UL 2 - Part I.

Airworthiness requirements SFD Aerodynamically controlled ultralight aircraft

Text as of: 26. 03. 2019

CHANGE SHEET

Date of issue of the change	Edited/deleted/new paragraphs:	Date of inclusion	Ranked

Content

DEFINITIONS, ABBREVIATIONS AND LABELS
TITLE A - GENERAL
SUBPART B - FLIGHT PERFORMANCE and CHARACTERISTICS
TITLE C - STRENGTH
TITLE D - DESIGN and CONSTRUCTION
TITLE E - PROPULSION SYSTEM
TITLE F - EQUIPMENT
TITLE G - OPERATING LIMITATIONS AND DATA
SUBPART J - PROPELLERS
APPENDIX I. - RESCUE SYSTEMS
APPENDIX II. – TOWING GLIDES
APPENDIX III. – UNIT LOADS OF THE REAR PART OF THE AIRCRAFT
APPENDIX IV. – BASIC LANDING CASES

 

DEFINITIONS, ABBREVIATIONS AND LABELS

I. General definitions

Tíha G = m∙g [N]

where: m mass [kg]
g gravitational acceleration [g = 9,81 m/s]

The International Standard Atmosphere (MSA) is defined as follows:

1. air is a perfect dry gas

2. the temperature at the height H = 0 m is 15 °C

3. air pressure at height H = 0 m is 1013,25 hPa

4. the temperature gradient from zero height to a height where the temperature reaches −56 °C is −0,0065 °C/m

5. the air density ρ under the given conditions is 1,225 kg/m³.

II. Definition of speeds

VA	Design speed of turns (maneuvering)
VB	Design speed for maximum gust size
VD	Steep descent design speed
VDF	Highest speed demonstrated by flight test
VF	Design speed with flaps extended
VFE	Maximum allowable speed for use of flaps
VH	Maximum speed in level flight at maximum permissible sustained power
VLO	Maximum allowable speed for extending the landing gear
VNE	Maximum unexceeded speed
VRA	Maximum speed in strong turbulence
VS	Stall speed or the lowest steady speed at which the airplane is still controllable
VSat	Stall speed or the lowest steady speed at which the airplane is still controllable in the landing configuration
VS1	Stall speed or the lowest steady speed at which the airplane is still controllable in the specified configuration
VSF	Calculated stall speed with fully extended flaps at maximum takeoff weight
VT	Maximum permissible speed for towing
VY	Best climbing speed
CAS	Calibrated airspeed (CAS = calibrated airspeed). Indicated airspeed corrected for instrument and installation error
EAS	Equivalent airspeed (EAS = equivalent airspeed). Calibrated airspeed, adjusted for adiabatic compressibility at given altitude. For MSA at sea level, the equivalent airspeed is equal to the calibrated airspeed.
IAS	Indicated airspeed (IAS = indicated airspeed) is the airspeed as indicated by the speedometer (connected to the pitotstatic system), calibrated with respect to the adiabatic compression of the air stream in a standard atmosphere at sea level, uncorrected for the errors of the airspeed system.
CAS	True airspeed (TAS = True airspeed). Flight speed in still air. Actual airspeed is the equivalent airspeed multiplied by (ró/ró0)½.

III. Definition from the field of strength

Supporting structure	There are those parts of the structure of an ultralight aircraft, the failure of which would seriously endanger the safety of the aircraft.
Maximum take-off weight MTOM	The largest weight at which an ultralight meets airworthiness directives.
Weight of empty aircraft	It is stipulated in UL 2 § 29.
Operating load	The maximum load that can be expected in operation.
Numerical load	The service load multiplied by the appropriate factor of safety, normally 1,5.
Operating multiple n	The ratio of the total aerodynamic force acting perpendicular to the flight path and the total weight of the aircraft. in straight-line steady flight this multiple is equal to one. Operating loads can be expressed in terms of aerodynamic forces or acceleration forces.

IV. General technical terms

UL	Ultralight aircraft.
Motor	A propulsion engine that is used to propel an airplane.
Buildable propeller	Designation of a propeller whose settings can be changed at rest or while running. It is divided into: a) propellers whose settings are directly controlled by the pilot (manually adjustable propellers). b) propellers whose settings are controlled by a regulator or other automatic device. This device can be either a fixed part of the propeller or a separate device that is or is not controlled by the pilot (constant speed propeller). c) propellers whose settings are changed by a combination of the methods listed in points a) and b).
Seat belts	Safety belts within the meaning of this regulation are four-point belts, consisting of a shoulder belt for each shoulder and two lap belts.
Multiple	The ratio of the specified load to the total weight of the aircraft. The determined load can be air forces, inertial forces, or reaction forces from the ground or water.
Maximum take-off power	Designation of power limited to a duration of 5 minutes, which is permissible at sea level under MSA conditions and at the maximum permissible speed and inflation pressure for take-off, missed approach and continuous take-off.
Maximum continuous performance	Designation of the power that is permissible for an unlimited time at a specified altitude under MSA conditions at maximum speed and boost pressure.
Incandescent	In a fire area marked "flame resistant", there is the ability of parts of the structure and equipment to withstand the thermal action of a "standard flame" for 15 minutes. For materials and building components that serve to limit fire in a given fire space, the designation "refractory" means the ability of the material, to the extent that corresponds to the purpose of use, to resist the temperature arising in a long continuous fire of the highest intensity in each zone at least as well as steel. For pipes and other parts of equipment, the designation "refractory" means the ability of the material, to the extent appropriate for the purpose of use, to resist the heat generated in a fire at least as well as steel.
Refractory	In a fire area marked "fire resistant", there is the ability of parts of the structure and equipment to withstand the thermal action of a "standard flame" for 5 minutes. For sheets and structural building parts, the designation "refractory" means the ability of the material, to the extent that corresponds to the purpose of use, to resist the temperature generated in a fire at least as well as Al alloy and in pipes for liquids, other parts of systems with flammable liquids, electrical lines, stops and engine controls the ability to perform their respective functions at the temperatures and extraordinary conditions that arise during a fire in a given location.
Hard-tempered	A material that does not support combustion in such a way that the flame spreads indefinitely after moving the material away from the burning place is defined as non-flammable.

V. Symbolism Used

B	surface depth [m]
CY	lift coefficient [-]
P	power, in general [N]
S	area, generally [m2]
Terms & Conditions	horizontal tail surfaces
SOP	vertical tail surfaces
STerms & Conditions	VOP area [m2]
SSOP	SOP area [m2]
W-	average surface operating load [Pa]
W	size of specific local load on the surface [Pa]
D	proportional deformation [%]
G	G = MTOM*g [N] weight of the aircraft at maximum take-off weight

 

TITLE A - GENERAL

1. Purpose

This building code sets the minimum airworthiness requirements for aerodynamically controlled ultralight aircraft, which need to be met so that the use of the UL aircraft for the specified purpose is problem-free and does not jeopardize the safety of air traffic as well as the safety of third parties.

UL 2 § 1 Applicability

These airworthiness requirements are intended for aerodynamically controlled UL airplanes,

1. whose take-off weight does not exceed 600 kg, including the rescue system and

2. whose minimum speed V_SO according to UL 2 § 49, it is not higher than 83 km/h (CAS).

UL airplanes are only authorized for non-acrobatic operations, which include:

1. any turn required for normal flight,

2. flight modes with a deviation of max. ±30° from the horizon around the transverse axis, including fall prevention training,

3. sharp turns with an inclination of up to 60°.

 

SUBPART B - FLIGHT PERFORMANCE and CHARACTERISTICS

I. In general

UL 2 § 21 Keeping a license

Each requirement of this Chapter must be proven by a type airplane test, for the most unfavorable combinations of weight and position of the center of gravity in their entire range.

A certificate must be made for all configurations in which the airplane will be operated, unless otherwise specified.

Warning:

Subpart B does not list all the flight tests that are required to meet the certification requirements. The responsible authority has the right to determine the scope of the flight tests.

Notes on point UL 2 § 21:

1. Test flight instrumentation

a. For the tests, the aircraft should be equipped with suitable devices that allow the necessary measurements and observations to be carried out in a simple way.
b. In the initial phase of the test program, the accuracy of the instruments and their correction curves must be determined. Particular attention should be paid to erroneous data of the speedometer system, while the relevant configuration of the aircraft must be considered.

2. The following ground tests must be performed prior to flight tests:

a. Engine test run.
b. Measurement of maximum deflections of control surfaces, ailerons, flaps and their control elements.
c. Weighing the aircraft and determining the operational positions of the center of gravity.

3. Functional tests

All ground functional tests must be performed before flight testing begins.

UL 2 § 23 Limitation of load distribution

1. The ranges of weights and positions of the center of gravity in which the safe operation of the aircraft is to be ensured must be determined by the applicant.

2. The range of center of gravity positions shall not be less than that corresponding to the weight of each crew member in the range from the minimum weight of 70 kg for the pilot alone to the specified maximum weight for the pilot and passenger, always taking into account the most unfavorable distribution of fuel and luggage. The specified maximum weight of a person must not be less than 110 kg for a single-seater aircraft and 2x100 kg for a two-seater aircraft (see Notice in UL 2 § 25).

UL 2 § 25 Weight limitation - maximum weight

The maximum weight must be determined so that

1. not to be higher than:

a. the highest weight proposed by the applicant,
b. design maximum weight, which is the highest weight for which a certificate is maintained considering all load cases and all flight performance requirements, and
c. was not less than the weight consisting of the weight of the empty airplane with the minimum required equipment, the minimum crew member weight of 110 kg for a single-seater airplane or the minimum crew member weight of 200 kg for a two-seater airplane and the fuel supply per hour of flight at maximum continuous engine power.

Warning:

1. The weight of the crew member must not be less than 100 kg for the strength certificate.

2. The maximum amount of fuel and any additional equipment must be considered (pay attention to the increase in weight during equipment changes, repairs, etc.)

UL 2 § 29 Weight of the empty aircraft and the corresponding position of the center of gravity

1. The weight of the empty airplane and the corresponding position of the center of gravity must be determined by weighing the airplane with:

a. a firmly built-in load,
b. minimum required equipment,
c. unusable amount of fuel,
d. and in case, if it is used, with the maximum amount of oil,
e. and in the case, if used, with hydraulic fluid,
f. and if used, with engine coolant,
g. and with the undercarriage in the retracted position, if retracting the undercarriage affects the change in the position of the center of gravity,

but without

h. weight of crew member(s), a
i. the weight of other easily removable parts of the load.

2. The configuration of the aircraft when determining the weight of the empty aircraft must be precisely defined and easily retrievable at any time.

3. The weight report must be accompanied by an equipment overview that describes the equipment (eg tire size, wheel covers, fuel volumes, etc.) and an equipment overview that lists all fixed equipment elements (eg avionics, auxiliary heating) , towing equipment if fitted, etc.). The position of the center of gravity of the empty airplane and the range of operating positions of the center of gravity for the extreme front and rearmost position of the center of gravity must be determined for the weight report.

II. Flight performance

UL 2 § 45 In general

Proof of compliance with the requirements of this Chapter for flight performance must be made for the maximum weight and refer to no wind and normal atmospheric conditions at zero altitude of the international standard atmosphere (hereafter referred to as MSA).

UL 2 § 49 Falling speed

1 V_SO is the stall speed (CAS) if achievable in flight, or the minimum steady speed at which the airplane is still controllable with the engine idling (inlet closed) or shut down. The configuration that shows the larger value of V is decisive_SO, and

a. the airplane is in landing configuration a
b. the weight corresponds to the maximum weight and the position of the center of gravity is in the least favorable position of the permitted range.

2 V_S1 is the stall speed (CAS), if attainable in flight, or the lowest steady speed at which the airplane is still controllable, at which the engine is idling (inlet closed) or shut down, while

a. the airplane is in the configuration it maintains during the test in which speed V will be used_S1 a
b. the weight corresponds to the maximum weight and the position of the center of gravity is in the least favorable position of the permitted range.

3 V_SO and V_S1 must be determined by flight tests according to the requirements set forth in point UL 2 § 201.

UL 2 § 51 Take-off

The take-off length must be determined for a take-off from rest to a height of 15 m from dry, flat and short-cut grass at maximum weight and in no wind. It can be a maximum of 450 m.

Note:
The take-off length given in the Flight Manual is to be determined as the mean value of six proof flights.

UL 2 § 65 Climbing

The best rate of climb shall be, after correction for MSA zero altitude conditions, with:

a. maximum take-off power,
b. with the undercarriage retracted,
c. maximum take-off weight, a
d. flaps in the position prescribed for climb and without exceeding any specified temperature limitation, higher than 1,5 m/s.

III. Handling and agility

UL 2 § 143 In general

1. The aircraft must be safely controllable and capable of maneuvering when:

a. take-off with maximum take-off power,
b. climb,
c. level flight,
d. descent,
e. landing with engine running and engine off, a
f. sudden stopping of the engine.

2. The aircraft must be able to perform a smooth transition from one flight position to another under all probable flight conditions (including turns, if they are possible based on the configuration) without extraordinary demands on the pilot's skill, alertness and strength of the pilot and without the risk of exceeding the operating multiples at each admissible change in engine performance or its sudden stoppage. Slight deviations from recommended procedures must not lead to an unsafe flight situation.

3. Any abnormal flight performance observed during the flight tests required to demonstrate compliance with the flight performance requirements and any significant change in flight performance due to rain must be verified for all permissible engine operating modes.

4. If the required forces of the pilot seem uncomfortably high, compliance with the limit values ​​of the pilot's forces must be proven by quantitative tests. The forces from the pilot must in no case exceed the maximum values ​​for normal steering around the three axes listed in the following table. These requirements must be met for all permissible engine operating modes.

	altitude control	lateral steering	directional control	flaps, landing gear
	[tax]	[tax]	[tax]	[tax]
a) short-term activity	20	10	40	10
b) longer-term activity	2	1,5	10

5. The deflections of the control surfaces and auxiliary rudders available to the pilot must not, under any circumstances, be reduced to such an extent by the elastic extension of the control system that the UL aircraft would be difficult to control.

Note:
When steering by changing the position of the center of gravity and other non-conventional methods of steering, during short-term effects on the control elements (e.g. the trapeze), an average physically fit pilot must be able to overcome the steering forces, and during long-term actions, the pilot's steering forces must not overload the pilot. In these cases, it is necessary to discuss these matters with the responsible authority.

UL 2 § 145 Height control

1. At every speed lower than 1,3V_S1 it must be possible to change the pitch of the longitudinal axis by the action of the altitude control so that the aircraft quickly regains a speed of 1,3 V_S1.

Test conditions: All possible aircraft configurations and engine modes, with the aircraft balanced at 1,3V_S1 (if balancing is taken into account).

2. It must be possible to change the configuration (undercarriage, flaps, engine operating modes, etc.) in the entire range of the turn envelope, without requiring special skill of the pilot and without exceeding the specified steering forces according to point UL 2 § 143.

3. At speed V_DF it must be possible to safely remove the airplane from steep flight for all permissible center of gravity positions and engine operating modes.

UL 2 § 147 Transverse and directional control

With adequate steering action, it must be possible to go from a 30° banked turn to a turn in the opposite direction within 5 seconds. The turn must be performed at a speed of 1,3 V_S1 and where applicable with the landing gear extended and the flaps extended.

UL 2 § 155 Force from altitude control during manoeuvre

The airplane must demonstrate such elevator forces that increase in a turn or when taking a turn with steady speed in proportion to the multiple at all speeds at which the required normal acceleration can be achieved without stalling, simultaneously with the flaps deployed and where is used, even with the undercarriage retracted.

For aircraft controlled by changing the position of the center of gravity and other non-conventional control methods, the control force required to derive the operating load must be discussed with the responsible authority.

UL 2 § 161 Balancing

The speeds to achieve a balanced equilibrium state around all three axes must lie between 1,3V_S1 and 2,0V_S1 for all engine operating modes and extreme center of gravity positions.

IV. Stability

UL 2 § 171 In general

The aircraft must meet the requirements of UL 2 § 173 through UL 2 § 181. In addition, it must have sufficient stability and "handling feel" (handling smoothness) under all normal expected operating conditions.

UL 2 § 173 Longitudinal static stability

For each combination of center of gravity position, flap position, and engine power, the slope of the force-to-steering versus speed curve must be positive over the entire speed range from minimum speed to maximum allowable speed. Any significant change in speed must cause a change in steering force that is significantly felt by the pilot.

UL 2 § 177 Transverse and directional stability

1. If the airplane is in steady straight flight, then every increase in deflection of the transverse rudder (ailerons) must correspond to an increase in the yaw angle, if the transverse and directional controls are gradually deflected in the opposite direction. This behavior may not be directly proportional.

2. During a yaw, no change in steering forces must be so great that the control of the airplane requires special skill of the pilot.

UL 2 § 181 Dynamic stability

All the rapid oscillations that occur between the stall speed and the V speed_DF s:

a. free, a
b. firm management

 they must be heavily damped.

These requirements must be met in all permissible engine operating modes.

V. Dragging

UL 2 § 201 Behavior when stalling in straight flight

Dragging properties must be tested for the maximum forward and maximum rear position of the center of gravity and for the maximum and minimum weight specified in UL point 2 § 25.

Drag behavior tests shall be carried out as follows:

1. The initial level flight speed is reduced by approximately 2 km/h per second either until a stall condition is reached, which is manifested by an uncontrollable nose-down pitch or a nose-down pitch and simultaneous pitching to one wing, or until after landing the elevator on stop. Until the stall condition is reached, it must be possible to induce and correct yaw and yaw by steering action, which means corresponding rudder deflections.

2. When resuming the normal flight attitude, it must be possible to prevent a bank of more than 20° during normal use of the controls. The aircraft must not show any uncontrollable tendencies to go into a spin.

3. The loss of altitude from the start of the stall to the resumption of level flight must be determined using normal procedures.

Note:
Stall loss is the difference between the altitude at which the stall occurred and the altitude at which level flight was regained.

4. Proof of compliance with the requirements of paragraphs 1 to 3 of this point must be made under the following conditions:

a. flaps in all positions,
b. chassis retracted and extended,
c. airplane balanced at 1,4V_S1 (if balancing is considered), a
d. engine power

idling,
75% of maximum continuous power, a
if 75% of max. continuous power causes a longitudinal inclination angle greater than 30°, then the power setting can be reduced to a maximum of 50% of max. continuous power.

UL 2 § 203 Dragging in a corner

1. When stalling in a cleanly flown 30° pitch turn, it must be possible to resume normal level flight without the aircraft exhibiting a tendency to pitch uncontrollably or to spin into an uncontrollable spin.

Note:
A bank will be rated as uncontrollable if the aircraft banks more than an additional 30° in the direction of the turn.

2. The height loss from the start of the stall to the resumption of level flight must be determined using normal procedures. This requirement must be met under the conditions prescribed in point UL 2 § 201, paragraph 4.a. to 4.d.

UL 2 § 207 Warning against drag

1. The aircraft does not need to give a stall warning if, during a stall from direct flight:

a. it is possible to induce and correct yaw by lateral control, while the directional control is held in the neutral position,
b. there is no significant drop in the bearing surface, with the rudder and yaw controls held in neutral.

2. An aircraft that does not meet the conditions under paragraph 1.:

a. must clearly and distinctly warn of stalling, both in straight flight and in a turn, with flaps and landing gear in any normal position,
b. must not give a stall warning at normal operating speeds, but the warning must occur sufficiently early before the stall configuration is reached for the pilot to return the airplane to level flight, and
c. stall warning can be given either by inherent aerodynamic properties (eg shake) or by a device that clearly signals stall.

VI. Behavior on the ground

UL 2 § 233 Directional stability and controllability

At any speed that can be expected when the aircraft is moving on the ground, uncontrollable behavior must not occur and the aircraft must have sufficient directional control during taxiing.

UL 2 § 234 Take-off and landing in crosswinds

The ability of the airplane to take off and land safely in crosswinds must be tested. Based on the results of these tests, the conditions for crosswind operation are determined in the Flight Manual.

VII. Special requirements for operating conditions

UL 2 § 251 Vibration and flutter

At all speeds up to V_DF excessive vibrations must not occur on any part of the aircraft. In addition, under no normal flight conditions should there be such violent flutter as to interfere with satisfactory control of the airplane, cause excessive fatigue to the crew, or damage the structure. A flutter that warns of overstretching is acceptable within the specified limits. This requirement must be met with the engine stopped and running in all permissible engine operating modes.

 

TITLE C - STRENGTH

I. In general

UL 2 § 301 Load

1. The strength requirements are determined in the form of operating load (the highest expected load during operation) and numerical load (operating load multiplied by the prescribed safety factors). Unless otherwise specified, the loads specified in the regulation are operational.

2. Unless otherwise specified, air and ground loads are always balanced with inertial forces, taking into account all larger isolated weights of the aircraft. Loads must be distributed in such a way that these distributions must either correspond to actual conditions or approach them on the safe side.

3. If the distribution of external load and internal forces will be substantially changed due to deformations under load, such a new distribution must be considered.

UL 2 § 303 Safety factor

1. Unless otherwise specified, a safety factor of 1,5 must be used.

2. The safety factor must be multiplied by an additional safety factor if:

a. there is uncertainty about the strength of the component (part),
b. loss of strength must be expected in time until replacement,
c. exact strength values ​​are not available due to unknown manufacturing and testing methods.

The size of this additional safety factor, unless otherwise stated below, must be determined for each type separately. The required time until the replacement of these component(s) is specified in the Operation Technical Manual.

d. the additional safety factor is mainly determined for:

i. every part that has clearance (does not apply to pressing) and is subject to impact stress or vibration,
ii. rudder hinges (except rolling and joint bearings),
iii. bearings (joints) in drawbar steering that are subject to angular movement (except rolling and joint bearings), and
iv. bearings (joints) in cable management.

Application	Additional safety factor	The resulting factor of safety fU
joints (shear bearing) with clearance, loaded by impact stress or vibrations	2,0	fU = 2,0 * 1,5 = 3,0
rudder hinges (except rolling and joint bearings)	4,44	fU = 6,67
bearings (joints) in the tie rod control	2,2	fU = 3,30
bearings (joints) in cable management	1,33	fU = 2,0
castings	2,0	fU = 1,5 * 2,0 = 3,0
fittings – applies to – all fittings, – all fasteners, – pressing (for fixed mounting)	1,15	fU = 1,5 * 1,15 = 1,725
seat belts and seats	1,33	fU = 1,5 * 1,33 = 2,0

The increasing coefficients are applied in the case of a strength certificate carried out numerically (not by test), with the exception of safety belts and seats.

Interpretation of the use of additional factors for composite structures:
The safety factor f for composite structures is in the range of 1,5 to 2,25, that is, an increasing factor of 1 to 1,5.

The use of the increasing factor depends on:

a. component or part to which it will be applied,
b. accuracy of calculations and their reliability,
c. submitted tests of composite material samples and their results, a
d. verification of production, control procedures and experience of the manufacturer.

The Technical Commission will decide on the use of the appropriate safety factor in cooperation with the chief engineer and the expert opponent of the project. In the case of amateur constructions, the chief technician and the construction supervision technician will decide on the use of the coefficient.

UL 2 § 305 Strength and deformation

1. The structure must be able to transfer operational loads without permanent deformations. Under all loads up to the operating load, the resulting deformations must not limit safe operation. This applies primarily to control systems.

2. The structure must be able to carry the numerical load for at least 3 seconds without failure. However, the three-second limit does not apply if the strength test is carried out by a dynamic test, during which real load conditions are demonstrated.

UL 2 § 307 Evidence of structural strength

1. Fulfillment of the strength and deformation conditions according to point UL 2 § 305 must be proven for all critical load cases. A theoretical, numerical certificate can only be recognized if the chosen type of construction is known based on experience that the calculation methods used give reliable results. Otherwise, a strength test must be carried out.

2. Certain parts of the structure must be proved as specified in Chapter D of this regulation.

Warning:
Subpart C does not list all the strength requirements for the certificate.

II. Flight loads

UL 2 § 321 In general

1. Flight multiples are given by the ratio of the air force component that acts perpendicular to the plane's flight path to the plane's weight. With a positive multiple, the air force is oriented upwards relative to the aircraft.

2. Proof of compliance with flight load requirements must be made for all possible weight and centering combinations.

UL 2 § 331 Symmetrical flight conditions

1. When determining wing loads and gravity and inertial loads under symmetrical flight conditions according to UL 2 § 333 to UL 2 § 345, the relevant balancing loads on the horizontal tail surfaces must be considered to match actual conditions or be on the safe side .

2. The increase in load on the horizontal tail surfaces during turns (when the rudders are in operation) must be balanced by the forces of the angular rotational acceleration of the aircraft in such a way that it corresponds to the actual conditions or is on the safe side.

3. When determining the load (turnover ratio) that occurs under the prescribed conditions, it is assumed that it is caused by a sudden change in the angle of attack while maintaining the flight speed. Angular accelerations need not be taken into account.

4. The aerodynamic values ​​that are required to determine the load conditions must be supported by tests, calculations or safe estimates.

a. If more accurate data is not available, the largest negative lift coefficient value for rigid airfoils in normal configuration of −0,8 may be used. In the case of non-rigid load-bearing surfaces, this must be consulted with the responsible authority.
b. If the overturning moment coefficient is C_mo less than ±0,025, the C factor must be used for the wing and tail surfaces_mo with a value of at least ±0,025.

III. Flight envelope of operating multiples (Vn diagram)

UL 2 § 333 In general

1. Compliance with the structural strength requirements must be demonstrated for all combinations of flight speeds and load multiples located on the boundary curve and inside the load envelopes described in paragraphs 2 and 3 of this section.

2. Reversal envelope (see Fig. 1)

Configuration:
Flaps in flight position (see Fig. 1).

3. Gust envelope (see Fig. 2)

Configuration:
Flaps in flight position (see Fig. 2).

a. At design speed V_B the UL aircraft must be able to withstand a positive gust (up) and a negative gust (down) up to 15 m/s, which acts perpendicular to the flight path.
b. At design speed V_D the UL aircraft must be able to withstand a positive gust (up) and a negative gust (down) up to 7,5 m/s, which acts perpendicular to the flight path.

UL 2 § 335 Design airspeeds

The following design airspeeds are equivalent airspeeds (EAS).

1. Design maneuvering speed V_A

where:
V_S1  = specified design stall speed at maximum design weight, flaps retracted and engine idling.

2. Design airspeed with flaps extended V_F

There must be no V in all landing configurations_F less than the greater of both of the following values:

a. 1,4 V_S1, where V_S1 is the calculated stall speed with the flaps retracted and at maximum weight,
b. 1,8 V_SF, where V_SF is the calculated stall speed with the flaps fully extended and at maximum weight.

3. Maximum design speed V_D

The maximum design speed can be chosen by the designer, but it must not be less than the greater of the following values:

a. 1,2 V_H, where V_H is the maximum level flight speed at maximum sustained engine power,
b. 1,5 V_A according to paragraph 1.

4. Design speed in strong gust V_B

The maximum design speed in a strong gust may be chosen by the designer, but:

a. must not be less than V_A ,
b. need not be greater than 0,9V_H, where V_H (EAS) is the maximum level flight speed at maximum sustained engine power.

UL 2 § 337 Operating turnover multiples

Operating turnover multiples according to the turnover envelope (see Fig. 1) must have at least the following values:

n1	+ 4,0
n2	+ 4,0
n3	- 1,5
n4	- 2,0

Negative operating multiples of turns for UL airplanes with non-rigid airfoils that have only limited ability to sustain negative acceleration in flight must be consulted with the responsible authority.

Deformation of non-rigid airfoils can lead to significant changes in the use of the turn envelope such that point A is not reachable below speed V_D. If such cases are proven, the operating multiplier may be reduced to the highest achievable multiplier below V speed_D.

UL 2 § 341 Gust multiples of turns

If a more accurate calculation corresponding to the actual conditions is not available, the gust multiples must be calculated as follows:

where:
U = gust speed [m/s]
V = flight speed [m/s]
a = slope of the lift line of the aircraft [1/rad]
g         = gravitational acceleration [m/s²]
S = wing area [m²]
lm = geometric mean chord [m]
r₀       = air density at sea level [kg/m³] r₀ = 1,225 kg / m³
r = air density [kg/m³]
m = weight of the aircraft [kg]
k = mitigation factor, which is determined as follows:

where μ is the relative mass ratio of the aircraft, which is calculated as:

It is not necessary that the value of n, which is determined by the above relation, be greater than:

UL 2 § 345 Loads with flaps extended

1. If the airplane has flaps, a positive operating factor of 2,0 must be considered; considering flap positions from "retracted" to "maximum positive deflection" and speeds up to design speed V_F.

2. It must be considered that the aircraft meets the conditions of points UL 2 § 321 and UL 2 § 331 and also points UL 2 § 333 to UL 2 § 337 with the flaps in the position from "retracted" to "maximum negative deflection".

UL 2 § 361 Engine bed load

1. The engine bed and its attachment must be designed for the following load cases:

a. operating torque load from the propeller, which corresponds to the take-off power and the respective propeller revolutions with the simultaneous application of 75% of the operating load from case A according to point UL 2 § 333.
b. operating torque load from the propeller, which corresponds to the maximum continuous power and the respective propeller revolutions under the simultaneous application of the operating load from case A according to point UL 2 § 333.

2. For conventional piston engines with a direct ("hard") propeller drive, the operating torque from the engine used in paragraph 1 above is calculated by multiplying the mean (average) torque by the appropriate factor according to the following table:

	two-stroke engine	four-stroke engine
1 cylinder	6,0	8,0
2 cylinders	3,0	4,0
3 cylinders	2,5	3,0
4 cylinders	1,5	2,0
5 or more cylinders	1,33	1,33

Note:
The term "hard" transmission means direct drive, gear or belt drive. For other types of drive (e.g. centrifugal clutch) and non-conventional motors, the relevant coefficient must be consulted with the responsible authority.

UL 2 § 363 Lateral loading of the engine bed

The engine bed and its mounting must be designed for a side load with an operating factor of not less than one-third of the operating factor from envelope point A (1/3 n1).

IV. Control surfaces and control systems

UL 2 § 395 Management systems

All parts of the main control system between the stop and the control surface must be designed for a load that corresponds to at least 125% of the control surface load according to points UL 2 § 423 and UL 2 § 441 and also according to point UL 2 § 455.

In no case may the load in any part of the system be less than 60% of the pilot's forces according to point UL 2 § 397.

UL 2 § 397 Loading by forces from the pilot

All control systems for direct control of the aircraft around its longitudinal, transverse or vertical axis (main control system) and other control systems that affect the behavior of the aircraft in flight, as well as their mounting or support points, must be designed up to the stops ( last connection) to the service loads that are defined in the force table from the pilot.

For non-conventional control systems (eg side control stick control), lower pilot forces may be permitted by the responsible authority if it can be demonstrated that the forces in the table cannot be used.

Management	Acting force [daN]	Method of introduction of forces (assuming the use of a single control stick)
Height control	35	By pulling and pushing the control lever
Lateral steering	20	Side-to-side movement of the control stick
Steering and other foot-operated controls	90	By pressing forward on the pedal (directional control

The steering system for turning must be designed for a load of 90 daN on each pedal while simultaneously acting on both forward pedals.

UL 2 § 399 Dual control system

The dual control system must be designed for the following loads:

a. simultaneous action of both pilots in the same direction a
b. simultaneous action of both pilots in the opposite direction,

while 75% of the forces specified in point UL 2 § 397 are considered for both pilots.

UL 2 § 405 Secondary control systems

Secondary control systems, eg for extending or retracting the landing gear, flap control, balance, etc., must be designed for the expected maximum forces that the pilot can exert.

Note:
Design forces when acting on hands or feet must not be less than:

a. for wheels, handles, etc. controlled by finger or hand P = 15 daN,
b. for levers and wheels controlled by the whole arm without using the body's own weight P = 35 daN,
c. for levers and handles controlled by the whole arm with support or using the body's own weight P = 60 daN,
d. force exerted by the legs while resting on the seat (e.g. force on the foot brakes) P = 75 daN.

UL 2 § 411 Stiffness and deformation of the control system

The range of movement of the control surfaces that can be used by the pilot must in no case be dangerously reduced by the flexible deformation of the control circuit.

Note:
It will normally be considered acceptable if the circumference of each main control meets the proportional deformations recommended in this paragraph in the stiffness tests. By introducing the loads shown in the following table, no part of the steering system should be lengthened or shortened by more than 25%.

The proportional strain is defined as:

a = movement of the steering controller in the cabin when the force is applied by the pilot with the corresponding control surface locked in the neutral position.
A = possible positive (negative) movement of the steering control in the cab measured from the neutral position valid for the disengaged steering.

Management	force [daN]	introduction of force
Height (tilting)	12	pull, push the handle
Transverse (bending)	8	handle movement to the side
Directional (turning)	15	push the pedal

If the proportional distortion in the primary control system exceeds 25%, then special attention must be paid to the thorough demonstration of Sections III and VII, Subpart B.

V. Horizontal tail surfaces

UL 2 § 421 Balancing load

1. Balancing load is the load required to maintain balance under any given flight conditions without pitching accelerations about the transverse axis.

2. Horizontal tail surfaces must be designed for such balancing loads as will occur at any point on the turn envelope and at flap positions in accordance with UL 2 § 335 and UL 2 § 345.

UL 2 § 423 Rotating loads

The horizontal tail surfaces must be designed for the turning loads that can be expected during pilot-induced turns at all speeds up to VD.

Note:
The load must be determined for the sudden deflection of the height control, taking into account the following cases:

a. velocity V_A, maximum upward deflection,
b. velocity V_A, maximum downward deflection,
c. velocity V_D, one third of the maximum upward deflection,
d. velocity V_D, one third of the maximum downward deflection.

In doing so, the following assumptions must be made:

1. The airplane is initially in level flight and neither its position nor its speed changes.
2. Loads are balanced by inertial forces.

Note:
If a more reliable calculation is not available, it is possible to use the recommended VOP loads listed in Annex III, paragraph 1.

UL 2 § 425 Gust loads

If a more accurate calculation corresponding to actual conditions is not available, the forces acting on the horizontal tail surfaces must be calculated as follows:

where:
P_{Terms & Conditions}   = force on horizontal tail surfaces [N]
P₀       = balancing force on the horizontal tail surface that acts before the gust load [N]
r_o       = air density at sea level [1,225 kg/m³]
k_{Terms & Conditions}   = mitigating factor, unless a more accurate calculation corresponding to actual conditions is carried out, the same value as for the wing can be used
S_{Terms & Conditions}   = area of ​​horizontal tail surfaces [m²]
a_{Terms & Conditions}   = slope of the lift line of the horizontal tail surfaces [1/rad]
U = gust speed [m/s]
V = flight speed [m/s]
de/da = derivative of the current skew according to the angle of attack at the VOP location

UL 2 § 427 Unsymmetrical loads

The effect of the propeller jet on the loading of fixed tail surfaces and rudders must be considered if an increase in loading due to this effect can be expected.

Unless proven otherwise, it is assumed that one half of the VOP is subjected to 100% and the other half to 70% of the maximum load under symmetrical flight conditions.

VI. Vertical tail surfaces

UL 2 § 441 Rotating loads

The vertical tail surfaces shall be designed for the turning loads which may occur under the following conditions:

1. maximum directional control deviation at speed V_A,
2. one third of full directional control deflection at V speed_D.

UL 2 § 443 Gust loads

1. The vertical tail surfaces must be designed for lateral gust loads up to the values ​​according to point 2 of UL § 333.

2. If a more accurate calculation corresponding to the actual conditions is not available, the forces acting on the vertical tail surfaces must be calculated as follows:

where:
P_SOP   = force on vertical tail surfaces [N]
r₀       = density of air at sea level (1,225 kg/m³)
V = flight speed [m/s]
S_SOP   = area of ​​vertical tail surfaces [m²]
a_SOP    = slope of the lift line of the vertical tail surfaces [1/rad]
U = gust speed [m/s]
k_SOP    = mitigating gust factor, which is determined as follows:

M = maximum weight of the aircraft [kg]
l_mS      = geometric mean chord of the vertical tail surface [m]

UL 2 § 444 T-shaped tail surfaces

1. For airplanes in which the horizontal tail surface is supported by a vertical tail surface, the tail surfaces and their attachment, including the rear part of the fuselage, must be designed for the prescribed load on the vertical tail surfaces and for the yawing moment induced by the horizontal tail surface acting in the same direction.

2. Unless a more accurate calculation is made, for T-shaped tail surfaces, the induced pitching moment from the gust load can be determined as follows:

Where:
M_r      = induced pitching moment of horizontal tail surfaces [Nm]
S_{Terms & Conditions}   = area of ​​horizontal tail surfaces [m²]
b_{Terms & Conditions}   = span of horizontal tail surfaces [m]

VII. Additional conditions for tail surfaces

UL 2 § 447 Combined loading of tail surfaces

1. Under the condition that the aircraft is in a load state corresponding to point A or D of the turning envelope (conditions with a higher balancing load must be considered), the load on the horizontal tail surfaces is combined with the load on the vertical tail surfaces according to point UL 2 § 441.

2. It must be assumed that 75% of the load according to point UL 2 § 423 (for horizontal tail surfaces) and 75% of the load according to point UL 2 § 441 (for vertical tail surfaces) are acting simultaneously.

3. For UL airplanes with "V" shaped (butterfly) tail surfaces, gust must be allowed for at V speed_B, which acts perpendicular to the tail surface.

VIII. Wings

UL 2 § 455 Wings

The aileron must be designed for a control load that meets the following conditions:

1. it must be considered that at the maximum deflection of the ailerons and at the speed of V_A acts on the plane by a factor of n = 2,66
2. it must be considered that at 1/3 of the maximum deflection of the ailerons and at a speed of V_D acts on the plane by a factor of n = 2,66.

Note:
If a more accurate calculation is not available, it is possible to use the recommended loads listed in Annex III, paragraph 2.

IX Ground load

UL 2 § 471 In general

The operational ground loads specified in this section are defined as the external loads and inertial forces acting on the structure of the airplane. Under all specified ground loading conditions, the external reaction must be in balance with the inertial forces and moments so that it corresponds to or approaches the actual conditions on the safe side.

UL 2 § 473 Ground loading conditions of the chassis

1. The requirements of the following paragraphs must be met for the maximum design weight.

2. The chosen operational multiplier, acting at the center of gravity of the aircraft, must not be less than the value that is reached during landing with descending speed

with the restriction that it must not be less than 1,5 m/s and must not be greater than 3 m/s.

3. During the impact, a lift force corresponding to a maximum of 2/3 of the weight of the aircraft may be considered at the center of gravity of the aircraft. When considering such a buoyant force, a multiple of inertial forces can be substituted for the ground load multiple, reduced by the ratio of the considered buoyant force to the weight of the aircraft.

The operational load factor at the center of gravity of the aircraft is determined from the relationship:

n_pr = n_k + 0,67

where n_k is the operational multiple on the chassis wheels:

where:
y = total path of shock absorption-total drop of the center of gravity [m]
(sum of expected tire compression y_PN and dampers y_TL)

y = y_PN + Y_TL

y_ef    = effective damping path [m]
y_ef = 0,5·y_PN + 0,5·y_TL for rubber or spring shock absorbers [m]
y_ef = 0,5·y_PN + 0,65·y_TL when using hydraulic shock absorbers [m]

Note:
If n_pr is greater than 4 according to the calculation, then it is necessary to check the fastening of all concentrated masses (engine, fuel tanks, crew seats) for a load corresponding to the calculated n_pr.

UL 2 § 479 Basic landing conditions for landing gear

The methods of loading the chassis are listed in Appendix IV.

UL 2 § 485 Lateral load conditions

To determine the landing gear lateral load, the airplane is assumed to be level with the main landing gear wheels touching the ground and:

1. a force equal to 1,34 times the maximum weight of the aircraft (G) acts at the center of gravity of the aircraft, equally distributed to the main wheels,

2. the operating lateral inertial forces of 0,83 G at the center of gravity of the airplane are distributed between the wheels of the main landing gear such that:

a. 0,5 G acts on one side towards the torso,
b. 0,33 G acts on the other side away from the hull. See Annex IV.

UL 2 § 493 Load conditions during braking

It must be demonstrated that the braked wheels of the chassis (with tires and shock absorbers in static position) will comply with the load when:

a. the vertical operating load per wheel is 0,67 G,

UL 2 § 497 Additional load conditions for stern landing gear

1. The maximum force from the aft landing gear load analysis acts at an angle of 45° in the wheel pivot towards the rear.

2. The maximum force from the static reaction acts in the vertical and lateral axes simultaneously - see Appendix IV.

For the construction of the spur itself and its attachment to the surrounding supporting structure, including the tail surfaces with fixed balancing masses, the load on the spur when landing on the spur (the main landing gear is not in contact with the ground) can be determined as follows:

where:
P = force on spur [N]
m = weight of the aircraft [kg]
g      = gravitational acceleration [m/s²]
i_y      = radius of inertia of the aircraft [m]
L = distance of the spur from the center of gravity of the aircraft [m]

Warning:
If you cannot value i_y determined in a more precise way, a value can be substituted

i_y = 0,225 * L_tr

In this case, L corresponds_tr total hull length without rudder.

UL 2 § 499 Additional load conditions for nose landing gear

To determine the ground load of the nose landing gear and their attachment, assuming that the compression of the shock absorbers and tires corresponds to the rest state, the following conditions must be met:

1. For rearward loads, the force components acting in the axis must have the following magnitude:

a. the vertical component corresponds to 2,25 times the value of the stationary static load of the wheel,
b. the resistance component corresponds to 0,8 times the vertical component.

2. For loads acting forward, the components of the service force acting in the axis must have the following magnitude:

a. the vertical component corresponds to 2,25 times the value of the stationary static load of the wheel,
b. the forward-oriented component corresponds to 0,4 times the vertical component.

3. For lateral loads, the operating components of the operating forces at the point of contact with the ground must have the following magnitude:

a. the vertical component corresponds to 2,25 times the value of the stationary static load of the wheel,
b. the lateral component of the load corresponds to 0,7 times the vertical component.

X. Emergency landing conditions

UL 2 § 561 In general

1. Although an airplane may be damaged in an emergency landing, it must be designed so that each person on board must be protected from the effects of the forces specified in the following paragraph.

2. The structure must be designed so that each person on board, if belts are properly used, has a good chance of escaping serious injury in the event of an emergency landing under the following conditions:

A person on board is exposed to the following numerical inertial forces, which act independently of each other:

up	4,5 g
ahead	9,0 g
to the side	3,0 g
Dolu	4,5 g

3. Fuel tanks must withstand the load of inertial forces specified in paragraph 2 of this point without damage.

XI. Other loads

UL 2 § 597 Loads with isolated masses

The fastening of all isolated masses that are part of the aircraft equipment (including the necessary load for adjusting the position of the center of gravity) must be designed to withstand the load corresponding to the maximum design multiples of flight and ground loads, including the conditions during an emergency landing according to point 2 of UL § 561.

 

TITLE D - DESIGN and CONSTRUCTION

UL 2 § 601 In general

The strength of parts that have a significant effect on operational safety and for which an unambiguous calculation cannot be performed must be proven by tests.

UL 2 § 605 Manufacturing methods

The production methods must guarantee continuously error-free strength connections that are permissible with regard to maintaining the required strength under normal conditions expected in operation. If manufacturing processes (such as gluing, spot welding, heat treatment or plastic processing) require precise control for this purpose, they must be carried out according to approved work methods. Non-conventional production methods must be proven by appropriate tests.

UL 2 § 607 Securing connecting elements

Approved securing means and methods must be used for all fasteners used in primary structural strength connections, controls, and other mechanical systems critical to the safe operation of the airplane. In particular, self-locking nuts must not be used for screws that rotate in operation, unless an additional locking element is used that works other than on the principle of friction.

Note:
Aviation pins must not be used in places where accidental disconnection may occur during folding, catching on vegetation, clothing or manipulation by unauthorized persons.

UL 2 § 609 Construction protection

Each part of the supporting structure must:

1. be sufficiently protected in service against harmful effects or reduction of strength due to any causes including:

a. weathering,
b. corrosion,
c. abrasion, a

2. contain sufficient means for ventilation and drainage.

UL 2 § 611 Inspections

Such provisions shall be made to permit the inspection (including inspection of major elements of the primary structure and control system), accurate testing, repair and replacement of each part that requires monitoring and adjustment to ensure accurate function and proper operation, lubrication or maintenance.

UL 2 § 612 Assembly and disassembly

The design of the aircraft must be such that the possibility of damage or permanent deformation during assembly or disassembly, which can be carried out even by a person without special experience, is reduced to a minimum, especially in places where such damage is not clearly visible. The possibility of incorrect assembly must be excluded by the design measures used. The correct assembly of the aircraft must be easily controlled.

UL 2 § 613 Strength properties of materials and calculated values

1. The strength properties of the materials used must be documented by a sufficient number of tests so that the calculation values ​​can be determined on the basis of statistics.

2. The calculated value must be chosen so that the probability of insufficient strength of any part of the supporting structure, including the considered dispersion of material properties, is very low.

Note:
Material specifications must either be established as part of the certification process or conform to known standards. When determining the calculation values, the material values ​​may be changed or extended by the designer if necessary for production reasons (e.g. with regard to the manufacturing method, to forming, machining or subsequent heat treatment).

3. If, under normal conditions, a temperature is reached on some part of the load-bearing structure or strength connection, which has a significant effect on strength, this effect must be taken into account.

Note:
The temperature of parts of the structure up to 54 °C is considered as normal operating temperature.

UL 2 § 627 Fatigue strength

The structure must be designed and constructed to exclude locations of stress concentrations and high stress levels and to take into account the effects of vibration. Materials that have poor crack propagation properties must not be used. All assemblies, especially in the supporting (primary) structure, must be easily controllable. Elastic varnishes or elastic protective coatings must not be used.

UL 2 § 629 Prevention of flutter and structural strength

1. It must not occur on the airplane in any configuration and at any permissible speed at least up to V_D flutter, aerodynamic shaking (divergence) and control reversal. The controllability and stability of the airplane must not be dangerously sensitive to deformation of the structure. in the range of permissible speeds, the structure must have such damping that the aeroelastic oscillations die out quickly.

2. Proof of agreement with the requirements of paragraph 1 must be made in the following way:

a. systematic tests for the excitation of flutter in flight at speeds up to V_DF. These tests must demonstrate that when approaching V_DF no drop in damping occurs.
b. test flights, during which it will be proven that when approaching V_DF there is no sharp drop in control effect about all three axes and that there are no signs of impending flapping of the wings, tail surfaces and fuselage during static stability and balancing positions.
c. for airplanes whose speed V_D is higher than 200 km/h, it must be demonstrated by means of a ground vibration test and subsequent flutter analysis that there is no flutter up to 1,2 V_D, even before the flight tests.

I. Control surfaces

UL 2 § 655 Development

1. Movable control surfaces must be designed so that there is no collision of either control surfaces with each other, or control surfaces with other fixed parts of the structure, if one of the surfaces is in any position and the others move in the entire range of deflections. This requirement must also be met when:

a. operating load (positive or negative) for all control surfaces and in the entire range of their deviations a
b. operational load on the supporting structure of the aircraft, with the exception of the control surfaces.

2. If an adjustable stabilizer is used, it must be equipped with stops that will limit the range of adjustment to allow safe flight and safe landing.

UL 2 § 659 Mass balancing

Fixings and connection nodes of concentrated mass balances on the rudders must be designed for the following operating loads:

1. 24 g perpendicular to the plane of the control surface,
2. 12 g forward and backward,
3. 12 g parallel to the axis of the rudder.

II. Control systems

UL 2 § 671 In general

Every control must work easily, smoothly and reliably in order to safely serve its purpose.

UL 2 § 675 Stops

1. Each control system must have stops that safely limit the range of deflection of each aerodynamic surface that is active in this system.

2. The stops must be placed so that due to wear, clearance or adjustment of the steering, there is no change in the steering characteristics, which would be caused by a change in the range of movement of the steering surface.

3. Each stop must bear a load that corresponds to the design conditions for the stop according to point 2 of UL § 397.

4. For airplanes controlled by moving the center of gravity, for which no conventional control stops can be used to limit the control forces of the pilot, it must be demonstrated that the range of center of gravity deflections or control movements is such that the pilot cannot induce dangerous loads on the surrounding structure.

UL 2 § 677 Management of the balancing system

1. Appropriate measures must be taken to prevent inadvertent, incorrect or sudden balancing action. A device must be installed near the balance controller that shows the pilot the position of the balance surface within the range of its possible deflection. These devices must be visible to the pilot and designed and located to prevent confusion.

2. The control of auxiliary control (balancing) surfaces must be self-locking, if the surface does not contain sufficient balancing and there are no dangerous tendencies to flutter. The self-locking control of auxiliary surfaces must show sufficient rigidity and reliability in the part of the system that lies between the auxiliary surface and the connection of the braking member to the supporting structure of the aircraft.

UL 2 § 679 Interlocking device in the control system

If a device is used which serves to lock out the control system while the airplane is on the ground, measures must be taken which:

a. clearly warn the pilot that the blocking device is in operation,
b. prevent the blocking device from being activated in flight.

UL 2 § 683 Functional tests of control systems

Functional tests must prove that in a system that is dimensioned for the load according to points UL 2 § 397 and UL 2 § 399, the following will not occur:

a. getting stuck or obstructed,
b. excessive friction, a
c. excessive bending,

if the steering is controlled from the cockpit.

UL 2 § 685 Structural elements in the control system

1. All structural elements of each steering system shall be designed and constructed to prevent jamming, jamming and blocking which may be caused by passengers, loose objects or frozen moisture.

2. Means must be built into the cockpit to prevent the penetration of foreign bodies into places where they could cause the system to block.

3. All parts of the flight control system must be designed or clearly and permanently marked in such a way as to minimize the risk of incorrect assembly which could lead to incorrect control operation.

UL 2 § 687 Springs (spring members)

The permissibility of the use of all spring elements used in the control system must be proven by tests during which operating conditions are simulated. It must be proven that failure of one spring does not lead to flutter or dangerous changes in operating characteristics.

UL 2 § 689 Ropes and rope systems

1. All ropes, rope ends, turnbuckles, rope joints and pulleys must conform to approved specifications. In particular, the following apply:

a. No rope less than 2 mm in diameter may be used in primary control systems. All ropes must be installed in such a way that there are no dangerous changes in the tension in the ropes in the entire range of deviations under operating conditions and also due to the expected temperature changes.
b. All rope lines, pulleys, terminals and turnbuckles must be accessible for visual inspection.

2. All pulley types and sizes must match the ropes with which they will be used. All pulleys must be fitted with tight-fitting guards to prevent slipping or jamming when the rope is loose. All pulleys must lie in the plane of the rope so that the rope cannot rub against the edges of the pulley.

Note:
The inner diameter of the guide groove of the pulley should not be less than 300 times the diameter of the individual wire of the rope.

3. Guides must be designed so that they do not change the direction of the rope by more than 3°, unless it is proven by tests or experience that a higher value is permissible. The radius of curvature must not be less than the sheave radius for the same rope.

Note:
In guides made of Teflon or similar material, the direction of the rope can be changed up to 10°.

4. On all parts that make an angular movement, tensioners must be connected so that they can be freely adjusted in the entire range of deflections.

UL 2 § 697 Control of flaps

1. Each flap control must be designed so that the flaps in each position necessary to meet the performance requirements do not change their position, except when moving to achieve the required position, unless it is demonstrated that such movements are not dangerous.

2. The control of the flaps must be designed in such a way that there can be no unobserved extension or rearrangement. The applied control forces and changes in speed must not be so great at any permissible speed as to affect the operational safety of the airplane.

UL 2 § 701 Connecting flaps

The movement of flaps located symmetrically from the plane of symmetry must be mechanically linked to ensure their simultaneous movement, unless otherwise ensured that the airplane has safe flight characteristics when the flaps are retracted on one side and extended on the other side.

III. Chassis

UL 2 § 721 In general

The aircraft must be designed to be able to land without endangering the people on board on a short grassy area.

IV. Flight deck design

UL 2 § 771 Pilot space - General

The cockpit and its equipment must allow each pilot to perform his tasks without excessive concentration or fatigue.

UL 2 § 773 View from the cockpit

Each pilot space must be designed to:

a. the pilot's range of vision was sufficiently large, clear and undisturbed with regard to the safe control of the aircraft, and
b. the rain did not significantly affect the pilot's view of the route during normal flight and during landing.

Note:
Fulfillment of the conditions can be achieved by a corresponding opening in the cockpit glazing.

UL 2 § 775 Wind shields and windows

Glazing and windows must be made of material that does not cloud and whose fragments cannot lead to serious injury to persons on board.

UL 2 § 777 Controls and controls in the cockpit

1. All controls and controls in the cockpit must be located to allow for optimal pilot operation and to prevent confusion and inadvertent or unobserved activation.

2. Controls and controls in the cockpit must be located so that the pilot can control each control organ up to its full deflection with lap and shoulder belts on. At the same time, he must not be restricted either by clothing (including winter gear) or by structural elements of the cockpit.

3. In dual-control airplanes, at least the following controls and controls must also be accessible from both pilot's seats:

a. power unit power control (gas inlet),
b. flaps,
c. balancing,
d. device for opening the cabin cover, a
e. rescue system.

4. Controllers of secondary control systems must maintain each set position without requiring the constant attention of the crew and must not be inclined to change position arbitrarily due to load or vibration.

Note:
If the airplane is equipped with a balance system, the requirement that the balance be controllable by both pilots may be waived if it is demonstrated that, at the most unfavorable position of the balance surface, the steering forces are sufficiently low and that no steering difficulties arise.

UL 2 § 779 Sense of movement and action of elements and controls in the cockpit

1. Controls and controls in the flight deck shall be designed to operate as follows:

Control and control element	A sense of movement and effect
lateral steering	right (clockwise): right wing down
altitude control	aft: bow up
side proceedings	right pedal forward: bow right
balancing	corresponding steering movement
flaps	drag: the flaps extend (extend) or fold down
gas inlet	forward: power increases
building a propeller	forward: increasing the number of revolutions
mixture	forward or upward: a rich mixture
switches	down: off
rescue system	pull: activation of the rescue system

2. For airplanes that are not controlled by aerodynamic means, the direction of movement of the main controls must be specified in the Flight Manual. Auxiliary control systems must correspond to the sense of movement established in the previous paragraph.

UL 2 § 780 Color markings of controls and controls in the cockpit

Emergency system controls must be red.

UL 2 § 785 Seats and safety belts

1. Each seat and its attachment to the supporting structure must be dimensioned for the weight of the person on board at least 100 kg and the maximum multiples that correspond to the specified flight and ground loads, including emergency landing conditions according to point UL 2 § 561.

2. Seats, including cushions, must not deform to such an extent that the pilot cannot safely reach controls and controls when loaded in flight in accordance with Subpart C, or that incorrect use may occur.

3. The strength of safety belts must not be lower than that which results from the numerical load of flight and landing cases, as well as from the conditions of an emergency landing, taking into account the geometry of the arrangement of the belts and the seat.

4. Each crew member must wear a four-point seat belt. Each seat belt must be designed to maintain the crew member in the original seated or supine position during all accelerations that may occur in flight or during an emergency landing.

UL 2 § 786 Protection against injury

Fixed parts of the supporting structure or fixed parts of the equipment must, if necessary, be padded to protect the person(s) on board from injury in a light emergency landing.

UL 2 § 787 Luggage compartment

1. Each baggage compartment must be sized for the maximum load weight shown on the label and for the critical load distributions at the respective largest multiples arising from the flight and landing load cases.

2. Each baggage compartment shall be so equipped as to protect the occupants against injury which could be caused by the movement of the contents under a forward acceleration of 9,0 g.

UL 2 § 807 Emergency exit (emergency exit)

1. The cockpit of the aircraft must be equipped in such a way that in case of danger it can be exited without any problems and quickly.

2. In the case of a closed cockpit, the opening system must be simple and with unambiguous controls. It must work quickly and be designed so that it can be operated by every person on board strapped to a seat and also be operated from the outside.

UL 2 § 831 Ventilation

1. If the cockpit is enclosed, it must be sufficiently ventilated under normal flight conditions.

2. The concentration of carbon monoxide must not exceed a concentration of 1 part in 20 parts of air.

3. In the case of an open or uncovered cockpit, care must be taken to ensure that engine exhaust cannot be inhaled by the crew.

 

TITLE E - PROPULSION SYSTEM

I. In general

UL 2 § 901 Definition and development

1. The applicant must demonstrate that each combination of engine, exhaust system and propeller on the airplane for which the airworthiness certificate is being carried out is compatible with that airplane, works satisfactorily and operates safely within the specified conditions.

2. The propulsion system includes all parts which:

a. are necessary to derive forward thrust, a
b. have an effect on the safety of the drive unit.

3. The propulsion unit must be designed, arranged and built in such a way that:

a. safe operation was ensured,
b. was accessible for necessary inspections and maintenance, a
c. the installation regulations of the engine manufacturer were followed.

Note to paragraph 3.a:
A test run of the complete propulsion system lasting at least 3 hours will be recognized as a certificate. First, the engine must run for 1 hour at 75% of maximum continuous power. Then it is necessary to proceed according to the following program:

start and stop 10 times,
start and idle for 5 minutes,
5 minutes full power,
5 minutes of cooling at low idle speed,
5 minutes full power,
5 minutes of cooling at low idle speed,
15 minutes 75% continuous power,
5 minutes of cooling at low idle speed, a
15 minutes full power.

Switch off the engine and let it cool down, repeat the program. At the same time, there must be no obvious damage to any part of the drive system or any of its components.

UL 2 § 903 Motors

1. In general

The provisions of this article apply to reciprocating engines which are designed and built in the usual manner. The regulations are applicable to piston engines for ultralight aircraft.

2. Design and construction

Engines intended for use on an ultralight airplane that have not yet been certified under JAR-E, JAR 22 Section H or FAR Part 33 may be certified as ultralight engines. The engine and its accessories must be designed and constructed in such a way that the possibility of failure during operation is reduced to a minimum.

a. Serial and modified engines

In the case of automobile or motorcycle engines, or their units, which have been introduced into series production and if no fundamental change has been made to these engines or units, an acceptable level of reliability is assumed, which is documented by the experience of operating such engines. These are permitted for a given UL aircraft category.

b. The installation and modifications of the engines will be judged according to experience from operation. Modifications not verified by the test will require an additional engine reliability test.

3. Engine tests prescribed in aircraft type tests

If an engine is selected for a particular type of UL airplane, it may be tested in that UL airplane as a 50-hour flight test. The flight test must include, as a minimum:

– 100 starts
– 10 flights lasting at least 1 hour
– 60 ascents to a height of at least 500 m above the terrain, while the flight with take-off power must last at least 5 minutes continuously.

At least 30 of these climbs must be done in summer temperatures (minimum 20˚C).

4. Engine tests prescribed in engine type tests for UL airplanes:

a. Engine Reliability Test.

The applicant must demonstrate that the engine is capable of operating at the prescribed duty cycle of 25 hours without significant defects. Work cycles follow each other periodically. The manufacturer specifies in advance the maintenance operations on the engine that will be performed during the test.

One long-term reliability test cycle will be performed as follows:

starting and idling	5 min.	Repeat 2 times
max. output	5 min.
cool down and stop	5 min.
starting and idling	5 min.
max. output	5 min.
max. continuous power	60 min.
cool down and stop	5 min.
total time of one cycle	1 hours 45 minutes

The reliability test is performed on the ground. Engine braking is performed using a test propeller.

b. Test operation in an ultralight aircraft.

The applicant must demonstrate that the engine in the proposed propulsion system on the aircraft for which the certified engine is intended is in accordance with the function of the aircraft and will demonstrate reliability in a 50-hour test according to paragraph 3 of point 2 of UL § 903.

UL 2 § 905 Propellers

1. The propeller must meet UL 2 Title J specifications.

2. Engine power and propeller shaft speed must not exceed the limits for which the propeller is certified or approved.

UL 2 § 925 Propeller safe distance

For an uncovered propeller, the safe distance at maximum weight, worst center of gravity and worst blade setting must not exceed the following values:

1. Ground clearance: at least 170 mm between the propeller and the ground (for each nose landing gear airplane) or 230 mm (for each spur landing gear airplane). In doing so, the landing gear must be statically compressed and the aircraft is either in the normal horizontal take-off position or in the taxiing position, whichever position is more critical. In addition, a safe distance must be maintained in a horizontal position at take-off, if

a. the critical tire is completely depressurized and the relevant chassis strut is statically loaded or

2. the critical strut of the chassis is at the stop and the corresponding tire is statically loaded.

3. Distance from other parts of the aircraft

a. The smallest radial distance of 25 mm between the end of the propeller blade and the adjacent parts of the aircraft plus the additional radial distance that is necessary to dampen harmful vibrations (pay special attention to the springing of flexible hinges).

4. At least 13 mm longitudinal distance between the propeller blades or their fairing and other parts of the aircraft.

5. A safe distance between other rotating parts of the propeller or propeller hub (including its cover) and other parts of the aircraft must be maintained under all operating conditions.

b. Distance from persons on board

There must be sufficient distance between the propeller(s) and the person(s) on board so that the person on board who is strapped into the seat cannot come into contact with the propeller(s) during careless operation.

II. Fuel system

UL 2 § 951 In general

1. Each fuel system must be designed and arranged to safely provide such fuel flow and pressure as are intended for proper engine operation under all normal operating conditions.

2. Each fuel system must be arranged so that fuel can be drawn from only one tank to supply the engine, unless the tanks are interconnected so that they are emptied simultaneously.

3. The fuel system must be constructed in such a way that it cannot be blocked by generated fuel vapors.

UL 2 § 955 Fuel flow

1. Fall system.

The fuel flow rate during auto-drop filling (both from the main and the reserve supply) must be at least 150% of the fuel consumption for the engine at the maximum take-off power of the engine.

2. System with fuel pump.

The fuel flow in each priming system (both the main and backup pumps) must be at least 125% of the fuel consumption at the specified maximum engine take-off power.

UL 2 § 959 Inexhaustible amount of fuel

The inexhaustible amount of fuel must be determined for each tank as the amount at which the first signs of fuel failure occur under the most unfavorable conditions for such fuel supply during take-off, climb, approach and during landing. This amount must not exceed 5% of the tank volume.

UL 2 § 963 Fuel tanks - General

1. Each fuel tank must withstand without failure the vibrations, inertial forces, hydrostatic load and external forces acting on it during operation.

2. If "spillover" of fuel in the tank can substantially change the position of the center of gravity of the airplane, measures must be taken to reduce the "spillover" to an acceptable level.

UL 2 § 965 Tests of fuel tanks

Each fuel tank must withstand a pressure of 0,01 MPa without damage or leakage.

UL 2 § 967 Fuel tank installation

1. Each fuel tank must be secured in such a way as to prevent concentrated loading caused by the fuel's own weight. In addition

a. a soft spacer must be placed (if necessary) between the tank and its mounting to prevent chafing of the tank and

2. the materials that are used to attach or cover the attachment must not be absorbent or must be modified in such a way as to prevent the absorption of fuel.

3. Each space in which the tank is built must be ventilated and equipped with drainage to prevent the accumulation of flammable liquids or vapors in it. Every space adjacent to the space in which the tank is built must also be well ventilated and equipped with drainage.

4. No fuel tank may be located where it could be hit in the event of an engine fire.

5. It must be proven that the location of the tank installation site will in no way limit the operation of the aircraft or the freedom of movement of the persons on board, and that the escaping fuel cannot directly affect the persons on board.

6. Damage to the structure due to a hard landing in which the landing gear load exceeds the required numerical value, but the total load does not exceed the values ​​for an emergency landing, must not lead to the destruction of the fuel tank or fuel lines.

UL 2 § 971 Fuel tank drain sump

1. Each fuel tank, if permanently installed, must have a drainable sump effective in all normal positions on the ground and in flight, the volume of which is either 0,10% of the tank volume or 120 cm³ (the larger value is decisive). If this is not the case,

a. an accessible sludge container or 25 cm drain tank must be built into the fuel system³ a

2. the fuel tank drain must be built in such a way that, in the normal position on the ground, water flows from all parts of the tank to the sludge tank or drain tank.

3. Drainage must be easily accessible and able to be easily put into operation.

4. Each fuel system outlet must be equipped with a device that can be safely manually or automatically locked in the closed position.

UL 2 § 973 Fuel tank filler necks

The filler necks of the fuel tanks must be outside the area intended for persons. Poured fuel must not flow into the space in which the fuel tank is located or into any other part of the aircraft, outside of its own tanks.

UL 2 § 975 Fuel tank venting

Each fuel tank must be vented at its top. Besides, they have to

1. each vent valve must be constructed and located in such a way that the risk of its clogging by ice or other foreign bodies is reduced to a minimum,
2. each vent must be designed to prevent suction of fuel due to negative pressure during normal operation,
3. every venting must be brought out into the free space.

UL 2 § 977 Fuel strainer and filter

1. A fuel cleaner (fuel filter) must be placed between the fuel outlet from the tank and the carburetor inlet (or engine driven fuel pump if installed).

2. At the outlet of each fuel tank there must be a cylindrical strainer with three to six meshes per centimeter (suction basket). The diameter of the strainer must be at least equal to the diameter of the outlet from the fuel tank and the length of the strainer at least twice the diameter of the outlet from the fuel tank.

3. Each strainer or each filter must be easily accessible for inspection and cleaning.

UL 2 § 993 Fuel lines and connections

1. Each fuel line must be installed and fixed in such a way as to prevent its excessive vibration and to withstand the loads created by the hydrostatic pressure of the fuel and from flight multiples.

2. Each fuel pipe that is fixed on such parts of the aircraft, the position of which can change with each other, must be equipped with a flexible member.

3. Flexible hoses must be proven to be usable for the intended purpose.

4. Leakage of fuel from any pipe or joint shall not directly impact hot surfaces or equipment so as to cause fire, or directly impact persons on board.

UL 2 § 995 Fuel taps and their control

1. A device must be installed that would allow the pilot to quickly shut off the fuel supply to the engine in flight.

2. The pipe section between the fuel shut-off valve and the carburetor must be as short as possible.

3. Each fuel shut-off valve shall have either firm stops or effective locking in the "open" and "closed" positions.

III. Oil system

UL 2 § 1011 In general

1. If the engine is equipped with an oil system, this system must supply the engine with a sufficient amount of oil at a temperature that does not exceed the maximum value for safe continuous operation.

2. Each oil system must have a usable volume that is sufficient to supply for the maximum flight time.

UL 2 § 1013 Oil tank

1. Oil tanks must be built in such a way that

a. have met the requirements according to point UL 2 § 967 paragraphs 1., 2. and 4. and
b. withstand all vibrations, inertial forces and hydrostatic loads that may occur in operation.

2. It must be possible to check the oil level without removing the cover parts (outside the filler cap) and without using tools.

3. If the oil tank is built into the engine compartment, it must be made of heat-resistant material.

UL 2 § 1015 Tests of oil tanks

Oil tanks must be tested in accordance with point 2 of UL § 965, whereby a pressure of 0,035 MPa must be used for pressure tests.

UL 2 § 1017 Oil pipes and connections

1. Oil lines must comply with UL Section 2 § 993 and each oil line and joint must be made of material capable of withstanding the maximum operating oil temperatures.

2. The vent pipe must be arranged so that

a. condensed water that could freeze or oil that could clog pipes could not accumulate in any place,
b. discharge from the vent pipe when oil foaming could not cause a fire hazard or so that the oil leaking from the pipe could not contaminate the wind shield in front of the person(s) on board or in front of the pilot(s).

IV. Cooling

UL 2 § 1041 In general

Powerplant cooling equipment must be capable of maintaining the temperature of all powerplant components and engine operating fluids within the limits established by the engine manufacturer for all expected operating conditions or required by the airplane manufacturer for those operating conditions.

V. Intake system

UL 2 § 1091 Air supply system

The air supply system must safely ensure the supply of the required amount of air to the engine under all expected operating conditions. The ingress of foreign bodies (grass, dirt, etc.) must be effectively prevented, preferably with a filter (sieve).

VI. Exhaust system

UL 2 § 1121 In general

1. The exhaust system must effectively ensure the safe removal of exhaust fumes without the risk of fire and without contamination of the space intended for people with carbon monoxide.

2. Any part of the exhaust system the surface of which is hot enough to ignite flammable liquids or vapors shall be so located and covered that leakage from any system through which flammable liquids or vapors pass cannot cause a fire by liquid contact or steam with any part of the exhaust pipes including their covers.

3. All parts of the exhaust system must be sufficiently far from adjacent flammable parts, or must be separated by heat-resistant covers.

4. The exhausts must not be located in the dangerous vicinity of the outlets of the fuel and oil system.

5. All parts of the exhaust system must be ventilated so that an unreasonably high temperature does not occur anywhere.

UL 2 § 1125 Exhaust pipes

1. The exhaust pipe must be made of refractory material and have provisions to prevent expansion damage after heating to operating temperature.

2. The exhaust and damping system must be mounted in such a way as to withstand all vibrations and inertial forces that arise during normal operation.

3. Parts of the exhaust system that are connected to parts that change their relative position must be flexibly connected.

VII. Engine control device and its accessories

UL 2 § 1141 In general

That part of each control device of the propulsion unit which is located in the engine compartment and must remain in operation even in the event of fire must be at least of heat-resistant material.

UL 2 § 1145 Ignition switch

1. Each ignition circuit must be equipped with a separate switch.

2. Each ignition current circuit must be switched on independently and its operation must not be conditioned by switching on any other switch.

3. Ignition switches must be arranged and designed to prevent their inadvertent use.

4. The ignition switch must not be used as a main switch for other circuits.

UL 2 § 1149 Propeller speed

Propeller speeds and settings must be limited to values ​​that effectively ensure safe operation under normal operating conditions.

1. During takeoff and climb at the speed recommended for best climb, the propeller must limit the engine speed with the throttle fully open so that it does not exceed the maximum allowable speed.
2. During gliding flight at speed V_NE with the inlet closed or the engine stopped, the propeller must not reach such revolutions that would exceed 110% of the maximum permissible revolutions of the engine or propeller (the lower value is decisive).

UL 2 § 1191 Fire wall

The engine must be separated from the rest of the airplane by a fire wall, cowling or equivalent means if the structural arrangement permits.

The fire wall or cover must be constructed in such a way that no dangerous amount of liquid, gas or flame can escape from the engine compartment to other parts of the aircraft.

The fire wall or cover must be heat resistant and protected against corrosion.

UL 2 § 1193 Engine covers and engine nacelles

The following applies to the engine cover:

1. Engine covers must be designed and secured to withstand the vibration, inertia and air forces that may be expected in service.

2. Enclosures must be equipped with such a device that the escaping substances from all parts of the enclosures can drain quickly and without residue in all normal configurations on the ground and in flight. Leaking substances must not be diverted to places where there could be a fire hazard.

3. All parts of engine covers exposed to high temperatures due to proximity to parts of the exhaust system or due to direct contact with exhaust gases must be made of heat-resistant material.

 

TITLE F - EQUIPMENT

I. In general

UL 2 § 1301 Function and development

1. Each piece of required gear must:

a. be of a type and design that will enable it to fulfill the expected function,
b. be built in such a way as to meet the specified limitations that apply to this equipment, a
c. work flawlessly after installation.

Note:
– Faultless function should not be impaired at temperatures below 0 °C, in heavy rain or in high humidity.
– If a radio station is built in, it must be proven that the aircraft's electrical system will not impair its function.

2. Instruments and other equipment must not endanger the safe operation of the aircraft either by themselves or by their effect on the aircraft.

UL 2 § 1303 Flight and navigation instruments

At least the following flight and navigation instruments must be installed:

1. Speedometer
2. Altimeter
3. Compass

UL 2 § 1305 Engine control devices

The following devices must be installed:

1. pressure gauge, thermometer and tachometer, which are recommended by the manufacturer or which are necessary for the engine to work within the specified limitations,
2. fuel quantity indicator for each fuel tank, which must be clearly visible even to the strapped-in pilot,
3. oil quantity indicator for each tank, e.g. dipstick.

UL 2 § 1307 Other equipment

Four-point seat belts must be installed for each person on board and must be capable of restraining them under the load that will produce the acceleration prescribed under emergency landing conditions.

Note:
See Fig. 3 for recommended installation of seat belts.

fig. 3: Recommended installation of safety belts

Warning:
1. If possible, it is recommended to install seat belts whose shoulder belts are adapted so that, in case of sudden braking, the lap belts (if they are not tightened tightly) cannot be pulled from the pelvis area to the level of the stomach, so that the user cannot stretch under the lap belt.
2. If there is a strap length of more than 152 mm between the attachment point of the shoulder straps and the upper edge of the backrest, such measures should be used to limit lateral movement, such as guides, which will safely establish an adequate distance between the shoulder straps so as to reduce the risk of injury or neck abrasion to the lowest possible degree.
3. If the seat back is strong enough and high enough that the geometry of the safety belts corresponds to Fig. 3 (i.e. 650 mm), the shoulder straps may be attached to the seat back or lead to the floor of the aircraft via a guide.
4. If the backrest is strong enough, it is possible to limit lateral movement of the body during the accelerations specified in point UL 2 § 561 for emergency landing conditions by using suitable means, e.g. shoulder belt guides.

II. Installation of devices

UL 2 § 1321 Arrangement and visibility

Flight and navigation instruments must be placed clearly and be easily legible for each pilot.

UL 2 § 1323 Speedometer system

1. The speedometer system should be calibrated so that the speedometer shows true sea level speed under MSA conditions with a maximum permissible system error of no more than ±6 km/h or ±5%, whichever is greater, within the following ranges speed:

a. from 1,2 V_S1 to V_NE with flaps retracted and
b. from 1,2 V_S1 to V_FE with flaps extended.

2. Determining the repair of the speedometer system must be done in flight.

3. The speedometer system must have a range of speeds from V_S0 to at least 1,05 times the V speed_NE.

Note:
Calibration of the airspeed system (IAS vs. EAS. speeds) must be specified in the flight manual.

UL 2 § 1325 Static pressure system

1. Any device whose body is connected to static pressure must be connected so that flight speed, opening and closing of windows, humidity or other extraneous influences do not significantly affect the accuracy of the device.

2. Total and static pressure systems must be designed and built in such a way that:

a. the safe drainage of condensed moisture was enabled,
b. wire penetration and excessive deformation or narrowing in bends were prevented, a
c. the materials used were durable, usable for the given purpose and protected against corrosion.

UL 2 § 1337 Power unit instruments

1. Devices and their management

a. Lines of engine control devices through which flammable liquids pass under pressure must meet the requirements according to point UL 2 § 993.
b. Every pipeline carrying flammable liquids under pressure must be fitted with baffles or have other safety devices at the source of pressure to prevent the release of excessive amounts of liquid in the event of a pipeline failure.

2. Any sight fuel sign used as a fuel quantity indicator shall be protected against damage.

III. Electrical systems and equipment

UL 2 § 1353 Design and installation of the accumulator

1. Accumulator batteries must be designed and built in accordance with the provisions of this section.

2. Explosive or poisonous gases that escape from the accumulator battery during normal operation or as a result of any possible malfunction of the charging device or the battery system must not accumulate in the aircraft in dangerous quantities.

3. Corrosive liquids or vapors that may escape from the accumulator must not cause damage to the surrounding supporting structure or to adjacent important parts of the equipment.

UL 2 § 1361 Arrangement of the main switch

1. The arrangement of the main switch must allow the power sources from the main bus to be promptly turned off. The disconnection point must be close to the sources controlled by the switch.

2. The arrangement of the master switch must be constructed so that the switch is easily recognizable and accessible to the pilot in flight.

UL 2 § 1365 Electrical conductors and accessories

1. Each electrical conductor must have a sufficient cross-section and must be properly routed, fixed and connected in such a way as to avoid the occurrence of a short circuit and the risk of fire as far as possible.

2. Every electrical device must have overload protection. No safety device may be intended for more than one circuit important for flight safety.

IV. Another gear

UL 2 § 1431 Radio communication and navigation equipment

Each used device from this equipment must meet the following conditions:

1. Equipment, including antennas, must not be dangerous in themselves, nor in the manner in which they are used, nor in their influence on the operational characteristics of the aircraft.
2. Equipment and devices for its operation and monitoring must be arranged so that they can be easily operated. The installation must be done in such a way that overheating is prevented, for example by means of sufficient ventilation.

 

TITLE G - OPERATING LIMITATIONS AND DATA

UL 2 § 1501 In general

1. The operational limitations listed in the following paragraphs and other limitations and data required for safe operation must be established.

2. Operating limitations and other data required for safe operation must be available to the pilot as prescribed in Subpart G.

UL 2 § 1505 Flight speeds

1. All airspeeds must be established as indicated airspeed (IAS).

Note:
The speed (EAS) that results from strength constraints should be recalculated accordingly.

2. Maximum speed V_NE must not exceed 90% of maximum speed (V_DF) proven by flight tests.

3. Maximum speed V_DF, proven by flight tests, must not exceed the maximum design speed V_D.

UL 2 § 1507 Turnover speed

The rotational speed must not exceed the design rotational speed V_A  according to point UL 2 § 335.

UL 2 § 1511 Speeds for operation of flaps

For each positive flap deflection setting, the maximum allowable speed for flap operation must not be V_FE greater than 90% of V speed_F according to point UL 2 § 335, on which the supporting structure is dimensioned.

UL 2 § 1515 Speeds for running gear

The maximum permissible speed for the operation of the landing gear V_LO must be determined for the retractable landing gear if it is lower than the maximum speed V_NE.

UL 2 § 1517 Speed ​​in strong turbulence

Maximum permissible speed in strong turbulence V_RA must not exceed the design speed V_B intended for the maximum requirements for gust in flight, determined according to Chapter C, point UL 2 § 335, paragraph 3.

UL 2 § 1519 Weight and position of the center of gravity

1. The maximum weight defined in point UL 2 § 25 must be established as an operational limitation.

2. The limitation of the position of the center of gravity defined in point UL 2 § 23 must be established as an operational limitation.

3. The weight of the empty aircraft and the corresponding position of the center of gravity must be determined in accordance with point UL 2 § 29.

UL 2 § 1521 Propulsion Unit Limitations

Limit values ​​for a power unit must be set so that the limits given by the engine or propeller manufacturer are not exceeded unless the applicant can satisfactorily demonstrate that higher limits can be safely used in connection with his aeroplane.

UL 2 § 1529 Operational – technical manual

The operational and technical manual of each UL aircraft must contain information that the applicant considers important for proper operation and maintenance. The applicant must include at least the following important information in the handbook:

1. device description,
2. the lubrication schedule, which must include the time between lubrication, the used lubricants and lubricating fluids that are applicable for individual devices,
3. pressures and electrical loads, permissible for individual devices,
4. tolerances and adjustment values ​​that are necessary for proper function, including deviations of control surfaces,
5. preparations for blocking, lifting and dragging on the ground,
6. distinction between supporting and secondary (auxiliary) structures,
7. the time between inspections and adjustments, which are necessary for the proper maintenance of the aircraft, and the method of their execution,
8. special preparations for aircraft repairs,
9. special control and adjustment preparations,
10. list of special tools,
11. data on weighing and determining the position of the center of gravity, which are necessary for the safe operation of the aircraft,
12. determination of running time limitations or lifetime limitations (replacement or maintenance) of parts, accessories and additional devices that are subject to these limitations,
13. materials that are necessary for small repairs,
14. cleaning and maintenance recommendations,
15. data on the construction, maintenance and checks of the rescue system,
16. procedures for disassembly, assembly and ground transport, information on support and anchor points,
17. list of labels and markings and their location.

I. Markings and labels

UL 2 § 1541 In general

1. The aircraft must be equipped

a. marking and labels according to points UL 2 § 1545 to UL 2 § 1547.
b. all other additional data, device markings and labels that are necessary for safe operation.
c. a fire-resistant registration plate, which contains at least the following information: type designation, manufacturer, serial number, year of manufacture, license plate.

2. All markings and labels specified in paragraph 1 of this point:

a. must be placed in a visible place, a
b. must not be easily erased, changed or not clearly visible.

3. The units of measurement used on the labels must be the same as those used on the gauges.

UL 2 § 1545 Speedometer marking

Marking	Range of speeds	Importance
green arch	1,1 VS1 up to VRA	normal operating range
yellow arch	VRA up to VNE	area with increased caution, only in calm weather
red vertical line	on VNE	maximum speed that must not be exceeded
white arch	1,1 VSO up to VFE	speed range with flaps fully extended
yellow vertical line	VA	maneuvering speed

UL 2 § 1547 Magnetic compass

If a compass is built in and the deviation is not lower than 5° for each course, there must be a table near the compass with the values ​​of the deviations for courses divided by a maximum of 30°.

UL 2 § 1549 Engine control devices

For each prescribed engine control device, if appropriate for the type of device:

– All maximum and, if given, minimum values ​​for safe operation must be marked with a red radial line.

UL 2 § 1553 Fuel quantity indicator

Each fuel quantity indicator must be marked to indicate "zero" in level flight if the fuel remaining in the tank corresponds to the inexhaustible quantity according to paragraph 2 of UL § 959.

UL 2 § 1555 Designation of control and control elements

1. Each steering and control element in the cockpit, with the exception of the main control, must be clearly marked according to its function and method of use.

2. The color marking of controls and control elements must agree with the colors specified in point UL 2 § 780.

3. The following applies to the control device of the fuel system:

a. Each fuel tank switching control shall be marked to visibly indicate its position for the applicable tank.
b. If for reasons of operational safety when using multiple tanks it is necessary to follow a certain order of use, the order in which the tanks must be used must be marked on the fuel tank switching controller or in its immediate vicinity.

UL 2 § 1557 Miscellaneous markings and labels

1. A label stating the maximum permissible weight of the cargo must be placed in each luggage compartment.

2. The fuel filler holes or their caps must have the marking of the mixing ratio for the fuel and oil mixture.

3. Every UL aircraft must have a label with the following text: This aircraft (sport flying device) is not subject to approval by the Civil Aviation Authority of the Czech Republic and is operated at the user's own risk. Acrobatic elements and intentional corkscrews are prohibited.

4. Every UL airplane must have an "Operational Data and Limitations" label clearly visible to the pilot with the following information:

– empty weight
– take-off weight
– payload
– weight in the luggage compartment
– the minimum weight of the pilot
– permissible speed V_NE
– fall speed V_S0
– design speed of rotation V_A
– if applicable, V speeds_FE, V_LO

Note:
If the maximum take-off weight of the aircraft can be exceeded by the amount of fuel, the weight of the crew and baggage must be specified on the label in connection with filling the tanks with fuel.

5. Marking of the pyrotechnic ZS (if the SFD is equipped with it):

a. Small symbol – place directly on the ZS, or in its immediate vicinity (for ZS built into the kite, place it on the outside of the fuselage in the area of ​​the shot).

Graphic form: a yellow isosceles triangle about 7 cm high with the inscription:

"PYROTECHNIC DEVICE.
BEWARE OF IMPROPER HANDLING.
HAZARD OF INJURY”.

6. Large symbol - place on the vertical tail surface from both sides, if possible on its fixed, immovable parts.

Graphic form: a yellow isosceles triangle about 13 cm high with the inscription:

"A PYROTECHNIC DEVICE IS PLACED IN THE AIRCRAFT.
BEWARE OF IMPROPER HANDLING.
HAZARD OF INJURY”.

II. Flight manual

UL 2 § 1581 In general

1. A Flight Manual must be created and submitted for each aircraft. It must contain at least the information specified in the two following paragraphs.

2. All data not specified in the two following paragraphs, as well as other data, if necessary for safe operation or based on an unusual type of construction, unusual method of operation or unusual operating characteristics, must be specified in the manual.

3. The airspeed data on the indicator and in the Flight Manual must be in the same units.

UL 2 § 1583 Operating restrictions

1. Speeds. The following information must be provided:

a. flight speed V_NE, V_RA, V_A, and if applicable, V_FE and V_LO along with defining their meaning,
b. traffic restrictions with regard to permissible wind speeds.

2. Weights. The following information must be provided:

a. maximum take-off weight,
b. the weight of the empty aircraft and the corresponding position of the center of gravity,
c. cargo distribution.

3. Power unit. Powertrain limitations must be established.

4. Load. The following information must be provided:

a. limitation of weight and extreme positions of the center of gravity according to point UL 2 § 25 together with parts of the aircraft included in the empty weight of the aircraft according to point UL 2 § 29,
b. data enabling the pilot to determine whether the position of the center of gravity and the distribution of the load with different load combinations is still within the specified permissible range,
c. data on the correct deployment of removable load for each load distribution that requires the use of removable load.

5. Turnovers. The permitted turns for which the certificate is made must be determined together with the permissible range of flap positions.

6. Multiples. The following positive operating turnover multiples are established:

a. multiple for V_A, which corresponds to the multiple at point A of the envelope according to Fig. 1
b. multiple for V_NE, which corresponds to the multiple at point D of the envelope according to Fig. 1

UL 2 § 1585 Operating data and procedures

1. Data on normal and emergency procedures must be provided, as well as any additional data necessary for safe operation.

2. Information must be provided on the procedures for performing a safe take-off and a safe landing at the appropriate specified distances according to point UL 2 § 51, including procedures for piloting in crosswinds and data on the highest permissible crosswind components. Safe engine-off landing procedures must be given.

3. The following information must be provided:

a. the speed for the best climb rate, which must not be lower than that which was used when proving compliance with point UL 2 § 65;

4. stall speed in various configurations;

5. loss of height from the beginning of the fall from direct flight until the resumption of level flight according to point UL 2 § 201.

6. If special procedures are required to start the engine in flight, these must also be listed.

7. Safe mounting, leveling and dismounting procedures that are expected to be performed by the pilot before and after flight and could cause unobserved damage must be reported.

8. Data on the function and operation of the rescue system must be provided.

 

SUBPART J - PROPELLERS

I. Design and manufacture

The manufacturer will suggest for which types, engine powers and installation method the propeller is intended.

UL 2 § 1917 Materials

The properties and durability of the materials used by the propeller manufacturer must:

1. be proven on the basis of experience or tests,
2. conform to specifications that safely establish that its strength and required properties agree with the proposed values.

UL 2 § 1919 Reliability

The design and manufacture must minimize the possibility of dangerous propeller conditions occurring between two inspections.

UL 2 § 1923 Regulation of propeller settings

The suitability of installing an adjustable or adjustable propeller must be consulted with the relevant responsible authority.

II. Strength tests of propellers

1. Tests of fixed wooden propellers prescribed in type tests

The new type of solid wooden propeller (monobloc) must undergo a strength test by spinning at a speed corresponding to 1,23 times the highest operating speed. The duration of the test is 5 minutes, the propeller must not show any damage or permanent deformation.

2. Tests of other types of propellers prescribed in type tests

a. The new type of propeller must be tested by overloading the hub and the root parts of the blades with a load corresponding to twice the value of the centrifugal force of the maximum permissible revolutions of the propeller. The load will act for min. 1 hour. A static load or a load by overturning to 1,4 maximum permitted revolutions is allowed, while the blade setting angle will be zero.
b. The new type of propeller must undergo a strength test by spinning at a speed corresponding to 1,23 times the highest operating speed for a period of 5 min. during operational adjustment of the angle of attack of the blade.
c. Depending on the design and type of materials used, other requirements (e.g. fatigue tests) will be determined by the responsible authority.

Note:
An individually made wooden propeller can be approved based on experience, if the composition and quality of the wood from which it is made can be seen.

An individually manufactured propeller made of composite materials must always be tested by spinning it to a speed corresponding to 1,1 times the highest operating speed. Exam duration 5 minutes.

The term individually manufactured means propellers that do not have a type test.

 

APPENDIX I. - RESCUE SYSTEMS

A. In general

1. For the installation of a certified rescue system (ZS), it is necessary to demonstrate that it fully meets the requirements of the regulation for rescue systems for ultralight aircraft.

2. The installation of a rescue system in a UL aircraft by its manufacturer or owner must be approved by the manufacturer of the type-certified ZS.

3. The structure and its attachment points must be described in the operational documentation of the UL aircraft.

B. Rescue system load

1. The structure between the attachment points of the rescue system's support ropes must be designed to withstand the dynamic impact that occurs when the rescue system is put into operation and corresponds to the values ​​specified by the manufacturer. Furthermore, it is required that the design and fastening of the seats, safety belts and subsequent structures up to the anchorage points of the vehicle must withstand the forces caused by the weight of the crew as a result of the dynamic shock from the vehicle.

2. Dynamic shock multiplied by a factor of 1,5 = safe load.

If the load-bearing ropes are fixed at several points of the load-bearing structure, then each individual fixing point must bear the load, which is defined as follows:

Main curtains are always front hinges, which must be dimensioned as follows:

– One main hinge – it must be sized for a safe load.
– More main hinges (usually 2) – each hanger must be rated for the following load: (safe load divided by the number of main hangers) multiplied by a factor of 1,33.

Rear curtains (stabilizing).

– The strength of each hinge – each individual fastening point must be dimensioned as follows: (safe load divided by the number of all fastening points, including the main ones) multiplied by a coefficient of 1,33.

Example of calculating the load from the rescue system:

Input data: flight weight 600 kg, speed V_D = 300 km/h, dynamic impact from ZS is 5 g, the aircraft has 4 hinges (2 front main and 2 rear stabilizers), basic safety factor = 1,5, additional safety factor = 1,33.

Safe load:

F_safety = F_dyn * 1,5 = (600 * 9,81 * 5) * 1,5 = 44 N

Main hinge load:

Rear suspension load:

3. When designing the ZS attachment points, it must be considered that the dynamic impact acts on the strength structure in all the following directions:

a. in a vertical plane from a direction parallel to the longitudinal axis of the aircraft backwards to a direction of 60° upwards, and
b. in the range of 30° to both sides from the axis of symmetry.

C. Construction of the rescue system

1. The attachment of the rescue system must be designed for a maximum multiple that corresponds to the specified flight and landing load cases, including the prescribed load during an emergency landing.

2. If the rescue system is located in front of the propeller, such a device must be installed to prevent interruption of the propeller support cables.

3. In the event of putting the rescue system into operation, the fastening and the surrounding strength structure must be able to absorb a possible recoil.

Warning: the amount of recoil can be thought of as a numerical value.

4. The device that activates the rescue system must be located so that it is easily accessible by the pilot and easy to use even under overload conditions.

5. The structure between the attachment points of the support cables and the seat, including the safety belts, must be able to absorb the shock when the rescue system is deployed according to Chapter B.

 

APPENDIX II. – TOWING GLIDES

Additional requirements for glider towing by ultralight aircraft

The following additional requirements apply to ultralight aircraft that can be used for towing gliders and for self-towing:

A. In general

1. Aerotow consists of a towed ultralight aircraft with a towing device and a towed glider.

2. As a rule, towing equipment consists of the following elements:

a. towing hitch,
b. disconnection device,
c. measuring device for measuring the critical temperature of the engine,
d. equipment for observing the towed glider in flight
e. safety tow rope

3. Gliders may only be towed by ultralight aircraft which are approved for towing and which comply with this approval.

4. Approval for towing will be granted if the applicant proves to the relevant UL technical inspector, either as part of the type or supplemental tests, that all the requirements listed here are met (for type-certified airplanes, the chief technical inspector decides). UL with type certification is not required for aerofoils of ultralight gliders, gliders with a technical certificate issued by LAA CR and hang gliders.

5. The approval for towing and the documents necessary for modification are stated in the airworthiness approval documents. Approval for towing is entered in the SFD Technical Certificate. The entry will be made by the Central Register of the LAA based on the request of the relevant technician inspector in the Registration Sheet (Remarks column). More detailed information about the towed aircraft is provided in the Flight Manual, or in the supplement to the Flight Manual.

B. Design and Construction

I. Tow rope switch

1. The hitch control must be located so that the hand that controls the throttle can also control this lever. The controller must be within easy reach and actuate with a "pull" without affecting the safe control of the tow aircraft.

2. The driver must be marked in yellow. a "Tow hitch" label must be placed near it. Its stroke should be at least 50 mm and may not exceed 120 mm.

3. The control between the lever and the towing hitch must operate smoothly.

4. The operating force to release the rope from the hanger must not exceed 200 N if the hanger is loaded with a force Q_name for load directions according to chapter E, paragraph 1.

5. The controller must be located in the cockpit in such a way that the force mentioned in the previous paragraph can be easily derived.

II. Indicator of critical engine running temperature

To check the critical temperature of the engine during towing, a critical temperature indicator must be placed in the pilot's field of vision with a warning of the maximum permissible temperature. The critical engine operating temperature is defined as the temperature at which the engine first reaches the maximum permissible value at the maximum continuous engine power.

III. A device for tracking a towed glider during towing

The pilot of the towing aircraft must be able to observe the towed glider without requiring special dexterity and without major body movements. The equipment used for this must show a calm and clear image of the towed glider. The glider must be visible in the entire range of the cone with an apex angle of 60° according to paragraph E.1.

IV. Towing rope and belay

Only non-metallic ropes (e.g. polyamide, polyester, etc.) may be used. The elongation of the towing rope under the permissible load may be no more than 30%. Rope joints should be protected against wear (abrasion) with a suitable covering (coating). The actual strength of the towing rope should not be higher than the rope load specified by the towing aircraft manufacturer. If a rope with a higher strength is used, it must have a belay with the maximum corresponding strength so as to ensure the protection of the towing aircraft and the glider. The rope for the aerial tug should be 40 m to 60 m long.

V. Towing hitch

1. The towing hitch must be rated for the loads in accordance with Chapter E. It must be installed so that the tow rope cannot collide with the UL control surfaces of the airplane in the load directions specified in Chapter E, paragraph 1.

2. Release of the rope (switching off) must be possible at the maximum permitted load in the entire range of the cone with an apex angle of 60° according to paragraph E.1.

3. The towing hitch must be adequately protected against contamination.

VI. Fuel pump

1. If the use of a fuel pump is necessary for trouble-free operation of the engine in accordance with point UL 2 § 955 paragraph 2., an emergency fuel pump must also be installed, which will immediately supply fuel to the engine in the event of a failure of the primary fuel pump. The drive of the emergency fuel pump must be independent of the drive of the primary fuel pump.

2. If the normal pump and the emergency pump are still in operation, an indicator or device capable of indicating a failure of one of the pumps shall be provided.

3. The operation of each fuel pump must not be influenced by the mode of operation of the engine in such a way that dangerous situations can arise, independently of the power of the engine or the operation of the engine or the operation of other fuel pumps.

C. Towed gliders

1. The gliders that are permitted to be towed are determined according to the weight and climb rate of the tow. Permissible take-off weights of a towed glider are determined by flight tests. The required speeds are determined according to Chapter D, Paragraph 5.

D. Flight characteristics of the aerofoil

1. tests with at least three representative types of gliders. At the same time, their maximum weight, aerodynamic characteristics, speed range and ground behavior should be combined in such a way as to obtain results lying on the safe side.

2. The take-off distance of a towed aircraft with an aerofoil at maximum weight from rest to a height of 15 m must be determined on a dry, flat, short-cut grass surface under normal conditions. The take-off length may be a maximum of 600 m.

Note:
The take-off length of the aerofoil listed in the Flight Manual shall be determined as the mean value of six proof flights.

3. The ascent time from unsticking to a height of 360 m above the starting point must not exceed 4 minutes, while

a. starting (maximum) power is used, a
b. the flaps are in the take-off position.

4. The best rate of climb of the aerofoil must be greater than 1,5 m/s after correction to the MSA values ​​at an altitude of 450 m, with:

a. using take-off (maximum) power,
b. with retracted landing gear (if the towing aircraft has retractable landing gear),
c. maximum take-off weight,
d. flaps in the position intended for ascending flight,
e. without exceeding all established temperature limits.

5. Minimum tow speed and tow speed for best climb must be determined by flight tests. The minimum speed of the aerofoil must not be lower than 1,3 V_S1 towing UL aircraft. The requirements of point UL 2 § 207, paragraph 1. and 2. also apply to the towbar.

6. The UL operating limits of the airplane must not be exceeded at any stage of the towing.

7. Engine operating limits must not be exceeded at any stage of the tow.

8. Take-off and towing must not require any extraordinary skills of the UL aircraft pilot or exceptionally favorable conditions. If the towed glider is out of the normal position in the airfoil inside LA 60° apex angle cone according to paragraph E.1., no extraordinary skill of the UL airplane pilot shall be required to restore its normal flight position.

9. The length of the towing rope must be determined. (40 to 60 m is recommended).

E. Fortress

1. It is assumed that the towboat is initially in steady level flight and a force of 500 N (if a more precise value is not available) acts on the towing line in the following directions:

a. backwards in the direction of the longitudinal axis of the trunk,
b. in a plane deviated from the horizontal plane of the trunk by 20° backwards and downwards,
c. in a plane deviated from the horizontal plane of the torso by 40° backward upwards,
d. in a plane deviated from the vertical plane passing through the longitudinal axis of the trunk 30° backwards to the side.

2. It is assumed that the aerobatic tow is in the same conditions as defined in paragraph 1 of this chapter and the load in the tow rope suddenly increases due to a shock to a value of 1,0 Q_name. The resulting rope load must be balanced by translational and rotational inertial forces.

3 Q_name is the maximum rated strength of the approved aerofoil towline belay.

Note:
The nominal strength should never be chosen lower than 200 daN, the recommended value is 300 daN.

4. The hitch must be designed for a working load of 1,5 Q_name, which act in the directions established in paragraph 1 of this chapter.

F. Operating Limitations and Data

1. The Flight Manual must contain the following information:

a. the maximum UL weight of the aircraft during towing,
b. maximum weight of towed glider,
c. the maximum rated strength of the tow rope fuse,
d. minimum air tow speed,
e. airfoil speed for best climb,
f. take-off lengths for at least 3 types of gliders that have been proven by tests. Furthermore, for example, other types of gliders can be listed, which have comparable corresponding properties with the tested types.

In addition, it must be stated how the take-off length is extended due to tall grass, raindrops or contamination of the airfoil (leading edge) and high air temperature.

2. A label must be placed next to the speedometer in the cabin

"Watch the speed in the aerofoil carefully!".

3. On the towed glider, a clearly visible label must be placed in the area of ​​the tow rope switch indicating the rated strength of the approved tow rope fuse.

4. Engine checks and periodic maintenance on UL airplanes used to tow gliders must be performed and documented by appropriate entries in the operational documentation to the extent prescribed by the engine manufacturer.

5. The requirements according to point UL 2 § 1585, paragraph 1. "Data on normal and emergency procedures" are also used for the tow truck, if they concern it.

 

APPENDIX III. – UNIT LOADS OF THE REAR PART OF THE AIRCRAFT

If more precise calculations are not performed, it is possible to calculate the load on individual parts of the structure according to the replacement methods listed below for conventional airplanes.

1. Loading of horizontal (vertical) tail surfaces

a. case of rudder deflection - maneuver

b. the case of a change in the angle of attack (yaw) – gust, damping

b₁ – depth of fixed part of tail surfaces (stabilizer, keel)
b₂ – depth of the moving part (rudder)
b_OP – depth of tail surface B_OP = b₁ + b₂

2. Aileron loading

b_KR – aileron depth
oo – axis of rotation

3. Flap loading

4. Aerodynamic brake (spoiler) load

V_VB – maximum speed with aerodynamic brakes extended
V_A – speed V_A from the Vn diagram

5. Load on the balancing surface (slot)

 

APPENDIX IV. – BASIC LANDING CASES

1. With stern wheel:

Landing assumption	horizontal landing	with a large angle of attack
The vertical component of the force at the center of gravity of the airplane	npr *G	npr *G
The horizontal component of the force at the center of gravity of the aircraft	0,25 * npr *G	0
The vertical component of the force on the wheels of the main landing gear	(npr – 0,667) *G	(npr – 0,667) * G * b/c
The horizontal component of the force on the wheels of the main landing gear	0,25 * npr *G	0
The vertical component of the force on the stern wheel	0	(npr – 0,667) * G * a/c
The horizontal component of the force on the stern	0	0

2. With bow wheel:

Landing assumption	horizontal landing		landing with a high angle of attack
	with inclined reactions	with the nose wheel above the ground
The vertical component of the force at the center of gravity of the airplane	npr *G	npr *G	npr *G
The horizontal component of the force at the center of gravity of the aircraft	0,25 * npr *G	0,25 * npr *G	0
The vertical component of the force on the wheels of the main landing gear	(npr – 0,667) * G * a1/c1	(npr – 0,667) *G	(npr – 0,667) *G
The horizontal component of the force on the wheels hl. chassis	0,25 * npr * G * a1/c1	0,25 * npr *G	0
Vertical component of force on the nose wheel	(npr – 0,667) * G * b1/c1	0	0
Horizontal component of force on the nose wheel	0,25 * npr * G * b1/c1	0	0

 

Undercarriage with stern wheel	Landing gear with nose wheel
Horizontal landing	Horizontal landing on three points
Landing with a high angle of attack	Horizontal landing with the nose wheel above the ground
	Landing with a high angle of attack

 

Lateral load on the wheels of the main landing gear:

Wheel load when braking:

Additional stern wheel load:

Additional nose wheel load: