Advanced Ultra-Light Aeroplanes Have a Type Design in Compliance With Set Standards
To be an advanced ultra-light aeroplane, an aircraft must meet the standards that are specified in the manual called “Design Standards for Advanced Ultra-Light Aeroplanes.”
Here is a link to that manual, from the “Light Aircraft Manufacturers Association of Canada.”
Use this link to apply to register ultralight aircraft in Canada.
You can use that link if you have a basic ultra-light aeroplane as well.
Laws Regarding Basic and Advanced Ultra-Light Aeroplanes
At National Aviation Registration, we’re committed to helping aircraft owners across Canada and internationally with all documentation requirements for their aircraft.
From initial registration and ownership transfers to address updates and more, we’re here to make the process easier. That’s true for those with basic ultra-light planes, advanced ultra-light planes, or any other kind of aircraft.
The following regulations may apply to your case. These excerpts are taken from “Design Standards for Advanced Ultra-Light Aeroplanes.”
Chapter A – General 1. Applicability (a) This publication contains standards for the design of Advanced Ultra-Light Aeroplanes. (b) Each person who manufactures an aeroplane or aeroplane kit for subsequent registration in the advanced ultra-light category shall demonstrate compliance with the applicable requirements of this publication. 2. Advanced Ultra-Light Aeroplane Category An Advanced Ultra-Light Aeroplane is an aeroplane which: (a) Is propeller driven; (b) Is designed to carry a maximum of two persons, including the pilot; (c) Has a maximum take-off mass, MTOmax, (weight, WTOmax) of: (i) 350 Kg (770 lb) for a single place aeroplane, or (ii) 560.0 Kg ( 1232 lb) for a two place aeroplane; (c) A maximum stalling speed in the landing configuration, VSO, at manufacturer’s recommended maximum take-off mass (weight) not exceeding 72 km/h (45 mph) (IAS); and (d) Is limited to non-aerobatic operations. Non-aerobatic operations include: (1) manoeuvres incident to normal flying (2) stalls and spins (if approved for type); (3) lazy eights, chandelles; and (4) steep turns, in which the angle of bank is not more than 60º 3. Minimum Useful Load Advanced ultra-light aeroplanes shall have a Minimum Useful Load, MU (WU) computed as follows: (a) For a single place aeroplane: MU = 80 + 0.3P, in kg; where P is the rated engine(s) power in kw; (WU = 175 + 0.5P, in lb; where P is the rated engine(s) power in BHP). (b) For a two place aeroplane: MU = 160 + 0.3P, in kg; where P is the rated engine(s) power in kw; (WU = 350 + 0.5P, in lb; where P is the rated engine(s) power in BHP)
- Maximum Empty Mass (Weight) The Maximum Empty Mass, MEmax, (Weight, WEmax) includes all operational equipment that is actually installed in the aeroplane. It includes the mass (weight) of the airframe, powerplant, required equipment, optional and specific equipment, fixed ballast, full engine coolant, hydraulic fluid, and the residual fuel and oil. Hence, the maximum empty mass (weight) = maximum take-off mass (weight) – minimum useful load. LAMAC 003 08 November 2004 DS 10141E Amendment 003 Amendment 003 2 Chapter B – Flight 5. Proof of Compliance Each of the following requirements shall be met at the most critical mass (weight) and CG configuration. Unless otherwise specified, the speed range from stall to VNE shall be considered. 6. Load Distribution Limits (a) Using comprehensive references, the following shall be determined: (1) the maximum empty mass (weight) and maximum take-off mass (weight) as defined in section 5. and 7., and a minimum flying weight; and (2) the empty CG, most forward and most rearward CG. Note: Standard occupant mass (weig
(b) Fixed and/or removable ballast may be used if properly installed and placarded. 7. Propeller Speed and Pitch Limits Propeller speed (RPM) and pitch shall not be allowed to exceed safe operating limits established by the manufacturer under normal conditions (i.e. maximum take-off RPM during take-off and 110% of maximum continuous RPM at closed throttle and VNE). 8. Performance, General All performance requirements apply in standard ICAO atmosphere and still air conditions. Speeds shall be given in indicated (IAS) and calibrated (CAS) airspeeds. 9. Stalling Speeds (a) Wing level stalling speeds shall be determined by flight test at a rate of speed decrease of 1.6 km/h/sec (1 mph/sec) or less, throttle closed, with maximum weight, and most unfavourable CG: (1) VS0: shall not exceed 72 km/h (45mph) (2) VS1: flaps retracted, shall not exceed 96.5 km/h (60 mph). (b) Level wing attitude and yaw control shall be possible down to VS0 or the speed at which the pitch control reaches the control stop. 10. Take-off With take-off at the maximum weight, full throttle, sea level, the following shall be measured: (a) Ground roll distance; and, (b) Distance to clear a 15.2 m (50 ft.) obstacle at 1.3 VS1. Note: The aeroplane configuration, including flap position, shall be specified. LAMAC 003 08 November 2004 DS 10141E Ame
- Climb With climb out at full throttle: (a) Best rate of climb (VY) shall exceed 93 m (300 ft) per minute; and, (b) Best angle of climb (VX) shall exceed 1/12. 12. Landing For landing with throttle closed and flaps extended, the following shall be determined: (a) Landing distance from 15.2 m (50 ft.) 1.3 VS0; and (b) Ground roll distance with reasonable braking if so equipped. 13. Balked Landing For a balked landing at 1.3 VS0 and flaps extended, the full throttle angle of climb shall exceed 1/30. 14. Controllability and Manoeuvrability (a) The aeroplane shall be safely controllable and manoeuvrable during take-off, climb, level flight (cruise), dive, approach and landing (power off and on, flaps retracted and extended) through the use of primary controls and normal displacements for the aircraft type. (b) Smooth transition between all flight conditions shall be possible without excessive pilot skills nor exceeding pilot force as shown in Figure 1.
Figure 1 (c) It shall be possible to trim the aeroplane at least for level cruise at an average weight and CG. LAMAC 003 08 November 2004 DS 10141E Amendment 003 Amendment 003 4 15. Longitudinal Control Longitudinal control shall allow: (a) Speed increase from 1.1 VSI to 1.5 VSI and from 1.1 VSO to VF in less than 3 seconds. This applies for both power-off and full power conditions. (b) Full control to be maintained when retracting and extending the flaps in the normal speed range; and (c) Stick forces per ‘g’ to steadily increase.
- Directional and Lateral Control (a) Reversing the roll from 30 degrees one wing low over to 30 degrees the other wing low shall be possible within 4 seconds at 1.3 VS0 (flaps extended and throttle idle) and at 1.2 VS1 (flaps retracted, throttle idle and full). (b) Rapid entry and recovery into/from yaw and roll shall not result in uncontrollable flight characteristics. (c) Where aircraft is so equipped, aileron and rudder forces shall not reverse with increased deflection. 17. Static Longitudinal Stability Longitudinal stability shall be positive from 1.2 VS to VNE at the most critical power setting and CG combination. 18. Static Directional and Lateral Stability (a) Directional and lateral stability and take-off and climb performance tests shall be performed to ensure the aeroplane complies with the requirements of this publication. (b) Directional and lateral stability is considered acceptable when the spiral stability of the aeroplane is neutral within the range specified in section 17. 19. Dynamic Stability Any short period oscillation shall be rapidly dampened with the controls free and the controls fixed. 20. Wings Level Stall It shall be possible to prevent more than 15 degrees of roll or yaw by normal use of the controls. 21. Turning Flight and Accelerated Stalls Stalls shall also be performed with power. After establishing a 30 degree co-ordinated turn, the turn shall be tightened until the stall. After the turning stall, level flight shall be regained without exceeding 60 degrees of roll. These stalls shall be performed with power on, flaps retracted and flaps extended. No excessive loss of altitude, nor spin tendency, nor speed build up shall be associated with the recovery
- Directional Stability and Control (a) Steering: Normal control inputs will achieve the desired steering results. In the case of aircraft equipped with rudders, pushing the right rudder pedal shall cause a turn to the right. (b) Ground handling shall not require special skills. No uncontrollable ground-looping tendency shall arise from 90 degrees of cross wind up to the maximum wind velocity selected by the applicant.
Chapter C – Structure 23. Loads (a) All requirements are specified in terms of limit loads. (b) Ultimate loads are limit loads multiplied by the factor of safety of section 24 (c) Loads shall be redistributed if the deformations affect them significantly. 24. Factor of Safety (a) The factor of safety is 1.5, except that it shall be increased to: (1) 2.0 x 1.5 = 3. on castings; (2) 1.2 x 1.5 = 1.8 on fittings; (3) 4.45 x 1.5 = 6.67 on control surface hinges; (4) 2.2 x 1.5 = 3.3 on push-pull control systems; and (5) 1.33 x 1.5 = 2. on cable control systems, seat belts and harness (b) The structure shall be designed as far as practicable, to avoid points of stress concentration where variable stresses above the fatigue limit are likely to occur in normal service. 25. Strength and Deformation (a) Limit loads shall not create permanent deformations nor large enough deformations which may interfere with safe operation. (b) The structure shall be able to support ultimate loads with a positive margin of safety (analysis), or without failure for at least three seconds (static tests). 26. Proof of Structure Each critical load requirement shall be investigated either by conservative analysis or tests or a combination of both. 27. Flight Loads (a) Appendix A of Chapter 523 of the Airworthiness Manual shall be used to determine the flight loads, except as noted in paragraphs 27(b) and (c). (b) Other design criteria may be used to determine the flight loads if their interpretation gives a level of safety equal to or exceeding Chapter 523 of the Airworthiness Manual. (c) For conventional designs, the simplified criteria of sections 28. to 34. may be used if they do not result in smaller load factors than the gust load factors of paragraph 27.(a), or in unrealistic values and the design falls within the limitations of Figure 2.
- Limit Load Factors The limit load factors shall be: (a) Positive: n = 4 (flaps retracted) and n = 2 (flaps extended); and (b) Negative: n = -2 (flaps retracted) and n = 0 (flaps extended). 31. Symmetrical Wing Loads 1. As a minimum, the following three conditions need investigation: (a) Point A normal load up = 4 x W tangential forward = W (b) Point D normal load up = 4 x W tangential rearward = W/5 (c) Point G normal down = -2 x W tangential forward = -2 x W/5 (d) Point F with flaps extended: normal up = 2 x W tangential forward = W
- Control System and Supporting Structure (a) The control system and supporting structures shall be designed for 125% hinge moments resulting from the surface load from section 35. but need not exceed the loads from the following pilot forces: (1) at the grip of the stick: (i) 445 N (100 lbs) in pitch (ii) 178 N (40 lbs) in roll limit loads: and (2) at the rudder pedals: 578 N (130 lbs) in yaw. (b) When dual controls are installed, the relevant system shall be designed for the pilots operating in opposition. (c) Control surface mass balance weights shall be designed for: (1) 24 ‘g’ ultimate normal to the surface; and (2) 12 ‘g’ ultimate fore and aft and parallel to the hinge line. (d) Right and left flaps shall be synchronized for symmetrical operation. (e) All primary controls shall have stops within the system to withstand the greater of pilot force, 125% surface loads, or ground gust loads. (f) The secondary controls shall be designed for the maximum forces a pilot is likely to apply in normal operation. 38. Ground Load Conditions (a) The basic landing conditions of Appendix C of Chapter 523 of the Airworthiness Manual are reproduced in Appendix B of this publication. (b) For advanced ultra-light aeroplanes the basic landing conditions of Appendix B of this publication are simplified as follows: L = ratio of the assumed wing lift to the aeroplane weight = 2/3; K = 0.25; n = nj + .67, load factor; and, nj = load factor on wheels, as defined in para (c) of this section.