wing sections with continuous ailerons, two wing extensions with ailerons and two winglets. The aps and ailerons of inner and outer wing/wing extension are interconnected (“aperons”).

The motorglider is in “T”-tail design.

The retractable two-wheel main landing gear is electrically operated and is equipped with hydraulic wheel brakes.

The engine of the STEMME S12 is based on the ROTAX 914 F2-01, which has been certicated by BRP-ROTAX® GmbH & Co.KG, Austria. STEMME modied the arrangement of some accessories (induction and exhaust system including turbo-charger, engine mounts etc.) to adapt the systems to specic installation requirements of the S12. These modications are certied in the STEMME S12 as engine model ROTAX 914 F2-01/S1.

ENGINE DESCRIPTION

Four-cylinder, four-stroke opposed type Otto-engine, turbocharged with electronic manifold pressure control (TCU); a center cam shaft and tappets; OHV; liquid cooled cylinder heads, cylinder barrels cooled by ram air; dry-sump lubrication; Dual Capacitive Discharge Ignition (DCDI); 2 CD-carburetors (variable-choke); integrated reduction gear with mechanical vibration absorber and overload clutch.

The engine is mounted in the fuselage in a center steel frame near the aircraft´s CG. Engine power is transmitted via a propeller shaft made of composites and a spur gear to the variable pitch propeller in the fuselage nose.

When the aircraft is in gliding conguration, the propeller blades are folded and covered by a movable nose cone (nose-cone). Propeller pitch change from take-o into cruise position is accomplished by electrically heated expansion elements, and from cruise back to take-o position (elements unheated) by springs as well as aerodynamic and centrifugal forces.

A fuel tank is located at each outboard end of the wing center section. Each tank supplies the feeder tank in the fuselage by use of electrically driven fuel pumps. From the feeder tank the fuel is supplied by two fuel pumps (one main and one auxiliary) to the engine to provide required redundancy.

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MAINTENANCE MANUAL STEMME S12

2.1 TECHNICAL DATA

FUSELAGE Design

Length 8.42 m 27.62 ft Width 1.18 m 3.87 ft Cockpit width 1.16 m 3.81 ft Cockpit height 0.93 m 3.05 ft Height of tail unit 1.75 m 5.74 ft

WING Design 5 sections CFRP-Spar with GFRP-shearweb,

Wing span 25.07 m 82.25 ft Center wing span 9.90 m 32.48 ft Wing area 19.95 m² 214.75 sqft Aspect ratio 31.5 Dihedral angle (V-form) 0.75° / 1.75° / 4° Mean aerodynamic chord 0.741 m 2.43 ft Wing Airfoil HQ41/14.35 Sweep of wing center section leading edge 0° Sweep of outboard wing leading edge up to the bend 0°

Front section CFRP-Kevlar-GFRP-structure, center steel frame, GFRP-fairing, tail boom with integrated vertical tail n (CFRP).

CFRP-sandwich skin.

AIR BRAKES Type Two blade Schempp-Hirth type, position outer

part of wing center section. Length 1.50 m 4.92 ft / 59.01 in Area 0.22 m² 2.37 sqft Maximum height above wing upper surface. 0.16 m 0.52 ft / 6.29 in

HORIZONTAL TAIL Design

Web GFRP, skin CFRP-Sandwich, elevator

GFRP-Sandwich. Span 3.10 m 10.17 ft Area 1.46 m² 15.72 sqft Aspect ratio 6.58 Airfoil FX 71-L-150/25

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VERTICAL TAIL Design Web GFRP, skin CFRP, rudder GFRP-sandwich. Span 1.60 m 5.25 ft Area 1.78 m² 19.16 sqft Airfoil FX71-L150/30 x*1.3

WEIGHTS Max. T/O-weight 900 kg 1984 lb Empty weight 690 kg

1556 lb

(approx., min. equipment.)

Maximum weight of

610 kg 1345 lb

non-lifting parts Total useful load

210 kg 463 lb

(Occupants, fuel, baggage)

WING FLAPS CENTER WING Span 4.39 m 14.40 ft Area 0.75 m² 8.07 sqft Flap positions: -10°, -5°, 0°, +5°, +10°, L (+16°)

AILERONS OUTER WINGS AND EXTENSIONS Span 5.70 m + 1.50 m 18.70 ft + 4.92 ft Area 0.78 m² + 0.13 m

8.40 sqft. + 1.40 sqft

LANDING GEAR 2 main wheels with brake discs, rim of wheel 127x127-30 Tire size 6.00-5 Wheel track 1.36 m 4.46 ft Tail wheel (steerable), tire size 210 x 65 Wheel base 5.46 m 17.91 ft

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ENGINE

MAINTENANCE MANUAL STEMME S12

Design

4-cylinder 4-stroke-Otto-motor, opposed type,

turbocharged, electronic manifold pressure

control integrated reduction gear. Type ROTAX 914 F2-01/S1 Engine reduction gear ratio i = 2.4286 Bore 79.5 mm 3.13 in Stroke 61 mm 2.40 in Displacement 1211 cm

73.9 cubic in Compression ratio. 9.0 Drive shafts turns. Clockwise Max. T/O power. 84.5 kW at 5800 RPM 113.2 hp Max. cont. power. 73.4 kW at 5500 RPM 98.4 hp Fuel ow at T/O power (115%). 33.0 l/h

8.72 US gal. / h

7.26 Imp.gal. / h

Fuel ow at max cont. power (100%). 27.2 l/h

7.19 US gal. / h

5.98 Imp.gal. / h

Fuel ow at 75% power. 20.4 l/h

5.34 US gal. / h

4.49 Imp.gal. / h

Specic fuel consumption at max. cont. power (100%).

276 g/kWh 0.454 lb / hph

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MAINTENANCE MANUAL STEMME S12

PROPELLER Design

2-blade folding propeller, CFRP blades, propeller hub and blade suspension. Forks aluminum. Pitch control via electrically heated thermo

elements. Type STEMME 11 AP-V Propeller diameter 1.63 m 5.35 ft Mass of the propeller

9350 g 20.61 lb (incl. outer-casing of needle bearing and rubber buers)

Overall mass of propeller blade 650 g ± 10 g 1.433 lbs ± 0.022 lb Max. propeller 2.650 RPM Propeller blade angle (station 0.7) 17.65° in T/O 24.05° in cruise. Propeller blade pitch (station 0.7) 114 cm / 44.9 in (T/O position)

160 cm / 63.0 in (cruise position) Max. current consumption of both resistor

10 A elements

Propeller turns Counter-clockwise

INFLIGHT CENTER OF GRAVITY RANGE In-ight center-of-gravity range aft of

265 - 420 mm 10.41 in - 16.5 in datum (See section 6.3 for center wing leading edge)

PROPELLER GEARBOX Design Helical spur gear,

aluminum alloy-housing,

elastic mounting in rubber elements. Type STEMME 11AG Drive shaft turns Clockwise Ratio of propeller gearbox i = 1.109

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FUEL SYSTEM Design

2 FRP-tanks in wing center section + 1 feeder tank, 1 electrical fuel transfer pump for each wing-tank + 2 electrical pumps from feeder tank to the engine.

Maximum fuel volume 126.0 l ( 2%)

33.28 US gal.

27.72 Imp.gal.

Unusable fuel 1.4 l

0.37 US gal.

0.31 Imp.gal.

NOTICE

For more technical data, please refer to the S12 Aircraft Flight Manual.

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2.2 THREE VIEW DRAWING

MAINTENANCE MANUAL STEMME S12

1.36 m / 4.46 ft

9.9 m / 32.48 ft

1.63 m / 5.35 ft

8.4 m / 27.56 ft 8.4 m / 27.56 ft

1.75 m / 5.74 ft

1.9 m / 6.23 ft

9.9 m / 32.48 ft

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25.07 m / 82.25 ft

Figure 2.2.a

Three View Drawing of the S12

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THIS PAGE IS INTENTIONALLY LEFT BLANK

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3. DESCRIPTION OF ASSEMBLIES

3.1 AIRFRAME, PRIMARY AND SECONDARY STRUCTURE

The Primary Structure includes:

• Wing spars, root ribs, and wing spar boxes

• Wing shells

• Center steel frame

• Tail boom and vertical stabilizer

• Front section of fuselage

• Horizontal stabilizer

• Fittings

The Secondary Structure includes:

• Control surfaces

• Cowlings, cooling air system ducts, cockpit components

3.1.1 WING

Structural design: 5 sections; CFRP-Spar with GFRP-shearweb, CFRP-sandwich skin.

The wing consists of four sections: A wing center section with a span of 9.9 m / 32.48 ft, two outer wing sections and two wing extensions including winglets with a length of 1.9 m / 6.23 ft each. Attachment of the wing center section to the fuselage is by means of four sliding bolts, attachment of the outboard wings to the wing center section as well as the connection of wing extensions to the outer wings is with one sliding bolt each.

A removable fairing covers the wing/fuselage junction. Beneath the fairings, free access to the wing attachment, the control system joints and the combined aileron/ap controls is possible. For disassembly, the wing center section can be lifted vertically.

The aps extend over the total span of the wing center section and the ailerons over the total span of the outboard wings. The controls of aps and ailerons are interconnected (“aperons”), the ailerons acting as dierential aps and the aps acting as dierential ailerons. The ap deection is reduced from the inner to the outer wing, the aileron deection is reduced from the outer wing to the wing center section and from up to down deections (dierential aileron deection). The ailerons of the wing extension are directly connected to those of the outer wing and have no separate actuation.

Two-blade Schempp-Hirth air brakes are installed on the wing center section upper surface. The aps and aileron gaps are sealed with elastic adhesive tape and a sliding layer on the upper surface of the wing and with a textile tape (optional: elastic adhesive tape and sliding layer) on the lower surface.

A boundary layer turbulator (adhesive 60° zigzag tape, leading edge at 69% of chord, 12 x 0.5 mm / 0.47 x 0.02 in) on the wing lower surface ensures a dened ow transition.

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3.1.2 FUSELAGE

The fuselage is a modular construction of three assemblies with bolted joints: The forward fuselage section (CFRP) construction, the center fuselage framework (steel frame) and the tail boom (CFRP construction). Loads from the fuselage front section, the wing, the landing gear, the power-plant and the tail unit are transmitted by center steel frame.

The center fuselage framework is covered by self-supporting fairings (upper fairing and engine cowlings LH / RH as well as lower fairing) of GFRP sandwich. The attachment of the fairings is by means of camlocks, which can be easily opened with a screw-driver. The upper fairing includes an oil service access door, which can be opened without a tool.

3.1.3 TAIL UNIT

Horizontal Tail

• T-arrangement, easily removable (simple spring-bolt connection),

• Stabilizer as sandwich construction of CFRP, elevator made of CFRP.

• Elevator gaps sealed by elastic tape.

•

Boundary layer turbulator (adhesive 60° zigzag tape, leading edge at 65% of chord, 12 x 0.5 mm / 0.47 x 0.02 in) on upper and lower side for dened ow transition.

Vertical Tail

• Stabilizer as a sandwich construction of CFRP, rudder as a sandwich construction of GFRP,

• Rudder gap sealed by elastic tape with integrated zigzag turbulator (combi-tape),

• COM antenna integrated in rudder.

3.2 COCKPIT

3.2.1 GENERAL

The two seats are arranged side by side (forward of the wing) and the S12 is equipped with dual controls. The FRP seat backrests are multi-adjustable. Each seat is equipped with 4-point seat belts with a center locking mechanism.

Between the seats a console covers both systems and the drive shaft tunnel. The instrument panel is designed in mushroom-style panel arrangement.

The one-piece canopy is hinged at the front and held in its opened position by two gas springs. Three canopy locks on each side of the canopy frame, left hand and right hand side, are operated by one locking lever on each side. The locking levers are mechanically interconnected by two bowden cables so that they act simultaneously. One additional lock to improve canopy emergency jettisoning (“Röger-hook”) is installed in the rear upper canopy frame, operated by adjacent handle.

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The emergency jettison will be activated by opening either lateral locking lever on canopy side (operating one is sucient) and the red T-shaped handle on the instrument panel. Canopy hinge opens and is lifted approximately 0.1 m / 4 in by a gas spring. The “Röger-hook” must remain closed, since it is the axis of rotation until the canopy is jettisoned.

The cockpit is ventilated via two nozzles (LH and RH side) in the instrument panel. The canopy ventilation is via openings in the instrument panel cover.

3.2.2 CONTROL ELEMENTS AND INSTRUMENTS

COCKPIT CONTROLS AT THE AIRFRAME

Following overview includes the controls at the airframe.

No. Name of Control Airframe Position

1. Control Stick Centered in front of each seat.

2. Rudder Pedals

3. Airbrake Lever At the LH side of each seat. Blue lever at LH cockpit side

4. Flap Lever At the LH side of each seat. Black lever at LH cockpit side

For each seat and adjustable. The pedals also steer the tail wheel, which is coupled to the rudder via spring device.

and on the center console between seats.

and on the center console. Indication of settings (-10, -5, 0, +5, +10, L) in center console. Unlocking is by moving lever to the right against a spring force which locks the ap positions.

5. Pedal Adjustment Handle In front of each seat. Unlocking is by pulling the handle.

6. Canopy Locks Two white handles with red stripe, one on left and one on right side of the canopy frame, to open and lock the canopy. The handles are mechanically connected so that both always move simultaneously. In addition a white handle at rear top center of the canopy is provided to insure correct movement of the canopy during jettison (“Röger-hook”).

7. Brake Lever

Lever on each control stick. Separate lever for parking brake valve on the oor panel console in front of the LH control stick.

8. Trim Knob A green switch on the control stick.

9. Throttle Lever One black lever on center console with two forward stops (For max. continuous and max. T/O-power). It is coupled with a spring acting forward in direction FULL POWER. Its position is xed by friction discs, which can be adjusted with a knurled knob on LH side of the center console.

10. Choke Lever

Black lever on center console, RH side of the throttle lever. It is coupled with a spring acting rearward in direction CHOKE OFF. Its position is xed by friction discs, which can be adjusted with a knurled knob on RH side of the center console.

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No. Name of Control Airframe Position

11. Propeller Pitch

Switch on center console. The forward position is the TAKE- OFF position. A green light located in the upper area of left instrument panel indicates if propeller pitch (not switch) is in T/O)-position.

12. Fuel Valve

Red handle on the rear console between the seat back rests. Turning the handle into its horizontal position (fuel valve CLOSED) cuts o the fuel supply for the engine.

NOTICE

Throttle positions for 115% and 100% can be selected by feel. The rst stop is the 100% throttle position. To select 115% the throttle lever must be moved through a throttle gate to the left and then pushed to the next stop.

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3.2.3 CONTROLS AT THE INSTRUMENT PANEL

The following overview includes controls in the lower area of the instrument panel. These elements are included in gure 3.2.3.a Arrangement of Elements on the Instrument Panel, see section 3.2.4.

No. Name of Control Airframe Position

1. Emergency Canopy

Release

2. Cowl Flap

Adjustment

3. Nose-cone

Operation

4. Propeller Brake Black T-handle to the RH side of the cowl ap adjustment to brake

5. Propeller

Positioning

6. Air Vents Two adjustable air vents for cockpit ventilation, one on LH and one

7. Canopy

Ventilation

Red pull-handle on LH side of the switch panel. It is pulled for emergency canopy jettison after opening the canopy locks on LH or RH side of the canopy frame.

Black T-handle on LH side of the lower middle section of the instrument panel to reduce engine cooling in cruise condition. The foremost position means cowl aps fully OPEN, 5 settings aft are available to reduce the opening of the cowl aps.

Black handle in the bottom center of the instrument panel to open, close and lock the nose-cone, linked to the engine electric master switch. Unlock by lifting, lock by pushing down the handle. In the forward position (Nose-cone OPEN) the engine master switch turns ON when nose-cone is LOCKED.

the propeller to a full stop after engine has been switched o in ight. Braking is by pulling the handle.

Black T-handle to the RH side of the propeller brake to position the propeller to t within the nose-cone contour. Operation is by a slow and steady pull on the handle to its stop.

on RH side of the Instrument panel, are provided. Knob on RH side of the ignition/starter switch to ventilate the can-

opy. The pulled position means canopy ventilation OPEN.

3.2.4 INSTRUMENT PANEL

The following description gives an overview of instruments, controls, monitor devices and CB’s installed on the instrument panel. The minimum Equipment List of the S12 Aircraft Fligth Manual is relevant. The positions of the elements are shown in gure 3.2.3.a Arrangement of Elements on the Instrument Panel, valid for the aircraft serial number as indicated on the title page of this Aircraft Maintenance Manual.

The ight control instruments include at least

1. One ASI (airspeed range 50 – 300 km/h / 27 – 162 kt),

2. One Altimeter,

3. One magnetic compass,

4. Trim indicator display.

These instruments are located directly in the view area of the PIC (in front of LH seat). Dual instrumentation is possible to provide an optimum view on ight control instruments from the RH

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seat (e. g. instruction ights). Additional avionics may be installed on customer demands and concurrent applicable approval.

Engine monitoring includes at least:

• Tachometer

• Oil pressure and oil temperature

• Cylinder head temperature (CCT) LH and RH

• Voltmeter and ammeter

• Fuel quantity in LH and RH wing tanks

• Engine hour-meter

These instruments are located as a rule, with the exception of the engine hour-meter, in the RH area of the panel, if not installed (i. e. with dual ight instrumentation) in the center area. The engine hour-meter is located on the center console between the seats.

The red re-warning light (test by pushing light for optic and acoustic signal) is adjacent to the engine instrumentation.

The following warning and monitoring lights are combined in a group, arranged independently of its location on the instrument panel. They inform the pilot about the proper condition of the aircraft at a glance. The group is always located at the upper instrument panel below the glare shield to allow for glare-free reading.

The following lights are arranged:

A. Red fuel pressure warning,

B. Yellow low fuel warning,

C. Green status indication for fuel aux pump operation,

D. Red warning light for manifold pressure (boost pressure),

E. Yellow caution light for malfunction of TCU,

F. Red warning light for malfunction of external alternator (battery charge control),

G. Yellow caution light for malfunction of internal generator.

The landing gear position and warning indication is also located below the glareshield and consists of two lights, indicating the situation with green or red steady or ashing light.

The following CB´s are combined in a group, arranged independently of its location at the panel:

• Master CB,

• CB for external alternator,

• Landing gear CB,

• CB´s for main, transfer and auxiliary fuel pumps.

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The lower center section of the instrument panel comprises a row of levers and switches. Any switches, except a switch to select voltmeter indication at the upper RH panel, are systematically arranged here. The red handle for canopy emergency jettison is installed LH of the row of switches. Sequence of levers and switches, starting from the left, is:

1. Red handle for canopy emergency jettison,

Landing gear lever with three positions (down: extension, center: neutral (electrically de-energized), up: retraction),

3. Switch for auxiliary fuel pump,

4. Fuel selector switch (positions “LEFT”, “BOTH”, “RIGHT”)

5. Electric master switch,

TCU emergency switch to isolate waste-gate actuator and TCU control in case of malfunction (switch is guarded with a red protecting plate for unintended operation),

The control elements for propeller and nose-cone (propeller brake, propeller positioning and nosecone handle) are arranged below the row of switches in the center console. In the same area the canopy ventilation knob and the ignition/starter switch (positions OFF, Right, Left, BOTH and START) are installed.

The following gure 3.2.3.a Arrangement of Elements on the Instrument Panel shows layout and arrangement of the instrument panel of the aircraft serial number as indicated on the title page, including control elements, monitoring devices and CB´s.

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EXAMPLE

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Figure 3.2.3.a

Arrangement of Elements on the Instrument Panel

(Related to S/N Indicated on Title Page)

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3.3 FLIGHT CONTROL SYSTEM

3.3.1 LONGITUDINAL CONTROL

Both control sticks are coupled by a connection tube. The control movements are transmitted via push-pull rods to the end of the tail boom and then straight up to the elevator tting. In the tail boom, the push-pull rod is supported by linear motion ball bearings. Adjustable stops for the longitudinal control are installed in the middle of the connection tube beneath the right control system cover in the cockpit. A control tube connection in the tail boom is linked to the spring trim system.

3.3.2 LONGITUDINAL TRIM 

Longitudinal trim is achieved by means of a spring system. The spring forces are adjusted by an electrical actuator. This actuator is controlled by a switch on the control sticks. The actuator together with the spring system and the linkage to the elevator control system are situated in the front part of the tail-boom.

3.3.3 WING FLAP CONTROL SYSTEM

Both ap control levers are coupled by a connection tube beneath the control system cover in the cockpit. Control inputs are transmitted from this connection tube via push-pull rods to a “mixing shaft” in the center fuselage. From this “mixing shaft”, the control inputs are transmitted via bell-crank levers, push-pull rods and quick release couplings to the control rods in the wing. The control rods in the wing are supported by means of linear motion ball bearings. The control movement is transmitted to the ap drive ttings via bell-crank levers. At the mixing unit, the wing ap control is supported by a gas spring against the center steel frame. This is to minimize loads on the ap lever at appropriate airspeeds related to the respective ap setting. A viscous damper acting in both directions isolates shock loads from the ap lever and the ap lock-in. The ap positions are locked at the output lever on the connection tube in a gate beneath the control system cover.

3.3.4 LATERAL CONTROL

Lateral inputs with the control sticks are transferred via an adjustable push rod to a center bell-crank lever beneath the connection tube of the longitudinal (elevator) control. From this bell-crank lever, control inputs are transmitted via push rods to the “mixing shaft” in center fuselage. Via this “mixing shaft”, bell-crank levers, push-pull rods and quick release couplings, the control rods in the wing are moved.

Both sides of the wing center section contain a straight-through control rod, supported by several linear motion ball bearings and equipped with a quick release coupling (L´Hotellier) at the division of inner and outer wing. From the push-pull rods in the outboard wing, the control movements are transmitted via two bell-crank levers to the drive ttings of the ailerons.

By means of the “mixing shaft”, the ailerons are moved together with ap position changes and the aps are moved together with aileron deections (“aperons”). The percentage of comovement depends on the position of the control surfaces. Stops for aileron control inputs (adjustment screws) are located LH and RH side of the elevator connection tube beneath the covers of the control system well in the cockpit.

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3.3.5 CONTROL OF AIR BRAKES

The air brake-levers are coupled by means of a connection tube. Travel of the levers is transmitted via push rods and bell-crank levers to a driving lever (over-center strut) in the center fuselage, from which it is transmitted via push rods and quick release couplings to push-pull rods in the wing, which then move the air brakes. The push rods in the wing are supported by linear motion ball bearings.

The air brakes are locked in retracted position by over-center-locking of the over-center strut. The locked position can be adjusted by a stop screw at the over-center strut. The fully extended position is determined by a xed rubber stop, which one end of the driving lever butts upon when this position is reached.

3.3.6 DIRECTIONAL CONTROL

From the LH and RH rudder pedal supports, control cables are led through the center fuselage to the tail boom behind the luggage compartment, where the control cables of the left pedals and the right pedals meet to be directed further to the rudder driving lever. The rudder driving lever is connected to the tail wheel by a spring assy. The stops for the directional control are mounted on the lower rudder support, the pertinent adjustment screws are located at the rudder on the drive tting.

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3.4 POWERPLANT

3.4.1 ENGINE

MAINTENANCE MANUAL STEMME S12

Engine description

See Maintenance Manual (Heavy Maintenance) for ROTAX Engine Types ROTAX 912 and 914 Series, section 71.

Upper (rear) Engine Attachment

A tubular steel frame at the upper engine anges, mounted in two vibration absorbing elements in the upper lateral tube of the fuselage frame.

Lower (front) Engine Attachment

A separate tubular steel frame at the lower engine flanges in two vibration absorbing elements in the forward lateral framework junctions.

Type ROTAX 914 F2/S1 (modied ROTAX 914 F Aircraft

engine, propeller shaft design with ange for xed pitch propellers).

The ROTAX 914 F2/S1 is modied by STEMME and based on the type ROTAX 914 F2. The modication was completed as a co-operation with ROTAX and with approval by ROTAX.

The base engine ROTAX 914 F2/01 is certied according to JAR-E / FAR 33. It is turbocharged and has an electronic dual ignition system.

The modied version ROTAX 914 F2/S1 was specially developed for the model S12, a derivative of the motorglider STEMME S10.

The engine modications are certied together with the S12 according to JAR 22. The modied version ROTAX 914 F2/S1 has the STEMME (internal) production No. 11AM-M.

Due to special requirements for installation in the center fuselage, the following modications were made:

•

Relocation of the turbocharger unit to the aft of the engine to stay within the outlines of the S12 fuselage. The turbocharger unit is supported by ve struts aft of the engine. A supercharger inter-cooler is installed. Relocation of the turbocharger unit required modications of the exhaust system. The exhaust headers are attached to the turbocharger by springs, positioning them in spite of the high thermal stress.

• The layout of the oil pipes for the turbocharger are modied.

•

The exhaust headers and the muer are shrouded by temperature-resistant material, thus thermally isolating the system from the engine bay. The turbocharger unit and the air box are isolated by radiation protective shields.

• The original ROTAX engine mounting is not used. STEMME specially developed an engine mounting for a center installation, consisting of two upper and one lower supporting element.

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• In the liquid cooling system, the combined function of the expansion reservoir and rell container was split up in two separate containers, with the expansion reservoir located above the engine and the rell container on the left side of the re-wall.

•

The ignition unit is installed above of the engine slightly behind the original ROTAX-position.

• The throttle levers on the carburetors have been slightly modied (modication does not aect the throttle rigging of the original engine) and additional springs have been installed (pulling towards full-power position) to compensate for friction due to the long control cables between cockpit and engine bay.

Engine performance data of the ROTAX 914 F2/S1 are identical to those of the ROTAX 914 F.

ENGINE PERFORMANCE DATA Max. T/O power. 84.5 kW at 5800 RPM 113.2 hp Max. continuous power. 73.4 kW at 5500 RPM 98.4 hp Engine reduction gear ratio i = 2.4286 Max. propeller 2650 RPM

3.4.2 LUBRICATION SYSTEM

SYSTEM DESCRIPTION

Refer to Maintenance Manual (Heavy Maintenance) for ROTAX Engine Types ROTAX 912 and 914 Series, section 79.

The ROTAX 914 is equipped with a dry-sump pressure lubrication system. The pumps are part of the engine. It was necessary to modify the routing of the oil pipes between oil pump and turbocharger. Oil is returned from the crankcase to the oil tank by compression blow-by pressure, and from the turbocharger via the secondary oil pump (scavenge pump). The oil tank is standard from ROTAX. It is installed behind the re-wall on the LH side of the fuselage frame. The oil ller cap is below a service access in the upper center fuselage fairing.

Oil is taken in exible pipes from the oil tank through the re wall to the engine bay. Any exible oil lines in the engine compartment are shrouded by re protective sleeves. The oil cooler is installed on the RH side of the fuselage frame and supplied with air from the RH cowl ap.

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3.4.3 COOLING SYSTEM

SYSTEM DESCRIPTION

Refer to Maintenance Manual (Heavy Maintenance) for ROTAX Engine Types ROTAX 912 and 914 Series, section 75.

Engine is cooled by liquid cooled cylinder heads, ram air cooled cylinders and by cooling the oil.

LIQUID COOLING

The radiator for the coolant is installed on the LH side of the fuselage frame, supplied by ram air from the LH cowl ap. The rell container is installed on the forward left side of the upper re wall. A thin tube links the relief valve in the rell container with the overow container, which is installed in the LH landing gear bay. The quantity of the coolant can be checked at the scale on the overow container and must be between min and max markings. The coolant lines consist of exible coolant hoses and aluminum tubes.

RAM AIR CYLINDER COOLING

Ram air for cylinder cooling is guided from the RH cowl ap through a duct to a distribution plenum on the engine upper side. The distributor is made of special heat-resistant GFRP.

INLET COWL FLAPS

Oil cooler, radiator, inter-cooler, ram air cylinder cooling and the carburetors are supplied with ram air from the LH and RH cowl aps in the lateral engine cowlings. To avoid low engine temperatures during high cruising speeds or descents from high altitudes, the cowl ap aperture can be reduced with a handle on the instrument panel. During T/O, climb and at high OAT´s, the cowl aps should be fully open. Five reduced apertures can be set for dierent conditions. It is not possible to close the cowl aps completely as long as the nose-cone is open. The cowl ap control is linked to the nose-cone operation: closing of the nose-cone also closes the cowl aps and opening of the nose-cone opens the cowl aps to the position set by the cowl ap handle.

OUTLET COWL FLAP

The warmed air streams exit the engine compartment via the lower outlet cowl ap. The aperture of the outlet cowl ap, like the inlet cowl aps, is controlled by the nose-cone operation as well as by the cowl ap handle on the instrument panel. The shape of the small fuselage fairing behind the outlet cowl ap maintains a slot when the cowl ap is closed with the nose-cone, to allow a permanent ventilation of the engine compartment.

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3.4.4 AIR INDUCTION SYSTEM

SYSTEM DESCRIPTION

Refer to the Maintenance Manual (heavy maintenance) for ROTAX Engine Types ROTAX 914 Series, section 73.

The engine induction air is routed from the RH inlet cowl ap via a duct, an air lter, and a short exible tube to the turbocharger unit. Compressed air from the turbocharger is routed via two tubes through the inter-cooler to the airbox. The inter-cooler, with a drain at its lowest point, is installed on the RH side of the center fuselage frame. Cooling air for the inter-cooler is taken from the RH inlet cowl ap, directed into the engine bay and to the outside via the lower outlet cowl ap.

The airbox is supported by two rubber mounts on the diagonal strut of the turbocharger unit. The airbox drainage lines are routed to the outlet cowl ap, in order to allow drained fuel to leave the engine compartment.

3.4.5 ENGINE EXHAUST INCL. TURBOCHARGER AND ATTACHMENT

SYSTEM DESCRIPTION

Refer to the Maintenance Manual (heavy maintenance) for ROTAX ENGINE TYPES 914 SERIES, section 78.

TURBOCHARGER ATTACHMENT

Five steel tube struts at aft engine mounts. The turbocharger unit, consisting of turbocharger and muer, is installed aft of the engine and supported by ve struts on the crankshaft housing. The turbocharger is supplied with oil by oil suction and oil pressure pipes from the engine oil pump.

The exhaust headers are installed below the engine, constructed using original ROTAX connecting elements. The headers are isolated from heat sensitive components. The exhaust headers are attached to the turbocharger by springs, positioning the headers in spite of the thermal eects. The exhaust headers are shrouded by a temperature-resistant glass-ber-tape (Frenzelit Isotherm

1000). The tape terminals are xed by special spring-loaded brackets.

The muer is covered by formed elements of an insulation material (Frenzelit „Isosafe“ 12 mm). To avoid heat radiation, heat protective shields (stainless steel) are installed close to the turbocharger unit. The exhaust pipe of the muer is routed through a hole in the LH engine cowling. In order to protect overowing fuel from contact with hot exhaust system parts, drain lines from the drip trays below each carburetor are routed to the lower cowl ap.

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3.4.6 FUEL SYSTEM

3.4.6.1 FUEL SYSTEM RELATED CONTROLS AND INDICATION SYSTEM

The S12’s cockpit provides the following fuel system related controls and indication systems:

The fuel shut-o valve (horizontal ”OFF”; vertical ”ON”) is located on the rear center console.

•

The fuel selector switch (with positions ”LEFT”, ”BOTH”, and ”RIGHT”) is located on the instrument panel.

• The switch for the auxiliary fuel pump (with positions ”ON” and ”OFF”) is located on the

instrument panel.

• The fuel pressure indication, the low-fuel light for the feeder tank as well as the green sta-

tus light for the operation of the auxiliary fuel pump, are located on the instrument panel.

• The fuel quantity indicators for the two wing tanks are on the instrument panel.

• The four fuel pumps (one main and one auxiliary pump plus one transfer pump from each

wing fuel-tank) have separate fuses, which are located in the Electronic System Boxes on top of the steel frame.

NOTICE

For the location of the listed switches, lights or indicators please refer to the S12 Aircraft Flight Manual, section 7.3 Instrument Panel.

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3.4.6.2 FUEL SYSTEM DESIGN

WING FUEL SYSTEM

Both wing tanks are located in the outer sections of the wing center section. The ends of the vent lines are located at the lower, outer ends of the wing center section. The ller caps are located on the upper surface of the wing center section. The fuel line from each tank includes a strainer, a ne lter and quick drain as well as a piston pump including a fuel-check valve for each tank.

All these components are accessible through an opening in the lower wing skin near the root rib. Depending on the position of the fuel selector switch and a oat switch in the feeder tank, the feeder receives fuel from either the right or the left tank separately, or from both tanks simultaneously. A single quick release connector is located between the fuel line from the wing and the corresponding connection in the fuselage.

3.4.6.3 FUEL SYSTEM OF THE FUSELAGE

FUEL SYSTEM OF THE FUSELAGE

Beginning at the quick release connector, fuel ows directly to the feeder tank. A oat switch in the feeder tank prevents feeder tank overlling. The feeder tank is vented through the right shell of the forward fuselage above the landing gear door at the rear edge. The feeder tank has a sump at its bottom that is connected to a quick drain. From the lower part of the feeder tank (approximately 1 cm above the bottom) the fuel is pumped by the main fuel pump (powered by the engine electric system) or the auxiliary fuel pump (powered by the main electric system) to the engine.

The main pump runs continuously whenever the engine electrical system is powered (nose-cone is pushed forward and locked). A fuel return line delivers surplus fuel back to the feeder tank. Contaminated fuel (i.e. particles, water, etc.) from the sumps can be drained by use of the quick drain in the landing gear bay or by the two quick drains in the wing tanks.

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3.4.6.4 ELECTRICAL LAYOUT AND OPERATION OF THE FUEL PUMPS

For normal operation, the auxiliary fuel pump is switched o, and only the main pump is operating. The main pump is electrically supplied by the internal generator of the engine, or the engine battery respectively.

The auxiliary fuel pump can be added if necessary, for example during take-o and landing. It is supplied by the external alternator, or the main battery respectively:

The right wing tank pump operates with the fuel selector switch in position ”RIGHT” or

•

”BOTH” tanks and the oat switch in the feeder tank switching the pumps “ON”.

• The left wing tank pump works with the fuel selector switch in position ”LEFT” or ”BOTH”

tanks, plus the oat switch in the feeder tank switching the pumps “ON”.

• The main pump works when the engine electrical system is powered.

•

The auxiliary fuel pump works when the main electrical system is powered and the auxiliary fuel pump switch is selected “ON”.

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Fuel-Tank Valves

Coarse-Filter

Fine-Filter

Right Transfer-Pump

MAINTENANCE MANUAL STEMME S12

Right Fuel Tank

Drainer

Sump

Pump o - Switch

Low fuel - switch

Feeder Tank

Drainer

Pressure-Dierence-Sensor

Fuel-Pressure

Air-Box-Pressure

Left Transfer-Pump

Fine-Filter

Coarse-Filter

Fuel-Tank-Vents

Check-Valves

Sump

Quick-Disconnect-

Fittings

Drainer

Main-Fuel

Aux-Fuel

Pump

Air-Box

Pressure-

Regulators at

Carburetors

Check-Valve Check-Valve

Engine-Compartment

Fuel-Shut-O-Valve

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Left Fuel Tank

Figure 3.4.6.a.

Diagram of the Fuel System

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3.4.7 ENGINE CONTROLS AND INSTRUMENTATION

REFER ALSO SECTION 3.2.2

The throttle and choke are operated by Bowden cables, linking the carburetors with the cockpit levers, which have adjustable friction control. The throttle and choke lever are installed in the center console between the pilots. The throttle lever has two stops, one for the 100% (max continuous) power position and one for the 115% (max takeo) power position.

Engine cooling is controlled by the cowl aps, opened and closed by a linkage to the nose-cone handle. In addition, the cowl ap aperture can be reduced in ve steps by the cowl ap operation handle on the instrument panel (see section 3.4.3).

Both engine ignition and engine starter are controlled by a key switch. A special feature is a three-second ignition delay after switching the key to position “START”, to allow the propeller blades to unfold before engine ignition (see section 3.7.11).

The engine instruments are located on the RH instrument panel. For positions and function of switches and instruments, refer to section 3.2.2 and 3.7.

3.4.8 FIRE PROTECTION

The engine including exhaust and induction system is isolated from the front, top and rear areas by rewalls. The rewalls consist of stainless steel sheets (thickness 0.38 mm / 0.015 in).

The insides of the engine cowlings on the LH and RH side of the engine and below the engine are coated with a re-protective painting (manufacturer: Courtaulds Aerospace). It consists of three coats of a white re resistant paint (type N 56582/T508) and one coat of clear varnish (type 4232-0303, hardener N39-1327 or N50/2509, thinner N39-3091). The fuel and oil hoses in the engine compartment are covered by re-protective sleeves.

The re warning system is triggered by two bimetallic temperature sensors above the carburetors. The re warning is by means of a red warning light (push-to-test function) on the instrument panel, combined with an acoustic warning tone via the loudspeaker.

3.4.9 ENGINE COWLINGS

COWL FLAPS: REFER TO SECTION 3.4.3

The engine cowling consists of three, (LH, RH and lower) parts. The cowlings are self-supporting and connected to the fuselage structure forward and aft by means of camlocks. The LH and RH cowlings are each fastened to the structure by 5 camlocks. Each includes an inlet cowl ap and air duct assy.

The cooling air duct of the LH cowling maintains a gap of approx. 5 mm / 0.19 in to the radiator with the cowling installed, to allow for relative motion.

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The air inlet of the RH cowling distributes the air for cylinder ram air cooling, for the intercooler and for the oil radiator.

When installing the RH cowling, the opening for the cylinder cooling is pressed against a rubber sealing on the coupling sheet of the upper support of the cylinder cooling air hose, routed to the distribution plenum for the cylinders. The induction air duct is prepared with a cut-out for the induction air lter. The lter assy. on the turbocharger inlet, consisting of lter, sealing plate and air hose, remain installed while cowling is removed.

The complete lower cowling forms the outlet cowl ap. It is fastened by its forward hinges and positioned by a rear control cable assy., which counteracts two springs acting towards the open position of the ap.

3.4.10 PROPELLER

The foldable propeller consists of a center hub and two propeller blades hinged to this central unit. The folding axis allows the propeller blades to move in the plane of propeller rotation. When the propeller is not rotating, the blades fold in by spring forces. The central part of the propeller is made of high strength aluminum. The propeller blades consist of carbon, Kevlar and glass composite.

During engine starting, the blades unfold automatically by centrifugal force. Rubber stops protect the blades in case of a possible over-swing. The fully folded position of the propeller blades also has rubber stops. The propeller blades may be retracted at any possible pitch angle.

After engine shut-down in ight, the propeller must be stopped by the propeller brake to allow the blades to fold in. The brake is operated by the propeller brake handle on the instrument panel. In the folded position the propeller can be positioned with the propeller positioning handle on the instrument panel to allow the nose-cone, which forms the fuselage nose, to retract to the closed position. After closing the nose-cone with the nose-cone handle in the center part of the instrument panel, the propeller is completely enclosed within the contours of the fuselage to achieve optimum soaring performance.

The propeller blade pitch can be changed from take-o (ne pitch) to cruise (coarse pitch) position. The pitch control is electrically actuated and operated by a switch on the center console behind the throttle assy. The take-o position of the propeller blades (not merely the switch) is indicated by the green light next to the switch.

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3.4.10.1 DESIGN OF VARIABLE PITCH PROPELLER

The numerical positions in the following text refer to the propeller diagram (see gure 3.4.10 a/b).

The propeller blades (1) are hinged in a forked mounting plate (4). The complete assembly, consisting of the fork, blade and hub, is rotated in order to set the pitch angle of the blade. The hollow axle (3) houses a spiral torsion spring (23) to fold the blade by means of a cam lever (22) that ts into the aperture for the buer-stop in the propeller blade.

Electrically heated expanding servo-elements (15) actuate the blade pitch mechanism. On achieving their activating temperature, these expanding servo elements drive a piston connected to the propeller blades. All components of blade angle mechanism are double-redundant and mechanically interconnected by a coupling ring (12) so that both propeller blades always have identical pitch.

3.4.10.2 PROPELLER BLADES

The propeller blades are manufactured from GFRP (Glass Fiber Reinforced Plastic) material in a twin shell construction. The shells are of hybrid laminate type (glass, carbon, aramid). Polyurethane tape is axed to the leading edges to improve protection against gravel.

To balance the pressure in the blades with the outside atmosphere there is a 1 mm bore hole in the blade tip. These holes also drain condensed moisture by centrifugal force.

3.4.10.3 PROPELLER BLADE ANGLE CONTROL MECHANISM

When disconnected from the power supply (unheated), the dilation eect actuator (15) is pushed inwards via the rocking lever (14) by the spring (20), thus moving the propeller blade into T/O position via the push-rod (13), the synchronizing ring (12) and the connector (11).

Heating of the servo-element by the heating element (16) generates piston pressure that pushes the swing-arm/push-rod system into the almost fully extended position, thus rotating the propeller blade against the spring towards high pitch. It is assured, that, in case of heating element malfunction, propeller blade position is automatically in T/O under all operating conditions.

The transmission of the actuating forces to the fork is eected by the drive pin (10) which is attached to the fork with an eccentric bolt to allow for accurate setting of propeller blades. The available range of adjustment of the servo-elements is 12.1 mm / 0.476 in. The stop setting for the T/O position is adjusted for the fully retracted piston. The range of the blade angle change is set to 6° 24’, limited by a stop switch.

A mechanical stop is adjusted so that only minimum override of the end position switch setting is possible (< 6’). The piston movement caused by the two-point control system, result in minor changes in propeller blade pitch angle with neglectable inuence on the propeller behavior.

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* The numbering of the position in accordance with the “Propeller Adjustment Report”.

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See

Figure 3.4.10.b.

Figure 3.4.10.a

Variable Pitch-Propeller

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Position

Part

Number

1 Propeller blade 2 Needle bearing (articulated joint) 3 Hollow axle 4 Blade suspension fork 5 Thrust plate 6 Propeller hub 7 Grooved ball bearing 8 Axial needle bearing 9 Grooved ball bearing

10 Drive pin

11 Coupling Part 12 Synchronization ring 13 Push rod

14 Rocking lever

15 Dilation eect actuator

16 Heating element

17 Slip ring (+) 18 Slip ring (-) 19 Fly-weight 20 Compression spring

21 Rubber stop 22 Cam lever 23 Torsion Spring 24 Stop 25 Adjustment screw 26 Yield bolt 27 Base 28 Nut 29 Ball washer

NOTICE

In order to prevent incorrect assembly, all parts relevant for the variable pitch of one propeller blade are marked with a red dot. All other parts remain unmarked.

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3.4.10.4 PROPELLER WEIGHT BALANCE

The variable pitch propeller is balanced by the manufacturer. Final balancing is accomplished by means of balancing weight washers fastened on the base-plate of the pitch change unit. The positions provided on the base-plate for fastening of the washers are shown in gure 3.4.10 c.

Figure 3.4.10.b

Propeller for Mounting

Reference Point for Propeller Adjustment

Red Side

* The numbering of the position in accordance with the “Propeller Adjustment Report”.

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Figure 3.4.10.c

Position of Balancing Weights

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WARNING

Balancing the variable pitch propeller or its assemblies may only be performed by the manufacturer or by authorized and licensed maintenance facilities according to the specic instructions and using appropriate equipment. The positions, number and weights of the balancing weight washers of the individual propeller must be entered in the maintenance form Propeller Adjustment Report (Annex D).

3.4.10.5 EXPANSION ELEMENT AND CONTROL SYSTEM

The thermo expansion actuators are heated electrically and cooled by convection and heat dissipation. Electric energy is transferred to the propeller blades by slip rings. The expansion elements actuate only towards the high pitch angles. The propeller blades are rotated in the direction of lower pitch position by the return spring at the lever mechanism assisted by mass and aerodynamic forces on the blades.

When the propeller switch in the cockpit is in „cruise“-position, the expansion elements are heated with about 50 Watts (with the engine running). The elements are heated until a temperature of 55º C is achieved and the element expands. As soon as the stop is reached propeller blades are in cruise position and are held in this position by the two-point control system. If the propeller switch is set to T/O position, the heating elements are switched-o and the expansion elements cool down. By duplicating any essential elements and realization of logic-OR the control system is fully redundant.

Heat insulation of operating elements is optimized to insure that the time for full change of pitch in either direction will not exceed 5 minutes under any operating conditions (from idle to full power). Typical actuating time for full change of pitch is about 2 ½ min.

The electrical energy to the expansion elements is interrupted (blade pitch automatically to T/O) as soon as the propeller switch is set to T/O or when the nose-cone handle is not locked in OPEN position (engine master switch is OFF).

If the nose-cone is closed, electrical energy to engine and propeller is isolated to save energy when soaring.

In order to check the pitch range mechanism, a push-button that can be operated from outside is provided on the right hand side (in ight direction) of the front gearbox frame. With the master switch “ON” and the button pushed and held, the expansion elements are heated and the propeller is brought into the cruise setting. When the cruise setting is achieved, the stop switch disconnects the electrical current to the heating elements. For this check, the push-button must be activated for about 3 to 4 minutes to heat the expansion elements. On releasing the push-button, the elements cool down and the propeller returns in the start position.

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3.4.10.6 IDENTIFICATION OF PROPELLER SYSTEM AND ASSEMBLY

Complete Propeller 11 AP V / XXXXX / YYYY Propeller Blade 11 AP VB / XXXXX - ZZ

X Number of ordering for production share (corresponds to the aircraft serial no. for complete propeller); 5 digits, starting with 30001, in addition 20038;

Y Month and Year of production, 4 digits;

Z Consecutive number of production share, 2 digits.

3.4.11 POWER TRANSMISSION

The components of the power transmission system are:

•

Centrifugal clutch (in between engine and drive shaft): A force transmitting clutch actuated by direction of rotation and rotating speed. Since the centrifugal clutch transmits torque by friction, it is also an overload protection.

• Drive shaft.

• Flexible coupling forward and aft, allowing for angular and torsional elasticity.

The rear terminal of the drive shaft is maintenance-free due to a hard lm coating (30 m), which also is anti-corrosive. The sliding friction coecient decreases with increasing torque by design. Optimum behavior of the coating is only attained if not greased. The condition of the coating should be checked regularly (refer to section 5.4.16).

3.4.12 FORWARD PROPELLER REDUCTION GEARBOX

The propeller reduction gearbox utilizes helical gears in an cast aluminum housing with a gear ratio of 1.109. It is almost maintenance-free. Propeller reduction gear oil level can be checked by means of an inspection glass with MIN/MAX markings. The extension shaft is connected to the lower gear drive shaft with a exible coupling. The propeller is mounted on the upper propeller shaft of the reduction gearbox.

The reduction gearbox is attached to a milled aluminum suspension, which is supported by 4 shock mounts in the fuselage front frame to reduce vibrations. The cut-outs in the front frame for the shock mounts are sealed by rubber adapter shapes.

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3.4.13 OPERATION MECHANISM OF THE PROPELLER FOLDING SYSTEM

3.4.13.1 NOSECONE

The nose-cone forms the fuselage nose in front of the cockpit section and covers the folded propeller during ight in glider conguration. To change from gliding to powered ight, the nosecone can be shifted forward to allow the propeller unfold and rotate after starting the engine. In the forward position locking and unlocking of the nose-cone lever simultaneously operates the engine master switch (see section 3.7.11).

The nose-cone consists of a CFRP shell, a GFRP frame and an aluminum tube, which provides suspension of the nose-cone at the front fuselage. The aluminum tube is guided through the hollow propeller drive shaft into an aluminum tubular guided pipe, including Teon slip rings to allow axial sliding of the nose-cone tube. The tubular guided pipe is rmly bonded to the GFRP guide block in the forward cockpit.

Shifting the nose-cone is by means of the nose-cone handle, installed in the lower section of the instrument panel (see gure 3.2.3.a). The pull rod of the handle is connected to the nosecone sliding tube via a milled aluminum bracket, which is screwed to the rear tube terminal and guided in a slot at the guide block. The bracket and the handle guide assy. prevent rotation of the nose-cone. The handle assy. includes a spring which acts towards the locked (down) position of the handle.

3.4.13.2 PROPELLER POSITIONING

The propeller positioning assy. allows rotation of the propeller drive shaft after engine shut down and braking the propeller to a full stop, to the position so that the folded propeller ts into the nose-cone contour. It is operated by a pull handle on the lower instrument panel next to the propeller brake handle (see gure 3.2.3.a).

The propeller positioning is driven by a Bowden cable. The bowden cable is routed from the handle to and around an aluminum disc, which is pivoted on the guide block of the nose-cone sliding tube in the forward cockpit. When the handle is pulled, the disc rotates against a circumferential spring force until reaching a screw stop, which is adjustable and denes the nal position of the propeller. The spring is wound around the nose-cone guide block and rotates the disc back toward its initial position when the handle is released.

A spring-loaded catch lever is installed on the disc and is connected to the bowden cable rotating the disc. If not operated, the spring acts on the catch lever, so it is clear of the propeller brake drum, mounted on the rear propeller drive shaft of the front gear. When the handle is pulled the spring force unlocking the catch lever is over-balanced by the circumferential spring force acting on the disc and the catch will engage in one of two notches of the brake drum, thus rotating the drive shaft until reaching the screw stop of the disc. The two notches of the brake drum are opposed; therefore the propeller rotates a maximum 180° as the handle is pulled to the stop.

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3.4.13.3 PROPELLER BRAKE

The propeller brake is used to decelerate the propeller rotation to a full stop after engine shut down in ight, to allow folding of the propeller by spring forces. It is operated by a pull handle on the lower instrument panel (see gure 3.2.3.a). The propeller brake is driven by a bowden cable, routed from the handle through the front fuselage frame to the propeller brake assy., which is mounted on the rear housing of the front gear.

The cable is connected to a spring-loaded shift lever, which is mounted on the lower front gear housing and pulls down one terminal of a brake belt during operation of the brake. The brake band consists of a steel sheet with a riveted brake lining. The brake band is installed around a brake drum mounted on the propeller drive shaft of the propeller gearbox. The second brake band terminal is fastened to the gear housing opposite the other, driven, terminal. Some adjustment of the clearance between brake lining and drive shaft is possible by loosening and shifting the bowden cable support.

3.4.14 FLUIDS

3.4.14.1 FUEL

It is recommended to use premium gasoline, unleaded, minimum RON 95.

Suitable fuels:

• Super DIN EN 228

• Super Plus DIN EN 228

ACCORDING DOT

Gasoline min. grade 1, AKI 90.0, acc. to Canadian General Standard Board CAN/CGBS 3.5 (Unleaded Automotive Gasoline)

ACCORDING FAA

Standard Specication for Automotive Spark-Ignition Engine Fuel, ASTM D 4814

AVGAS 100LL

When using AVGAS 100LL, valve seats are stressed by the high amount of lead and in addition combustion chambers will accumulate residues. AVGAS should be used only in case of fuel vapor problems or if other fuel is not available.

The engine manufacturer recommends not to use AVGAS for an extended period, because an increased amount of residues may accumulate in the engine.

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CAUTION

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CAUTION

Danger of fuel vapor lock when using “winter” fuel during summer time.

CAUTION

Use only the appropriate fuel, recommended for the climate zone.

NOTICE

In accordance with the EASA SIB 2011-01 R2 “Unleaded Aviation Gasoline (AVGAS) UL91” is also approved for operation. (See STEMME SB A31-10-104 latest approved revision).

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3.4.14.2 COOLANT FLUID

Mixture of 50% concentrated antifreeze agent with anti-corrosion additives and 50% water. Freezing point of this mixture is about -38°C / -34°F. BASF Glysantin Antikorrosion has proved to be good; this or equivalent coolant may be used.

WARNING

As the cooling system is pressurized, to avoid the risk of scalding by escaping hot coolant, do not open the locking cap on the coolant expansion tank unless the engine is cold (or barely lukewarm).

NOTICE

Waterless coolant is not approved for operation.

CAUTION

Check of the coolant uid: The quantity in the overow container (lower left side in the landing gear bay) must be between “min.” and “max.” markings. Missing coolant in the overow container must be added.

CAUTION

If the level of coolant in the overow container is below “min.” marking, proper supply to the breather valve and coolant tank is not assured and air may have been introduced into the cooling system. To check , conrm that the engine is cold, open the locking cap on the expansion tank (left side on re-wall in engine compartment), and add coolant if necessary.

NOTICE

If the engine is warm, the indicated quantity in the overow container is noticeably higher. An excessive coolant level in the overow container will not result in engine damage, but will result in overow of coolant into the landing gear bay.

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3.4.14.3 LUBRICANTS

ENGINE WITH INTEGRATED REDUCTION GEAR

According to the ROTAX Operation Manual the following oil specication should be met:

• Use only oil with API classication SG or higher!

•

Due to the high stresses in the reduction gears, oils with gear additives such as high performance motor cycle oils are recommended.

• Heavy duty 4-stroke motor cycle oils meet all the requirements. These oils are usually not mineral oils but semi- or full synthetic oils.

•

Oils primarily for Diesel engines have insucient high temperature properties and additives which favor clutch slipping, and are generally unsuitable.

For more detailed information refer to appropriate ROTAX publications (SI-914-019 latest approved revision).

Oil quantity 0.8 US gal / 0.66 Imp. gal. / 3 l

(minimum 0.53 US gal / 0.44 Imp gal / 2 l)

Oil consumption max. 0.026 US gal/h / 0.022 Imp gal/h / 0.1 l/h

CAUTION

The reduction gear additives, specied “GL4” or “GL5”, are required for a safe lubrication of the integrated reduction gear. Never use other oil additives!

CAUTION

A full synthetic oil in combination with AVGAS results in abnormally high abrasion and/or residues. During utilization of AVGAS, only semi-synthetic oils should be used.

NOTICE

Full- or semi-synthetic oils are to be preferred because of the temperature stability and less residue formation.

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a) Viscosity:

Use of multi-grade oils is recommended. Viscosity of multi-grade oils is less depending on temperature compared to single-grade oils. Multi-grade oils can be used year-round; after an engine start at low temperatures the engine components are lubricated faster and at higher temperatures the oil is heavier. Temperatures of adjacent SAE-classes overlap, so for short-term temperature variations there is no need to change oil. The suitable oil grade can be chosen from the table.

Figure: 3.4.14.3.a

Engine Oil Table

PROPELLER GEARBOX

Use reduction gear oils specied according to MIL-L-2105C (Grade 90) or automotive gear oils HYP 80W (ARAL), HYP 85W-90, HYP 85W 140 according to API-GL5 or Shell Aviation Oil S 8350 according to DTD 900/4981.

NOTICE

Oil specication: Only use “GL5” oils according to the API-system.

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3.5 LANDING GEAR

3.5.1 MAIN LANDING GEAR

The landing gear (L/G) consists of a tail wheel and two retractable main landing gear legs, hinged at the center fuselage frame with the hinge axis in ight direction and locked in the extended position by means of an over-center locking strut (“elbow lever”) for each leg. The wheel is mounted on a trailing arm that is supported against the leg’s frame by a pre-loaded elastomeric spring for shock absorption purposes.

Retracting of the L/G legs and doors is managed by an electrically driven linear actuator for each leg that is built up around a high precision ball screw. Each of the linear actuators is hinged with the top end at the fuselage frame; the bottom end is coupled to the respective elbow strut by means of a locking mechanism which can be released for emergency landing gear extension by pulling a T-handle in the cockpit (one for each of the legs) and via a bowden cable. In case of an emergency extension the two legs should be released in succession (right-side-rst order is suggested, but wrong order is not critical), they then extend out by gravitational force. Secure locking in the extended position is achieved by a spring that forces the over-center strut into its end position.

The actuators are controlled and sequenced by stop switches, the switches for EXTENDED be

ing integrated in the elbow struts and detecting the over-center position, those for RETRACTED mounted at the fuselage frame and detecting the top position of each L/G leg. All these switches are in duplicate, the second one giving the signal for the indication and warning system, which is displayed by focused green and red LED’s on the right face of the instrument panel (refer to the S12 Aircraft Flight Manual).

Both L/G doors are actuated by the landing gear legs. The RH landing gear door is coupled directly to the RH landing gear leg via a spring device. The LH door is controlled by a cable mechanism. During retraction, the LH landing gear leg starts closing the LH door by means of a cable, but it remains open enough to allow retraction of the RH landing gear leg. The RH landing gear leg eects complete closing of the LH door via the cable during the last portion RH main gear retraction. Opening of the LH door is by a spring loaded roller strut, which rolls on the LH door. It pushes the door to the outside against the cable to keep the door from moving due to air loads, and is blocked with the landing gear retracted, thus locking the door. In closed position the doors are additionally locked at the rear by means of magnets.

The disk brakes on the main L/G wheels are operated hydraulically. The master cylinders for both the left and right wheel are located on the control sticks. The pressure line from the master brake cylinders to the brake calipers of the wheel brakes in the center fuselage are designed as metal-shielded brake hoses. The brake uid reservoir is located in the landing-gear bay, cabin rear wall. The parking brake valve to set and to release the parking brake is located on the oor panel console in front of the LH control stick. The parking brake valve is operated by a lever respectively rotary handle. Braking action is simultaneous on both main wheels. Maximum brake pressure for the system layout is 115 bar / 1668 psi, maximum allowed system pressure is 200 bar / 2900 psi.

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MAIN LANDING GEAR EMERGENCY EXTENSION

By pulling two T-handles in the cockpit, one after the other, the landing gear actuators are disconnected from the locking elbow struts by means of Bowden cables and the landing gear legs extend by gravity to the “gear down” position. The secure locking of each leg in the extended position is achieved by a spring that forces the over-center strut into its over-center position.

In case of emergency extension the two legs should be released in the recommended sequence (T-handles marked 1 & 2, wrong order is not critical). The RH gear strut is equipped with a mechanism to avoid jamming of the struts in case of an incorrect sequence.

3.5.2 TAIL WHEEL

The tail wheel is not sprung and ts in a trailing fork mounted in two bearings. The upper bearing is a combined radial/axial sleeve bearing. The journal is constructed so that a certain friction damping is produced at the axial sleeve surfaces when loaded in axial direction in order to avoid tail wheel utter whilst taxiing. For steering on the ground the tail wheel fork is coupled with the rudder by means of two pre-loaded tension springs.

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139 mm / 5.47 in

1360 mm / 53.54 in

Figure 3.5.1.a

Main Landing Gear

779 mm / 30.67 in

372 mm / 14.64 in

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3.6 FLIGHT CONTROL INSTRUMENTS, PITOT STATIC SYSTEM

Instruments: See Equipment List.

CAUTION

For modications sections 3.8 and 9 must be respected.

Pitot pressure, static pressure and total energy compensation are measured by means of a multi-purpose probe in the nose-cone. Its orices are plumbed to the instrument panel through the nose-cone sliding tube.

NOTICE

Some instruments, e. g. Bohli-diaphragm-type rate-of-climb indicator, need additional static ports for proper operation at the widest area of the forward fuselage. These additional static ports are installed only, if they were required for instrumentation at time of delivery.

All hose lines of the pitot-static system are equipped with a water separator/lter in the front cockpit area below the instrument panel.

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Figure 3.6.a

Pitot and Static Pressure System

NOTICE

Any line must be as short as possible. The water/separator lters must always be located upstream of the device and upstream of any junction.

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3.7 ELECTRICAL SYSTEM

3.7.1 GENERAL SYSTEM ARCHITECTURE

The electrical system of the S12 consists of two independent, separate sub-systems: the main system and the engine system. Each system incorporates a dedicated battery and generator (actually an alternator and rectier for the main system). Failures of one system will not interfere the other system. There is no provision to cross-tie the two sub-systems.

The engine system powers exclusively those devices needed to operate the engine or which are available only when the engine is running. Under normal conditions the engine system battery is charged by the internal generator of the ROTAX engine. In gliding ight the engine system is shut o completely. The operational mode of the engine battery is a stand-by mode, as it is normally used only for engine starting. Under normal conditions it will not be discharged.

The main system powers all other devices of the aircraft. During powered ight its battery is charged by the external belt driven alternator. During gliding ight the main system battery powers all active devices. Once the main system battery is exhausted, the pilot is able to restart the engine using the still fully charged engine system battery.

After successful restart of the engine, the external alternator will power the main system and will recharge the main system battery. The operational mode of the main system battery is as a storage battery in cycle mode. Under normal conditions it will be regularly discharged and recharged.

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MAIN SYSTEM

COM / NAV TRIM LAND. GEAR LIGHTS AUTOPILOT AUX.FUEL PUMP

EXT. ALT

ENGINE SYSTEM

ENGINE MAIN AND TRANSFER FUEL PUMP

ENG MONITOR TCU PROP PITCH CTRL.

MAIN SYSTEM BATTERY

INT. GEN

START

ENGINE SYSTEM BATTERY

Figure 3.7.a

Electrical System S12

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3.7.2 DC GENERATION ATA 243000

3.7.2.1 MAIN SYSTEM DC POWER GENERATION

The DC power for the main system is generated by the belt-driven alternator of the engine and stored in a LiFePO4 battery.

MAIN SYSTEM BATTERY

Type Super B SB12V19E-EC STEMME p/n 831.489 Nominal Voltage 12.8 V Nominal Capacity C1 19.2 Ah Technology Lithium Iron Phosphate LiFePO4 Protection device Over-voltage protection device OVP-15.5 Qualication Airworthiness and eligibility are proven in /A/ Location Battery case at front ring frame of tail boom

MAIN SYSTEM ALTERNATOR

Type Nippon Denso 42A, STEMME p/n 830.600 Drive V-Belt Excitation External, controlled automatically depending on running engine Field circuit protection 5 A fuse Nominal Voltage: 14 V Max Output Power 42 A Voltage Regulator built-in Regulated Voltage 14.2 – 14.8V @ 25°C Qualication The alternator is part of the certied aircraft engine ROTAX 914 series Location Left side of propeller gear housing of engine

FUNCTION

When aircraft master switch is switched ON the main system master relay connects the battery to the system. From that moment on all devices supplied by main system are powered, except those with a dedicated ON/OFF switch. When the engine is running, a pulse train representing the engine’s rpm is present at the input of a frequency-controlled relay. The relay switches on the eld circuit of the alternator and the alternator starts to generate voltage. The alternator output is controlled by a voltage regulator built in the alternator. A red warning light indicates failures at the alternator, e.g. a broken V-belt.

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3.7.2.2 ENGINE SYSTEM DC POWER GENERATION

The DC power for the engine system is generated by the crankshaft-mounted generator of the engine and stored in a LiFePO4 battery. The battery is used to crank the engine to start.

ENGINE SYSTEM BATTERY

Type Super B SB12V7800P-CC STEMME p/n 831.488 Nominal Voltage 13.2 V Nominal Capacity C1 7.8 Ah Technology Lithium Iron Phosphate LiFePO4 Protection device Over-voltage protection device OVP-15.5 Qualication Airworthiness and eligibility are proven in /A/ Location Battery case at front ring frame of tail boom

ENGINE SYSTEM INTERNAL GENERATOR

Type ROTAX (Part of engine) STEMME p/n n/a Drive crankshaft, directly mounted Excitation permanent magnet Nominal Voltage 14 V Max Output Power 18 A Voltage Regulator external rectier-regulator (ROTAX) Regulated Voltage 13.7 – 14.3 V Qualication The generator is part of the certied air raft engine ROTAX 914 series Locations Coils and magnets mounted on aft end of crankshaft housing, voltage regulator installed on TCU-Starter-Board.

FUNCTION

If aircraft master switch is switched ON and the aircraft nose cone is opened and locked, a micro switch at the nose cone lever activates the engine system master relay. The relay connects the battery to the engine system. The TCU and the main fuel pump operate immediately. Pushing the ignition-starter key to START activates the starter relay and connects the battery to the starter. When the engine is running the rotating magnets generate voltage in the generator coils. The regulator limits the output voltage. A yellow warning light indicates failures of the generator system.

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3.7.2.3 OVER VOLTAGE PROTECTION

The outputs of the belt driven alternator and of the internal generator are monitored by a dedicated over-voltage device OVP. The device senses the voltage of the related system and disconnects the related alternator resp. generator in the case of system voltages of more than 15.2 V.

Each OVP is located in the related distribution box. A yellow light at the box indicates a triggered over-voltage protection.

3.7.3 DC INDICATION ATA 243700

3.7.3.1 DESCRIPTION

Each electrical system includes a voltmeter to indicate the system voltage and an ammeter to indicate the current between system and battery. Both instruments can be toggled between main and engine system by a common toggle switch. The ammeter is supplied by shunts in the positive lines of the batteries. Positive reading indicates battery charging, negative reading indicates battery discharging.

A red warning light indicates failure of the main system alternator. A yellow warning light indicates failure of the engine system generator.

Voltmeter, ammeter, toggle switch and warning lights are located on the right hand part of the instrument panels and easy to read and to operate by the crew. Switch and warning lights are clearly labeled.

3.7.4 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n Voltmeter 831.120 UMA N141100917V0A0 Ammeter 831.121 UMA N15110003CA000 Shunt 831.122 UMA 1C1

3.7.5 OPERATION

Use toggle switch to choose indication of engine system voltage and current or main system voltage and current.

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3.7.6 DC POWER DISTRIBUTION SYSTEM ATA 246000

The system is designed to distribute electrical energy from the source to the electrical loads. Because the engine of the S12 is located in mid-fuselage, there are long distances between voltage sources and the loads.

Therefore, special attention is paid to subdivide the system into functional assemblies which includes sub-distributions with related protective devices.

3.7.6.1 WIRE SPECIFICATIONS

Cross Section Specication Battery and Starter Cables 16 mm² EN 50525-2-81 All other, unshielded and

shielded

AWG 24...8 MIL-22579/16 or

Raychem 44A

The cross section of each single wire is optimized individually.

The wires are bundled with spiral wrap to a wire harness. This harness is routed and xed along the aircraft’s steel tube frame from the electrical distribution rack to the instrument panel. Inside the cockpit, the wire bundle is routed along the tunnel between the seats.

Sucient distance to moving parts and heat sources is provided. Mechanical damage by tight bending or chang is prevented.

3.7.6.2 DESIGN OF ELECTRICAL CONNECTIONS

Electrical connections are made by screw terminals, push-on contacts or multiple position connectors. Terminals and contact pins are generally crimped to the wires. Use of soldering techniques is minimized. Terminals and connector systems meet industrial standards.

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3.7.6.3 PROTECTION DEVICES

Fuses and circuit breakers of dierent industrial standards are used to protect each electrical circuit. The protective devices are rated according to the maximum allowed electrical load of the protected wire. Protection devices are located as close as possible to the voltage source to minimize the length of unprotected wire.

Fuses are grouped in fuse holders at dierent locations at the aircraft. The fuse holders are not reachable by the pilot in ight. Devices and circuits protected by fuse are considered as non-essential to safe ight. The loss of the related function caused by a blown fuse will not cause hazards to aircraft or crew. On the other hand, because of the fuse ratings there must be a severe reason for a blown fuse. It is considered improbable, that this could be corrected by the pilot by replacing a fuse in ight. A sucient number of spare fuses is provided with the on-board tool set for a repair on ground.

The circuits of landing gear and electronic equipment are protected by resettable circuit breakers, located on the instrument panel and easy to reach by the crew in ight. This enables the pilot to evaluate the state of power supply of such devices in the case of malfunction. In addition, it allows the pilot to deactivate certain devices during gliding ight to save battery power (load shedding).

Each circuit breaker is clearly labeled with its function.

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3.7.6.4 FUSE AND CB SPECIFICATIONS: MAIN SYSTEM

Circuit Location of Fuse/CB Fuse /CB

STEMME p/n Qualication

Rating (A)

Main System Master Fuse

Main System Box 50 832.247 ISO 8820-5

External Alternator Main System Box 50 832.247 ISO 8820-5 Position Lights

5 Landing Headlight 5 ACL

(Anti-collision light) Landing

Headlight

Fuse holder Main System

Box

831.430 ISO 8820-3

Auxiliary Fuel Pump 5 Ammeter Hi 5 Ammeter Lo 5 COM

CB 5 XPDR CB 5 EFIS CB 5 GPS CB 5

Instrument Panel Front

832.241 IEC 60934

Autopilot CB 5 Landing Gear

System Trim System

CB 7.5 832.235 LN 29886

5 831.430 Alternator Field 5 831.430 Fire Warning 2 Alternator

Warning Light Soaring

Computer

Fuse holder Instrument

Panel

831.429

ISO 8820-3

Volt / Ammeter 2 Stall warning 2

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3.7.6.5 FUSE SPECIFICATIONS: ENGINE SYSTEM

Circuit Location of Fuse/CB Rating (A) STEMME p/n Qualication

Starter

Engine System Box System Master Fuse 50 831.247 Engine Bus Supply Internal Generator 15 Prop Pitch Control Power 15 Main Fuel Pump 5 TCU Power 5

Fuse holder Engine Fuel Transfer Pump RH 5

System Box

Fuel Transfer Pump LH 5 Ammeter Hi 5 Ammeter Lo 5 Internal Generator

warning Starter Relay

80 831.491

831.432

831.430

2 831.429

5 831.430

ISO 8820-5External Power Supply 80 831.491

ISO 8820-3

Propeller Pitch Indication 2

Low Fuel Warning Light 2 Tachometer 2 Ignition retarder 2 Fuel Pressure Indicator 2

Fuse holder Instrument

Panel

831.429

ISO 8820-3

Fuel Level Indicator Left 2 Fuel Level Indicator Right Oil Press + CHT Left 2 Oil Temp + CHT Right 2 Fuel Flow Meter Volt / Ammeter 2 Engine Hour Meter 2

Fuse holder Instrument

Panel

831.429 ISO 8820-3

Prop Pitch Sense 2

3.7.7 SWITCHING DEVICES

Power switching devices are located as close as possible to the voltage source. Due to the distances between voltage source and cockpit, the most electrical loads are remotely controlled. With the manually operated toggle and rotary switches on the instrument panel the pilot controls the negative (ground) lines of control circuits of electro-mechanical or semiconductor relays, which are located on the electric distribution board. The wiring of the load circuits is routed directly from distribution board to load device the voltage drop between power supply and load to minimize.

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3.7.7.1 MANUAL SWITCHES

Manual switches are located on the instrument panel and easy to reach by the crew. The position of a toggle or rotary switch is indicated by the position of the operating lever or knob. The use of momentary switches or push-buttons is limited to certain functions of electronic equipment, as such electrical elevator trim, the PTT button of the communication radio, autopilot disengage and the remote control unit of the soaring computer.

Each switching element is suciently rated to carry the electrical load of the related circuit. None of the electrical circuits controlled by manual switches on the instrument panel has an electrical consumption of more than 3 A. Each switch is clearly labeled with its function.

Function STEMME p/n Electrical Load Rating Master Switch 831.151 Relay Coil, < 1 A 3 A Auxiliary Fuel Pump ON 831.151 MOSFET Gate, < 0.1 A 3 A Engine System Back-Up 831.151 Relay Coil, < 1 A 3 A TCU Emergency 831.151 Relay Coil, < 1 A 3 A Transfer Pump Selector 831.507 Waste Gate Servo, < 2A 0.15 A Landing Gear UP/DOWN 831.225 MOSFET Gate, < 0.1 A 20 A Ignition-Starter-Key-Switch 831.107 Relay Coils, < 2 A Position Light Switch 831.493 MOSFET Gate, < 0.1 A 3 A ACL Switch 831.493 MOSFET Gate, < 0.1 A 3 A Landing Light 831.493 MOSFET Gate, < 0.1 A 3 A Propeller Pitch 831.151 MOSFET Gate, < 0.1 A 3 A

3.7.7.2 ELECTROMECHANICAL RELAYS

Power relays are used as master relays for the main system and the engine system, as starter relay and as contactor for the optional ground power supply. The relays are typically ground-controlled, the control switch is connected to the negative side of the relay coil. This improves the system reliability.

Each relay is rated suciently for its electrical load.

Function STEMME p/n Rating Fuse Qualication Main System Master 831.416 70 A 50 A Engine System Master 831.416 70 A 50 A Starter Relay 830.725 75 A cont. / 300 A

short time

Ground Power Relay 831.111 75 A 80 A

80 A Part of certied

engine

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3.7.7.3 SEMICONDUCTOR SWITCH BOX

An 8-channel MOSFET-based switch box is installed to control the power for fuel pumps, lights and variable pitch propeller. The use of semiconductors improves the system reliability, reduces electromagnetic emissions and saves weight and space. The switch box is located close to the batteries and to the distribution boxes.

The box includes 8 channels with a switching capacity of max. 36 A each.

Connector System Channel # Function Fuse

1 Position lights 5 A

Gray Main

Black Engine

2 Anti-Collision Lights 5 A

3 Landing Light 5 A

4 Auxiliary Fuel Pump 5 A

1 -spare2 Propeller Pitch Control 15 A 3 Transfer Pump Left Wing 5 A

4 Transfer Pump Right Wing 5 A

3.7.8 MAIN SYSTEM DISTRIBUTION BOX

4.7.8 Main System Distribution Box This box includes a metallic bus bar to which the main system battery and the alternator are connected via master fuses. The main system master relay and a number of fuses are located in a combined fuse-relay-holder. The box is made of plastic with a sealed cover and is mounted on a rack on the aft area of upper re wall.

3.7.9 ENGINE SYSTEM DISTRIBUTION BOX

This box includes a metallic bus bar to which the engine system battery, the starter, the generator and the external power supply are connected via master fuses. The engine system master relay and a number of fuses are located in a combined fuse-relay-holder. The box is made of plastic with a sealed cover and is mounted on a rack on the aft area of upper re wall.

3.7.10 TCU BOARD

The following elements of engine control are located on this board:

• The turbocharger control unit TCU.

• The waste gate servo.

• The ambient pressure sensor.

• The starter relay.

• The voltage regulator of the internal generator.

• The power relay for the ground power supply.

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The elements are screwed to a mounting plate, which is xed to the steel tube frame, aft of the engine compartment.

1 2 3 4

1. Batteries

2. Shunts

3. Engine System Box

4. Main System Box

Electrical Components Center Fuselage

Figure 3.7.a

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52 3 4

2. Shunts

3. Engine System Box

4. Main System Box

5. TCU Box

Figure 3.7.b

Electrical Components Center Fuselage

3.7.11 INSTRUMENT PANEL BUS SYSTEM

Two fuse holders, one for each system, are located behind the instrument panel. Each fuse holder forms a distribution bus. See chapter 7.3.1 and 7.3.2 for a list of the connected devices on each bus. From the main system bus a wire leads to the group of circuit breakers on the instrument panel front. These circuit breakers form the avionics bus. Radio communication and navigation equipment is connected here.

All engine monitoring instruments are connected to the engine system bus.

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1 2 3

1. Circuit Breakers for Avionics

2. Engine Bus

3. Main Bus

Figure 3.7.c

Electrical Components Instrument Panel

3.7.12 ENGINE STARTING SYSTEM ATA 800000

3.7.12.1 DESCRIPTION

The electrical starter is connected via starter relay to the engine system battery. The starter relay is controlled by the combined ignition-starter-key-switch located on the instrument panel. The control circuit for the relay triggers an ignition retarding module which allows the foldable propeller to unfold the blades completely before the engine res. This protects the blade folding mechanism from damages caused by engine starting torque.

3.7.12.2 PARTS SPECIFICATION

Starter and starter relay are part of the engine, their specications can be found in the ROTAX 914 Maintenance Manual.

3.7.12.3 OPERATION

The starter relay is triggered when the key switch is pushed to the “Start” position. The procedure to start the engine is described in the S12 Aircraft Flight Manual.

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3.7.13 ENGINE IGNITION SYSTEM ATA 740000

3.7.13.1 DESCRIPTION

The ROTAX 914 engine is equipped with a dual ignition. The ignition system works independently from the aircraft’s electrical system and energy sources. The ignition system is part of the engine. A detailed description can be found in the ROTAX 914 Maintenance Manual.

The ignition system is controlled by the combined ignition-starter-key-switch located on the instrument panel. It includes the magneto check function. The ignition switch is connected with two shielded AWG22 wires to the electronic ignition control boxes, one for each box.

When the switch is closed, the related control box is grounded and deactivated. An ignition retarding module is connected in parallel to the ignition control boxes. It is actuated by the key switch. When in “Start” position, the ignition retarder deactivates the ignition control boxes for approximately 3 seconds. This allows the foldable propeller to unfold the blades completely before the engine res. After the delay time the ignition is activated automatically.

The ignition system is not fused. The fuse of the ignition retarding module is located at the instrument panel fuse holder (engine bus).

3.7.13.2 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n Ignition-Starter-Key-Switch 831.107 ACS 11-03170 Ignition Retarder Module 124.211 n/a

3.7.13.3 OPERATION

• In “OFF” position the control inputs of both ignition boxes are connected to ground and the ignition is deactivated.

• In “LEFT” position the control input of the right side ignition box is connected to ground and the right ignition channel is deactivated.

• In “RIGHT” position the control input of the left side ignition box is connected to ground and the left ignition channel is deactivated.

• In “BOTH” position the control inputs of both ignition boxes are open and both ignitions are deactivated.

• In “START” position the control inputs of both ignition boxes are momentarily connected to ground and the ignition is inhibited for approximately 3 seconds.

The engine starting procedure and operation of the ignition-starter-key-switch is described in the S12 Aircraft Flight Manual.

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3.7.14 ENGINE RPM INDICATING SYSTEM ATA 771400

3.7.14.1 DESCRIPTION

The RPM signal is delivered by a pick-up coil at the rotating magnets of the internal generator to the tachometer input of the Turbocharger Control Unit (TCU). The TCU provides an output to which the electrical tachometer on the instrument panel is connected. The tachometer is wired with shielded AWG22 wire. The RPM signal is secondarily used to trigger a frequency relay which activates the eld circuit for the belt-driven alternator. The tachometer fuse is located at the instrument panel fuse holder (engine bus).

3.7.14.2 PARTS SPECIFICATION

Pick-up and tachometer are part of the engine, their specications can be found in the ROTAX 914 Maintenance Manual.

3.7.14.3 OPERATION

The engine RPM indicating system does not require operation actions by the pilot.

3.7.15 CYLINDER HEAD TEMPERATURE INDICATING SYSTEM ATA 772200

3.7.15.1 DESCRIPTION

The CHT indicating system measures the cooling water temperature on two of the four cylinder heads of the ROTAX 914 engine. The temperature signal is generated by resistive type probes (Pt-100 characteristics) and displayed to the pilot by one pointer-style gauge for each side of the engine. The indicators share their fuses with the oil temperature resp. pressure indicator. The fuses are located at the instrument panel fuse holder (engine bus).

3.7.15.2 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n Oil Temperature Indicator 831.241 UMA T12113R150C050 TSO Temperature probe RTD

(Resistance Temperature Detector RTD)

831.243 UMA 1B4R

3.7.15.3 OPERATION

The CHT Indication system does not require operation actions by the pilot. Scale markings of operational limits and instructions to the pilot about corrective actions in the case of exceeded limits are dened in the S12 AFM.

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3.7.16 OIL PRESSURE INDICATING SYSTEM ATA 793100

3.7.16.1 DESCRIPTION

The oil pressure indicating system measures the pressure at the output of the mechanical oil pump of the ROTAX 914 engine. A transducer generates a voltage of 0…5 V proportional to the pressure. This voltage is displayed to the pilot by a pointer-style gauge.

The indicator shares a fuse with the CHT-left indicator. The fuse is located at the instrument panel fuse holder (engine bus).

3.7.16.2 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n Oil Pressure Indicator 831.115 UMA T04113U010B030 TSO Pressure Sender 831.116 UMA T1EU150G

3.7.16.3 OPERATION

The Oil Pressure Indication system does not require operation actions by the pilot. Scale markings of operational limits and instructions to the pilot about corrective actions in the case of exceeded limits are dened in the AFM.

3.7.17 OIL TEMPERATURE INDICATING SYSTEM ATA 793300

3.7.17.1 DESCRIPTION

The oil temperature indicating system measures the oil temperature in the crankcase of the ROTAX 914 engine. The temperature signal is generated by a resistive type probe (Pt-100 characteristics) and displayed to the pilot by a pointer-style gauge. The indicator shares a fuse with the CHT-right indicator. The fuse is located at the instrument panel fuse holder (engine bus).

3.7.17.2 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n Oil Temperature Indicator 831.241 UMA T12113R150C050 TSO Temperature probe RTD

(Resistance Temperature Detector RTD)

831.243 UMA 1B4R

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3.7.17.3 OPERATION

The oil temperature indication system does not require operation actions by the pilot. Scale markings of operational limits and instructions to the pilot about corrective actions in the case of exceeded limits are dened in the AFM.

3.7.18 FUEL DISTRIBUTION SYSTEM ATA 282000

3.7.18.1 DESCRIPTION

The fuel is stored in two wing tanks of equal size, one in each side of the wing center section. Each tank is vented to the atmosphere to prevent overpressure or vacuum situations. Each tank has an over wing fuel ller. Each tank is connected via a coarse lter and quick drain to a dedicated electric plunger pump with downstream check valves. These check valves close o the fuel line when the pumps are not powered. The transfer pumps deliver the fuel into the feeder tank. The fuel line between check valves and feeder tank includes a quick-disconnect to facilitate handling during wing rigging and de-rigging. The feeder tank incorporates two oat switches to control the transfer pumps and to generate a low fuel warning signal to the pilot. From the feeder tank the fuel is delivered by a turbine pump (feeder pump) to the fuel pressure regulator, which provides the required fuel pressure to the carburetors. The regulator’s return line ends in the feeder tank. The feeder pump is backed up by another turbine pump of the same type.

3.7.18.2 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n

Main and Auxiliary Feeder Pump

11AB-M02 (830.209)

Pierburg E1F

Fuel Transfer Pumps 830.521 HKT HEP-02

3.7.18.3 OPERATION

a) Main Feeder Pump Control

The main boost pump is supplied by the engine’s electrical system. It automatically starts to work when the engine’s electrical system is activated by the pilot. Proper operation of the main boost pump is indicated by the fuel pressure indicating system (refer 28-44-00). The electrical power line of the main feeder pump is protected by a fuse located inside the engine system junction box (refer 24-60-00).

b) Auxiliary Feeder Pump Control

The auxiliary boost pump is controlled by a switch to be manually activated by the pilot when required. To achieve better redundancy of the fuel system, the auxiliary boost pump is powered by the main electrical system, independently from the engine’s electrical system. The proper function of the auxiliary boost pump is indicated by a green light and by the fuel pressure indicating system (ref 28-44-00). The electrical power line of the auxiliary boost pump is protected by a fuse located inside the main system junction box (refer 24-60-00).

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3.7.18.4 TRANSFER PUMP CONTROL AND TANK SELECTION

The transfer pumps are powered by the engine electrical system. They transfer pumps are controlled by a oat switch in the feeder tank and by a pump selector switch. At feeder tank fuel levels below 90% the transfer pump control is active. The pilot selects the pump on a rotary switch with the positions LEFT-BOTH-RIGHT. Due to the check valves downstream from the transfer pumps, the related tank outlet will be closed o completely when the associated transfer pump is not powered. Thus the pump selector switch functions as a tank selector.

The selected transfer pump/pumps stops automatically when the fuel level reaches 90% in the feeder tank. On feeder tank-fuel level below 90% the selected transfer pump/pumps start automatically.

The transfer pumps can be activated manually by a test push-button in the cockpit center console when the Master Switch is ON. This function is used for daily inspection or Maintenance Tasks. The transfer pump/pumps to be tested can be selected by using the fuel selector switch (Left, Right, Both).

3.7.19 FUEL QUANTITY INDICATION ATA 284000

3.7.19.1 DESCRIPTION

A oat type fuel level sender is installed in each fuel tank in the wing. The electrical resistance of the sender changes depending on the oat position. This resistance is measured and displayed to the pilot on one pointer-type fuel level indicator for each tank. The indicator scale is marked with lines, each representing one eighth of full.

A separate oat type switch is installed in the feeder tank which activates a yellow low fuel warning light on the instrument panel at feeder tank fuel levels below ca. 2.5 ltr. The fuses for the fuel level indicator system are located in the instrument panel fuse holder (engine bus).

3.7.19.2 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n Fuel Quantity Indicator 831.435 UMA T18112F1F4, TSO Fuel Level Sender 851.503 VDO 226 801 015 001 G, modied

3.7.19.3 OPERATION

The fuel quantity indication system does not require any operation actions by the pilot. Instructions to the pilot in case of low fuel level situation can be found in the AFM.

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3.7.20 FUEL PRESSURE INDICATING ATA 733200

3.7.20.1 DESCRIPTION

The ROTAX 914 engine requires the measurement of fuel pressure in relation to the “upper deck” air pressure of the turbocharger system. Therefore a dierential pressure sender is installed which generates a voltage output of 0 V … 5 V proportional to the dierence between regulated fuel pressure and air box pressure. This voltage is displayed to the pilot on a dedicated digital voltmeter in the engine portion of instrument panel. The operational range of 150 mbar … 350 mbar is marked on a placard close to the voltmeter.

The fuse for the fuel pressure indicator system is located in the instrument panel junction box (engine bus).

3.7.20.2 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n Dierential Pressure Indicator 831.123 UMA N1EU07D Indicator 851.147 Conrad 134473

3.7.20.3 OPERATION

If the fuel pressure exceeds the limits, the pilot has to check the available amount of fuel and the state of the tank selector and the auxiliary fuel pump. Instructions about appropriate corrective actions are included in the S12 Aircraft Flight Manual.

3.7.21 FIRE DETECTION SYSTEM ATA 261200

3.7.21.1 DESCRIPTION

Two bimetallic thermal switches are positioned on the upper rewall directly above the carburetors to detect increasing temperatures. At temperatures above 160°C the switches (or at least one) open the control circuit of a relay. This activates a red warning light and a piezo signal horn. The system includes a test switch on the panel.

3.7.21.2 PARTS SPECIFICATION

Part STEMME p/n Manufacturer p/n Thermoswitch Microtherm R20 10AM-F01 on request Switch with indication light - housing 10AM-F02 on request Switch with indication light - cap 10AM-F03 on request Switch with indication light - bulb 10AM-F04 on request

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3.8 COM AND NAV EQUIPMENT

COM and NAV equipment is installed in the instrument panel. The loudspeaker is installed on the rear wall of the cockpit above the LH baggage compartment. The gooseneck microphone is fastened between the backrests at the center console. When using headsets it can be deselected.

Antennas are installed for:

• VHF-COM in rudder,

• VOR-NAV on cockpit oor. (Kevlar shell, optional equipment, retrot is not possible.)

• Transponder antenna in the front part of the tail cone or on nose-cone.

Only equipment listed in section 9 of this Maintenance Manual and approved for installation in the S12 may be installed without additional certication. Only this equipment has been proven by the manufacturer in accordance with the relevant certication requirements regarding initial airworthiness and proper function. For modications, the installation instruction and the original wiring harness of the aircraft manufacturer must be used.

This is also relevant for equipment approved for operation in motorglider or equipment without certication requirements due to possible inuence on power consumption, electromagnetic inuence and structural characteristics of the instrument panel. Outside Germany it must be conrmed that the equipment is certied by the local authority in charge.

The weight limits for the equipment installed in the instrument panel and the inuence on CG must be respected. The Equipment List and the Weight and Balance Report must be updated.

When equipment is modied, conformity with the type must be checked and conrmed by an inspector. If equipment is installed which is not certied for use in the relevant aircraft it must be proven by the authority, when applicable requirements are fullled (modication of a single aircraft).

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3.9 OXYGEN SYSTEM

One or maximum two oxygen system mountings (optional equipment) for one oxygen bottle each are installed in the upper baggage compartment.

The mountings are suitable for oxygen bottles from various manufacturers, provided the diameter is within a minimum of 0.132 m / 5.20 in, and the total length including regulator is approx.

0.45 m / 17.72 in up to a maximum of 0.52 m / 20.47 in.

The certication of the motorglider does not include a dened oxygen system and fulllment of the requirements must be demonstrated to the authority by the supplier or the facility, which modied the aircraft (normally as a modication of a single aircraft).

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3.10 WING FOLDING DEVICE

The Stemme S12 optionally can be equipped with a wing folding device.

A description of the functioning and the handling is to be found in the associated supplement to the S12 Aircraft Flight Manual.

The wing folding device is maintenance-free.

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4.1 GENERAL

This chapter denes the mandatory overhaul (TBO – Time Between Overhaul), replacement (TBR – Time Between Replacement), and inspection intervals for components of the STEMME S12. The chapter also denes the structural inspection requirements for the STEMME S12 composite airframe. It contains information about the lifetime of all STEMME designed parts or components that must be overhauled, rebuilt, inspected, or replaced at either specic ight time limits or specied calendar time limits or that require monitoring through scheduled maintenance.

Compliance with the specied times and intervals is mandatory for maintaining continuing airworthiness of the aircraft.

For airworthiness limitations of other parts installed in the aircraft not designed by STEMME or with separate TC refer to chapter 5 and respective manufacturer documentation as amended from time to time.

4.2 MAINTENANCE LIMITATIONS

Be sure to verify compliance with chapter 4 Airworthiness Limitations when performing scheduled maintenance per chapter 5. The following requirements must be adhered to consistent with applicable national regulations.

4.2.1 PAINT FINISH

To ensure that the temperature of the composite structure is kept below 54°C / 130 °F, the airframe must be painted white as approved. Only the nose cone may be a darker color. On the rest of the airframe, you may only paint registration letters and numbers and placards in a color other than white.

When re-painting is needed on a cowling, you must paint the engine side (interior side) of the lower and side cowlings with aviation grade re-resistant paint approved by STEMME.

Refer to P520-901.502 “Small Repairs of Composite Materials” (latest approved issue), Annex A of this manual, for guidance on the paint approved for use by STEMME.

CAUTION

Aircraft control surface balance is critical to ight safety. Repair and removal or addition of any paint or body ller to a control surface requires that the control surface is rebalanced according Annex D of this manual.

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4.3 COMPONENT REPLACEMENT AND OVERHAUL LIMITATIONS

This chapter denes replacement intervals and maintenance requirements for aircraft components, systems, and structures determined to be life-limited or to be monitored by scheduled maintenance.

Unless otherwise specied, the following components must be overhauled, rebuilt, or replaced with components that have service life remaining, within the life-time limits specied. To monitor remaining useful service life-time of components, maintenance work, e.g., overhauling, rebuilding, repairing, or replacing each item, must be recorded in accordance with applicable national regulations.

The IPC chapters mentioned below refer to the Illustrated Parts Catalog (IPC) for Type S10, Model S12 Doc.-No- L530-912.830 or latest revision.

MAXIMUM ALLOWABLE OPERATING TIMES

No. Part/Assembly/

Equipment

1 Variable pitch

Manufacturer STEMME

p/n

IPC

Chapter

Overhaul Replacement Notes

STEMME AG 11AP-V 61-10 5 years - (1)

propeller 2 Propeller hub STEMME AG 10AP-V01 61-10 - 2000 hr 3

Propeller fork

and fasteners/

attachment

components

4 Propeller gear

STEMME AG 10AP-V88,

61-10 200 hr 400 hr (1) 10AP-V76, 10AP-V77, 10AP-V78, 10AP-VU

STEMME AG 11AG 72-10 1000 hr - (1)

box (gear

wheels)

5 Propeller gear

STEMME AG 11AA 72-10 1000 hr - (1)

box mounting

6 Propeller Gear

Box mounting

STEMME AG 11AA-S

11AA-36

72-10 - 12 years (1)

elastomer parts

7 Centrifugal

STEMME AG 12AK 72-10 400 hr - (1)

clutch

(3 Flyweights)

Drive shaft

STEMME AG 10AS-07

10AS-W

72-10 800 hr - (1)

10AS-F

9 Flexible disk of

STEMME AG 11AS-09 72-10 - 12 years (1)

the drive

shaft system

(Cardan Joints)

(2)

(3)

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NOTES

1. Overhauls of STEMME–manufactured components may only be performed in accordance with manufacturer–approved data.

Under extreme environmental conditions (hot, dusty) greasing of the propeller blade bearings as part of the overhaul can be required more frequently if indicated during Annual Inspection.

Dierent Overhaul and Replacement intervals in the variable pitch propeller assy. are, dened for specic individual mechanical components. Respective components are sorted by their applicable TBO/TBR.

If the limitation is given in engine operating hours and in a calendar period, the rst occurring case applies for the aected components.

Even under ideal conditions, replacement or overhaul of parts or components may be required before attaining the chapter 4 Life Limits. Even life–limited parts are subject to “on–condition” inspection before reaching TBO intervals.

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