testing generic

GENERIC FLIGHT TEST REPORT GUIDE

FOR

CERTIFICATION COMPLIANCE ASSESSMENT OF SIMPLE LIGHT FIXED WING AIRCRAFT

TABLE OF CONTENTS

SECTION PAGE

PART 1 AIRWORTHINESS FLIGHT TYPE REQUIREMENTS

1. Introduction 1-1

2. General Requirements 1-4

3. Control Systems 1-7

4. Pilot Compartment 1-9

5. Ventilation 1-13

6. Fuel System 1-14

7. Carburettor Air Heat Rise 1-15

8. Engine Cooling 1-16

9. Propeller Speed 1-18

10. Ground and Water Handling 1-19

11. Airspeed Calibration 1-20

12. Stall Speeds 1-22

13. Stall Characteristics 1-25

14. Controllability and Manoeuvrability 1-27

15. Trim 1-31

16. Stability 1-32

17. Spinning 1-37

18. Vibration and Buffeting 1-38

19. Flutter 1-39

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20. Take-Off Distance 1-40

21. Climb Performance 1-42

22. Glide Performance 1-44

23. Landing Distance 1-45

PART 2 FLIGHT TEST REPORT GUIDE

1. Introduction 2-1

2. Flight Test Log 2-4

3. Certification Data 2-5

4. Test Configurations 2-11

5. Equipment and Flight Operations 2-13

6. Ventilation 2-18

7. Powerplant 2-20

8. Ground and Water Handling 2-23

9. Airspeed Calibration 2-25

10. Stall Speeds 2-26

11. Stall Characteristics 2-28

12. Controllability and Manoeuvrability 2-37

13. Trim 2-Error! Bookmark not defined.

14. Stability 2-44

15. Spinning 2-51

16. Vibration and Buffeting 2-53

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17. Flutter 2-54

18. Take-Off Distance 2-55

19. Climb Performance 2-56

20. Glide Performance 2-60

21. Landing Distance 2-61

Annexes

A. Terms and Abbreviations

B. Cooling Climb Test Data

C. Airspeed System Calibration Test Data

D. Longitudinal Static Stability

E. Take-Off Performance Data

F. Climb Performance Data

G. Glide Performance Data

H. Landing Performance Data

Light Aircraft Flight Test Report Guide 1-1

PART 1 AIRWORTHINESS FLIGHT TYPE REQUIREMENTS

SECTION 1 - INTRODUCTION 1. Background

The aircraft type certification process, as applicable in Australia, is set out under Civil Aviation Safety Regulations (CASR) Part 21 and explained further at Advisory Circular (AC) 21-13(0) Australian Designed Aircraft Type Certification. A major phase in any aircraft development project is that involving the flight testing. AC 21-13(0) divides this phase into three supplementary segments the developmental test flying, the manufacturers test flying to demonstrate that the aircraft meets the requirements of the applicable airworthiness standard, and any further testing carried out by CASA to validate the compliance claims of the manufacturer. Each of these flight test segments should result in the production of a Flight Test Report (FTR). In addition, other aircraft production projects that do not necessarily progress ultimately to the certification stage, for example those undertaken by amateur-builders, should also involve the generation of reports as they progress through their flight test phases. While there is no designated or required format for any of these flight test reports this document provides a template, and some basic guidance, that may be adapted for the task. 2. Purpose

This document includes background information and guidance for planning or checking compliance with those parts of the basic airworthiness standards requiring flight tests and pilot judgements. It then presents a Flight Test Report Guide (FTRG) that can be used as the basis for producing either a general flight test report or one specifically aimed at satisfying the type certification process (CASR 21.035 refers). Notwithstanding, the information provided herein is neither mandatory nor regulatory in nature and does not constitute a regulation or order. It is intended as an optional ready reference for manufacturers, amateur-builders, test pilots (TPs), CASA evaluation engineers and authorised persons. 3. Scope

Flight test items of interest during the evaluation of small fixed-wing aircraft are covered. Since all such aircraft are to some extent unique some sections of the guide may not be applicable. Conversely some relevant information, for example multi-engine considerations, may not be included. Before commencing any flight test program, the manufacturer or amateur-builder is encouraged to consult the Authority or the authorised person to determine which sections are relevant to the particular aircraft. 4. Applicability

The FTRG is intended to be generic and adaptable. No attempt is made in the document to define small fixed-wing aircraft. The airworthiness standards referenced at paragraph 6 below give an indication of the category of aircraft to which the guide may be applicable however additional information will probably be relevant


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when considering those more sophisticated or capable aeroplanes that could be covered by those standards. The generic nature of this guide should be acknowledged and if it is to be used in association with a type certification project the applicable airworthiness standard and its relevant guidance material must be consulted to ensure the actual requirements are being addressed. The currency of the information in this guide, when compared with that in the applicable airworthiness standard, cannot be guaranteed. In addition, certification projects will almost certainly require the provision of supplementary or supporting information. 5. Contents

The document contains two parts. This first provides basic flight test guidance information that can be used as background to the actual FTRG which comprises the second. 6. Certification Standards

The information provided in Part 1 and the FTRG itself are meant to be generic. It does not relate specifically to any one certification airworthiness standard. However its basis lies in the following five standards:

a. Civil Aviation Orders Section 101.55 (CAO 101.55),

b. Civil Aviation Orders Section 101.28 (CAO 101.28),

c. British Civil Airworthiness Requirements Section S (BCAR-S),

d. European Aviation Safety Agency (EASA) Certification Specification for Very Light Aeroplanes (CS-VLA), and

e. Federal Aviation Regulations (of the USA) Part 23 (FAR 23).

7. References

The primary and best reference for flight test information relating to small aeroplanes is AC 23-8B Flight Test Guide for Certification of Part 23 Airplanes, published by the Federal Aviation Administration (FAA). AC 23-8B is the major source of the advisory and flight test technique information reproduced in this guide. FAA Advisory Circular 23-15A Small Airplane Certification Compliance Program provides advice on acceptable means of meeting FAR 23 requirements for simple light aircraft. CASA AC 21-40(0) Measurement of Airspeed in Light Aircraft Certification Requirements, provides information regarding the installation and calibration of airspeed measuring systems. For amateur builders guidance on developmental flight test programs (as opposed to certification flight testing) can be obtained from the FAA Amateur Built Aircraft Flight Testing Handbook (AC 90-89A). If this Handbook is used as the basis of a test plan then sufficient data should be gathered to generate an effective Flight Test Report.


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The Sports Aircraft Association of Australia (SAAA) also publishes a useful SAAA Flight Test Guide. A specific version of this FTRG, applicable to amateur-built aircraft constructed in accordance with CAO 101.28 is provided as the Flight Test Guide for Assessment of Amateur-Built Aircraft Accepted under an ABAA (Amateur-Built Aircraft Acceptance). CASA AC 21-47(0) Flight Test Safety provides general safety information as well as guidance regarding hazard analysis / risk management procedures. Terms and abbreviations used in this document are defined at Annex A.


SECTION 2 GENERAL REQUIREMENTS 1. Operating Limitations

Aircraft operating limitations should have been established. These limitations will include all those upon which the aircraft design is based and all those that are applicable to:

a. weight and centre of gravity (CG) limits,

b. loading,

c. powerplant,

d. airspeed, and

e. flight handling including aerobatic manoeuvres.

2. Weight and CG Limits

The maximum weight must have been established such that it is:

a. Not more than:

(1) the highest weight selected by the applicant, or

(2) the design maximum weight, which is the highest weight at which compliance with each applicable structural loading condition and each applicable flight requirement is shown.

b. Assuming a weight of 77 kg for each occupant of each seat, not less than the weight which results from the empty weight of the aeroplane, plus required minimum equipment, plus:

(1) the required minimum crew and fuel and oil to full tank capacity, or

(2) each seat occupied, oil to full tank capacity and fuel for one half hour operation at rated maximum continuous power (MCP).

The approved maximum weight should be included in the appropriate type data package. The ranges of weight and CG within which the aeroplane is to be safely operable must have been established and should be as demonstrated by the applicant. The CG range must not be less than that which corresponds to the weight of each occupant, varying between a minimum of 55 kg for the pilot alone up to the maximum placarded weight for a pilot and passengers, together with a variation in fuel contents from zero to full fuel. The placarded maximum occupant weight must not be less than 77 kg per person.


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The empty weight and corresponding CG must be determined by weighing the test aeroplane:

a. With:

(1) fixed ballast,

(2) required minimum equipment,

(3) unusable fuel, maximum oil and, where appropriate, engine coolant and hydraulic fluid.

b. Excluding:

(1) weight of occupants, and

(2) other readily removable items of load.

The condition of the aeroplane at the time of determining empty weight must be one that is well defined and easily repeated. The minimum weight (the lowest weight at which compliance with each applicable requirement has been shown) must have been established such that it is not more than the sum of:

a. the empty weight,

b. the weight of the minimum crew (assuming a weight of 77 kg for each crew member), and

c. the fuel necessary for half an hour of operation at rated maximum continuous power.

Each requirement must be met by test on the aircraft at the most adverse combination of weight and CG within the range of loading conditions for which certification was or is to be demonstrated. 3. Ground Tests

Prior to flight testing, the following ground tests should be conducted:

a. Measurement of:

(1) control circuit stiffness and stretch,

(2) control circuit friction,

(3) control cable tension of closed control circuits, and

(4) maximum deflection of control surfaces, wing flaps and their respective cockpit controls.

b. All ground functional and systems tests should be completed. FAA AC 90-89A provides excellent guidance on conducting these tests.


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4. Instrumentation

For test purposes the aircraft should be equipped with suitable instruments to conduct the required measurements and observations with appropriate accuracy. If reliable results cannot be obtained the Authority, or the authorised person, will require the installation of special test equipment. At an early stage in the program the accuracy of the instruments and their correction curves should be determined. Particular attention should be paid to the position error of the airspeed indicating system. The influence of the configuration of the aeroplane should also be accounted for in all calibrations. CASA AC 21-40(0) Measurement of Airspeed in Light Aircraft Certification Requirements, should be consulted. 5. Performance

Compliance with performance requirements must be shown at maximum weight for still air in standard atmosphere at sea-level conditions using engine power not in excess of the maximum declared for the engine type. 6. Test Pilots

The minimum qualification a pilot must hold to carry out the initial flight testing on an Australian aircraft is a Private Pilot Licence (PPL) with the appropriate endorsements. However, flight testing of aircraft, for either development, compliance demonstration or post construction purposes, is an exacting task and, while the regulations do not call for the TP to have any specific test flying qualifications, it is recommended that a pilot with at least some such knowledge and experience be engaged. Further guidance is contained in CASA AC 21-47(0) and FAA AC 90-89A. 7. Hazard Analysis / Risk Management

There are hazards involved with all flight testing. Some sequences (eg spinning, flutter) will involve elevated risk levels. This guide does not include specific risk management information. The user is strongly urged to conduct a detailed Hazard Analysis / Risk Management exercise as part of the test planning and the ongoing flight testing processes. Guidance is contained in CASA AC 21-47(0) or can be gleaned from the Internet using search terms such as flight test risk or hazard management. Project managers and pilots are encouraged to contact the CASA TP for further information or assistance.


SECTION 3 CONTROL SYSTEMS

1. Requirements

a. Each control must operate easily, smoothly and positively enough to allow proper performance of its function.

b. Each control system must have stops that positively limit the range of motion of each movable control surface and prevent over-centre locking tendencies.

c. Proper precautions must be taken to prevent inadvertent, improper, or abrupt trim operation. There must be markings near the trim control to indicate the direction of trim control movement relative to the aeroplane motion.

d. There must be means to indicate to the pilot the position of the trim device with respect to the range of adjustment. This means must be visible to the pilot and must be located and designed to prevent confusion.

e. If there is a device to lock the control system on the ground there must be means to;

(1) give unmistakable warning to the pilot when the lock is engaged, and

(2) prevent the lock from engaging in flight.

f. All control systems must be designed and installed to prevent jamming, chafing, and interference from baggage, passengers, loose objects, or freezing of moisture.

g. Each wing flap control must be designed so that, when the flap is placed in any position upon which compliance with the performance requirements is based, the flap will not move from that position except when the control is adjusted unless movement is demonstrated not to be hazardous.

h. The pilot forces and rate of movement of the wing flaps at any approved speed must not impair the operating safety of the aeroplane.

i. There must be means to indicate all flap positions upon which compliance with the performance requirements are based.

j. The wing flaps must be mechanically interconnected unless the aeroplane has safe flight characteristics with the wing flaps retracted on one side and extended on the other.

2. Explanation

ABAA Aircraft - Flight Test Report Guide 8

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This section requires an assessment of the aircraft control systems. Compliance with paragraphs 1.a to 1.f and paragraph 1.i can be shown by inspection. Compliance with paragraphs 1.g, 1.h and 1.j must be demonstrated through flight test.


SECTION 4 PILOT COMPARTMENT

1. Requirements

a. The design of the cockpit and the cabin shall be such as to give each occupant every reasonable chance of escaping serious injury in a crash.

b. Seats shall not have hard edges or protruding parts which in a crash could be in a position likely to cause serious head injuries to a person seated and wearing correctly adjusted restraint equipment.

c. Stowage shall be provided for the flight manual, if required, either in a container of the glove box type or in a special fixed container easily accessible to the pilot.

d. The scale graduations of airspeed indicators shall be in knots only with the scale as close as is practical to the periphery of the dial. The numbered graduations shall commence at a speed lower than the power-off indicated stalling speed of the aircraft in the landing configuration and at the minimum weight operationally possible.

e. The height scale of altimeters shall be graduated in feet.

f. The barometric subscale of sensitive altimeters shall include a calibration in millibars in increments not exceeding 2 millibars.

g. Magnetic compasses shall be corrected and calibrated in accordance with Airworthiness Bulletin (AWB) 34-8.

h. A landing gear position indicator is required if the aeroplane has retractable landing gear.

i. Each pilot, with her seat, safety harness and any adjustable controls correctly adjusted for normal flight, shall be able to;

(1) without interference produce full and unrestricted movement of each control which he or she may be required to operate in flight, both separately and with all practical combinations of movement of other controls, and

(2) at all positions of each control exert adequate control forces for the operation to be performed.

j. The cockpit and its equipment must allow each pilot to perform his or her duties without unreasonable concentration or fatigue.

k. Each cockpit must be designed so that;

(1) there is no glare or reflections that interfere with the pilots vision,

(2) the pilots field of view is sufficiently extensive, clear and undistorted for safe operation, and


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(3) rain does not unduly impair the field of view along the flight path during normal flight and landing.

l. Each cockpit control must be located to provide convenient operation and to prevent confusion and inadvertent operation.

m. In aeroplanes with dual controls it must be possible to operate the following secondary controls from each of the pilot seats;

(1) throttle lever,

(2) wing flaps,

(3) trim, and

(4) opening and jettisoning device for the canopy.

n. Secondary controls must maintain any desired position without requiring constant attention and must not tend to creep under loads or vibration.

o. Fuel valves must be provided to allow the pilot to rapidly shut off fuel to the engine in flight. The fuel valves must be designed to prevent inadvertent operation and allow the pilot to rapidly open each valve after it has been closed.

p. Each fuel valve must have positive stops or effective detents in the on and off positions.

q. All emergency controls must be coloured red.

r. The cockpit must be designed to allow the occupants to make a rapid and unimpeded escape in an emergency.

s. Where the airspeed limitations are marked on the airspeed indicator they shall be as indicated airspeed and the following colour code shall be used;

(1) for the never-exceed speed, VNE, a red radial line,

(2) for the caution range, a yellow arc extending from the red line specified in (1) to the upper limit of the green arc specified in (3),

(3) for the normal operating range, a green arc with the lower limit at the stalling speed or minimum steady flight speed , VS1, corresponding with the maximum take-off weight and landing gear and flaps retracted, and the upper limit at the maximum structural cruising speed or normal operating limit speed, VNO, and

(4) for the flap operating range, a white arc with the lower limit at stalling speed or minimum steady flight speed in the landing configuration, VS0, corresponding with maximum take-off weight, and the upper limit at the maximum flaps extended speed, VFE.


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t. Where the powerplant limitations are marked on the powerplant instruments, the following colour code shall be used;

(1) each maximum and, if applicable, minimum safe operating limit shall be marked with a red radial line,

(2) each normal operating range shall be marked with a green arc not extending beyond the maximum and minimum continuous safe operating limits,

(3) each take-off and precautionary range shall be marked with a yellow arc, and

(4) each engine speed range that is restricted because of excessive vibration shall be marked with a red arc.

u. The aircraft shall contain the following placards;

(1) the required operating limitations placard,

(2) except where aerobatic manoeuvres have been permitted, NO ACROBATIC MANOEUVRES (INCLUDING SPINS) PERMITTED,

(3) where the airspeed indicator is not marked with the colour code described above a placard listing the airspeed limitations,

(4) where the powerplant instruments are not marked with the colour code described above a placard listing the powerplant limitations,

(5) identification of the various functional positions of the controls of the fuel valves or cocks,

(6) on or adjacent to each fuel filler cap FUEL and the minimum fuel grade designation for the engine and the useable capacity of the fuel tank,

(7) the compass calibrations,

(8) loading instructions if necessary to ensure that in all conditions of operation the aircraft CG will remain within limits,

(9) NO SMOKING unless;

a) fuel tanks fitted in the cockpit or cabin are isolated by means of vapour and fuel proof enclosures,

b) the cockpit and cabin linings, floorings and furnishing materials are at least flame resistant, and

c) self-contained ashtrays are provided, and

(10) where the aircraft is to be operated under an Experimental Certificate, WARNING. THIS AIRCRAFT IS NOT REQUIRED


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TO COMPLY WITH THE SAFETY REGULATIONS FOR STANDARD AIRCRAFT. YOU FLY IN THIS AIRCRAFT AT YOUR OWN RISK.

v. Each marking and placard required shall be displayed in a conspicuous and appropriate position, shall be capable of being easily read and shall not be easily erased, disfigured or obscured.

w. Markings shall be placed on both the inside and the outside of each exit door, hatch or canopy, indicating the position of the opening handles with the locks fully engaged and also providing essential operating instructions for opening. (Note: Where opening is achieved simply by turning a handle a curved arrow pointing in the correct direction and the word OPEN will provide adequate instruction.)

x. Where the cockpit is enclosed, the opening system must be designed for simple and easy operation. It must function rapidly and be designed so that it can be operated by each occupant when strapped in and from outside the aircraft.

2. Explanation

This section requires an assessment of the aircraft cockpit. Compliance with all paragraphs can be shown by inspection.


SECTION 5 VENTILATION

1. Requirements

When there is an enclosed cockpit suitable ventilation must be provided under normal flying conditions. Carbon monoxide (CO) concentration must not exceed one part per 20 000 parts of air (ie 50 parts per million). 2. Explanation

The level of CO contamination must be determined under all normal operating conditions using an approved CO detector. 3. Procedures

The level of CO contamination must be measured in front of the pilots face and at the instrument panel during ground operations, climb, cruise and approach.


SECTION 6 FUEL SYSTEM

1. Requirement

The unusable fuel quantity for each tank must be established. It must not be less than the quantity at which the first evidence of malfunctioning occurs under the most adverse fuel flow conditions expected during take-off, climb, approach and landing. 2. Procedures

The unusable fuel quantity for each tank may be determined by ground tests that accurately simulate the following flight attitudes and conditions:

a. Level flight at Maximum Continuous Power (MCP).

b. Climb at MCP at the take-off safety speed with full rudder sideslip.

c. Glide at idle power with landing gear and flap extended at VFE.

d. Glide at idle power with landing gear and flap extended with full rudder sideslip at 1.3 VS0.

e. Any combination of the above or any other approved manoeuvre that may be critical to the design.


SECTION 7 CARBURETTOR AIR HEAT RISE

1. Requirements

The induction system of the installed engine must provide effective means to prevent and eliminate icing. Unless this can be achieved by other means it must be shown that in air free of visible moisture at a temperature of -10C the induction system preheater can provide a heat rise of 500C relative to the outside air temperature with the engine at 75 percent of MCP. 2. Explanation

Tests of engine induction system icing protection provisions are conducted to ensure the engine will operate throughout its flight power range. Experience has shown that it may be difficult to achieve the required heat rise on uncowled engines and/or on two-stroke engines. In such cases the Authority or authorised person should be approached for guidance. 3. Procedures

All tests should be conducted in air free from visible moisture. The temperature sensing probe should be installed in the induction system downstream of the heater and upstream of the venturi. It may be necessary to make a tapping into the intake duct. Heat rise requirements should be met at an outside air temperature (OAT) of -10C at an altitude where the engine can develop 75 percent MCP. If it is not possible to obtain these conditions tests should be conducted at the required power setting at the lowest OAT achievable and reduced to the required condition. At the test altitude stabilise the aircraft in level flight at 75% MCP with the carburettor heat control in the cold position. Allow all parameters to stabilise then record pressure altitude, OAT, RPM, manifold pressure and carburettor inlet air temperature. Apply full carburettor heat and allow temperatures to stabilise before recording the new carburettor inlet air temperature. Repeat this procedure two or three times to ensure consistent results. Air temperatures should be measured with calibrated temperature probes. All other quantities may be recorded from the standard aircraft instruments.


SECTION 8 - ENGINE COOLING

1. Requirements

The engine cooling provisions must be able to maintain the temperatures of engine components and engine fluids within the temperature limits specified by the engine manufacturer, or within the limits determined to be necessary by the aircraft manufacturer. Compliance with the cooling requirements must be shown under all likely operating conditions. 2. Explanation

Compliance with the engine cooling requirements must be demonstrated by verifying that there is sufficient cooling to remain within temperature limits during climb at the take-off safety speed. The hottest cylinder should be determined before commencing climb tests if only one cylinder is to be monitored. Test data can be corrected to sea level 300C conditions with an assumed lapse rate of 20C per 1000 ft altitude. The correction formula (presented below) is necessarily conservative, so it is in the builders interest to test in the highest possible ambient temperatures to minimise the corrections required. However, temperatures should not exceed ISA + 230C. 3. Procedures

Tests must be conducted in air free of visible moisture at the maximum take-off weight and most forward CG. At the lowest practical altitude establish level flight at 75% MCP until temperatures stabilise. Record cooling data. Apply take-off power and climb at the take-off safety speed for one minute then reduce power to MCP and continue climbing at the same speed. The climb should continue to an altitude where the temperatures stabilise, plus 500 ft, or to the maximum operating altitude. The following cooling data should be recorded at one minute intervals throughout the test:

a. Time,

b. Pressure Altitude,

c. Outside Air Temperature,

d. Hottest Cylinder Head / Coolant / Exhaust Gas Temperature (selected depending on the limits specified by the engine manufacturer),

e. Engine Oil Inlet Temperature,


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f. Engine RPM, and

g. Indicated Airspeed.

Temperatures should be measured with calibrated temperature probes. All other data may be recorded from the standard aircraft instruments. Temperatures should be corrected to the Standard Hot Day Conditions using the following formula:

Corrected Temperature = Measured Temperature + (38 0.002 HP OAT) Example: Measured Cylinder Head Temperature = 2000C

Measured oil Inlet Temperature = 1500C Pressure Altitude (1013.2) = 3000 ft Outside Air Temperature = 100C Corrected CHT = 200 + (38 0.002*3000 10) = 2220C Corrected Oil Inlet Temperature = 150 + (38 0.002*3000 10) = 1720C


SECTION 9 PROPELLER SPEED

1. Requirements

The propeller speed and pitch must be limited to values that ensure safe operation under normal operating conditions. During take-off and climb at the take-off safety speed the propeller must limit the engine RPM, at full throttle, to a value not greater than the maximum allowable RPM. During a glide at VNE, with the throttle closed or the engine stopped, the propeller must not allow the engine RPM to exceed 110% of the maximum engine or propeller RPM, whichever is lower. 2. Procedures

The wording of the requirements sufficiently describes the tests required to show compliance.


SECTION 10 GROUND AND WATER HANDLING

1. Requirements

Landplanes must have no uncontrollable tendency to nose over under any reasonably expected operating conditions, including rebound during take-off or landing. Seaplanes or amphibians must have no dangerous or uncontrollable porpoising characteristics at any normal operating speed on the water. The ability to take-off and land safely in crosswinds should be investigated. There must be no uncontrollable ground or water-looping tendencies in 90 degree crosswinds, up to the demonstrated limit. This limit must be shown at any speed at which the aircraft may be operated on the ground or water. Advice and limitations on operations in crosswinds based on the results of these tests shall be provided in the flight manual. Landplanes must be satisfactorily controllable in power-off landings at normal landing speed, without using brakes or engine power to maintain a straight path. The shock-absorbing mechanism may not damage the structure of the aircraft when it is taxied on the roughest ground that may be reasonably expected in normal operations. 2. Explanation

The longitudinal characteristics and control requirements are self explanatory. The highest 90 degree crosswind component satisfactorily tested should be put in the flight manual as performance information. The power-off landing requirement need only be met in zero crosswind. 3. Procedures

With the most adverse weight and CG combination for ground or water handling, the aircraft should be taxied at low and high speeds upwind, downwind and crosswind. A series of crosswind and into-wind take-offs and landings should be made in all configurations proposed as cleared take-off and landing configurations. Landings should be made power-off for the into-wind cases. Since it is not always possible to take-off or land in a direct crosswind it is acceptable to test in crosswinds of other than 90 degrees provided the crosswind component is accurately measured. The determination of compliance is primarily a qualitative one. However, wind readings (velocity and direction) should be taken at various points along the runway to determine that the minimum 90 degree component has been tested. Landplanes should be operated from all types of runways applicable.


SECTION 11 AIRSPEED CALIBRATION

1. Requirement

The aircrafts airspeed indicating system (the ships system) must be calibrated in flight to determine the system error. Any other airspeed measurement system, independent of the ships system, used during flight testing must also be calibrated. 2. Explanation

The airspeed indicating systems must be calibrated to enable accurate definition of the stall and limiting airspeeds. Airspeed calibration and stall speed determination should be accomplished at an early stage in the flight test program because many of the other test airspeeds are functions of the calibrated stall speed. Airspeed measurement system calibration is discussed in detail at CASA AC 21-40(0) Measurement of Airspeed in Light Aircraft Certification Requirements. 3. Procedures

A number of test techniques are available for airspeed system calibration. Those applicable to small aircraft are described in AC 21-40(0). The simplest techniques are the speed course method or the GPS method which basically require a comparison of indicated airspeed with an accurately derived ground speed. More precise results may be obtained using dedicated test equipment such as a pilot-static boom or a trailing static source. For the simple speed course method the following points should be noted:

a. The true airspeed of the aircraft can be determined by timing the aircraft flying over a known distance marked on the ground. The speed course should be flat with its length dependent on the range of speeds to be measured (a surveyed runway of approximately 1500 metres would be suitable for aircraft with cruise speeds in the 100 KIAS range).

b. The aircraft should be flown along the course at constant altitude and constant indicated airspeed. A height of between 200 ft and 500 ft is recommended.

c. To allow for wind effects, two runs at the same airspeed in opposite directions are required. The final applicable groundspeed being the average of the speeds on the two runs. The aircraft should be allowed to drift with any crosswind, ie fly the aircraft on a constant heading parallel to the speed course.

d. Calibrations must be conducted in the take-off, cruise and landing configurations at speeds between a minimum speed comfortably above the relevant stall speed, at least 1.2 VS0, and the maximum level flight speed. At least five test points should be covered in each speed range.


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e. All testing should be conducted in smooth stable air with the aircraft loaded to the maximum take-off weight. Testing in the calm conditions of an early morning is recommended.


SECTION 12 STALL SPEEDS

1. Requirements

The stalling speed, VS0, at maximum take-off weight shall not exceed:

a. 61 KCAS in the case of an aeroplane with a type certificated engine, or

b. 55 KCAS in any other case1.

The stalling speed, VS0, or minimum steady flight speed, in KCAS, should be determined with;

a. engine(s) idling with throttle(s) closed,

b. propeller(s) in the take-off position,

c. landing gear extended,

d. wing flaps in the landing position,

e. most forward CG, and

f. weight used when VS0 is being used as a factor to determine compliance with a particular standard.

The stalling speed, VS1, or minimum steady flight speed, in KCAS, should be determined with;

a. engine(s) idling with throttle(s) closed,

b. propeller(s) in the take-off position,

c. aeroplane in the configuration relative to the condition in which VS1 is being used,

d. most forward CG, and

e. weight used when VS1 is being used as a factor to determine compliance with a particular standard.

VS0 and VS1 should be established by flight test in accordance with the following procedures:

a. The aeroplane should trimmed, power off, at 1.5 VS or the minimum trim speed, whichever is higher, and

b. then slowed to approximately 10 knots above the stall from whence the airspeed should be reduced with the elevator control at a rate of one knot per second or less until the stall occurs or the control reaches the stop.

1 Different speeds are applicable for different categories of aircraft. Check the applicable airworthiness standard.


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2. Explanation

The aircraft is considered to be stalled when either of the following conditions occurs:

a. an uncontrollable downward pitching motion,

b. a downward pitching motion resulting from the activation of some device (eg a stick pusher), or

c. the longitudinal control reaches the aft stop.

It is important to use a standard test technique to determine the stall speeds accurately as they are used as a reference for many other performance and handling requirements. Aircraft airspeed indicators can be unreliable in the stall speed regions. All performance stall speeds are based on calibrated airspeed (CAS) and for accurate results an independent flight test airspeed measurement system can be used. Details are provided at AC 21-40(0) Measurement of Airspeed in Light Aircraft Certification Requirements. Aircraft which have a sloping forward limit to their weight / CG diagram should be evaluated at the most forward CG regardless of weight (the forward regardless point) as well as the most forward CG at maximum take-off weight. The higher of the stall speeds produced at either of these configurations will be taken as the stall speed. FAA AC 23-8B provides detailed information regarding stall speed measurement. 3. Procedures

The aircraft should be trimmed, power off, at 1.5 times the anticipated stall speed or at the minimum trim speed, whichever is greater. The aircraft should then be slowed to about 10 knots above the stall and from there the speed should be reduced at a rate of one knot per second or less until the stall occurs or the longitudinal control reaches its stop. If a calibrated flight test airspeed system is being used both the stalling calibrated airspeed and the indicated airspeed, using the production airspeed system installed in the aircraft, should be noted. Where exact determination of stalling speed is required, a number of stalls should be conducted during which the entry rate is varied to bracket one knot per second. The resultant stall speeds can then be plotted against entry rates to determine the one knot per second value. The time history of all stall flight tests should be recorded so the actual weight of the aircraft at the time of each test can be determined. Where an aerodynamic stall2 has

2 Not applicable if the stall is defined by the longitudinal control reaching the stop i.e. a minimum steady flight speed.


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been demonstrated the test weight stall speeds can be corrected to the maximum take-off weight stall speed using the following formula:

VS = VST * WS/WT where VS = Corrected Stall Speed (CAS) VST = Test Stall Speed (CAS) WS = Maximum Take-Off Weight WT = Test Weight


SECTION 13 STALL CHARACTERISTICS

1. Requirements Wings Level Stall

The behaviour of the aircraft during stalling from a wings level attitude must be investigated at the forward and aft CG limits. Priority should be given to the aft CG case and testing should be conducted at both maximum and minimum weights. The light loading case may be critical in aeroplanes with high thrust to weight ratios. Stall demonstrations should be conducted by reducing the speed at no more than one knot per second from straight and level flight until either a stall results as evidenced by a downward pitching motion, or a downward pitching and rolling motion not immediately controllable, or until the control reaches the stop. It must be possible to produce and correct roll and yaw by unreversed use of the controls until the stall occurs. It must be possible to prevent more than 150 of roll3 by normal use of the controls during recovery. There must be no tendency to spin. The loss of altitude from the beginning of the stall until regaining level flight by applying normal procedures, and the maximum pitch attitude below the horizon, must be determined. Stalls must be demonstrated under the following conditions:

a. Take-off, cruise and landing configurations.

b. Initial trim speed at 1.5 VS.

c. Power at idle and 75% MCP.

2. Requirements Turning Flight Stalls4

When stalling during a coordinated 300 banked turn it must be possible to regain normal level flight without encountering uncontrollable rolling or spinning tendencies. The roll will be considered to be uncontrollable if the aeroplane rolls more than a further 300 into the turn or more than a total of 600 out of the turn. The loss of altitude from the beginning of the stall until regaining level flight by applying normal procedures must be determined. Turning flight stalls should be demonstrated under the same conditions laid down for wings level stalling. 3. Requirements Stall Warning

There must be clear and distinctive stall warning with wing flaps and landing gear in any normal position when approaching the stall in both straight and turning flight. The stall warning must occur sufficiently in advance of the stall to provide the pilot adequate forewarning. 3 200 for aircraft certificated to BCAR-S 4 Aircraft certificated to FAR 23 or CS-VLA must also meet accelerated stall handling requirements.


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The stall warning may be furnished either through inherent aerodynamic qualities of the aeroplane or by a device that will give clearly distinguishable indications under the expected conditions of flight. However, a visual stall warning device that requires the attention of the crew within the cockpit is not acceptable in itself. 4. Explanation

It should be possible to achieve the stall conditions described above, and the subsequent recoveries, without an exceptional degree of pilot skill, alertness or strength. The requirement for unreversed use of the controls implies the aircraft will maintain positive stability throughout the sequence. A lightening in the longitudinal control force at the stall is acceptable however a push force, either as a transient to prevent pitch-up or as a steady push, is not. During recovery from the stall power should not be changed until flying control is regained. This is interpreted to mean not before a speed of 1.2 VS is attained. Successful investigation of the stall handling requirements will necessitate construction and completion of a matrix of data points covering all aircraft configurations, power-on and power-off, for both straight and turning flight. FAA AC 23-8B provides detailed information regarding investigation of an aircrafts stall characteristics. 5. Procedures

All tests investigating stall characteristics should be commenced with the aircraft trimmed at 1.5 VS or the minimum trim speed if greater. During entry to and recovery from the stall the following quantitative data should be recorded:

a. stall warning speed if stall warning is required,

b. stall speed,

c. pitch attitude changes,

d. roll attitude changes, and

e. altitude loss.

The following qualitative determinations should be made:

a. stick force curve remains positive, and

b. controls remain effective.

A sufficient number of stalls should be made in all configurations so as to produce repeatable data.


SECTION 14 CONTROLLABILITY AND MANOEUVRABILITY

1. General Requirements

The aircraft must be safely controllable and manoeuvrable during:

a. take-off at maximum take-off power,

b. climb,

c. level flight,

d. descent,

e. landing, with both power on and power off, and

f. in the event of a sudden engine failure.

It must be possible to make a smooth transition from one flight condition to another (including turns and slips) without exceptional piloting skill, strength or alertness and without danger of exceeding the limit load factor. This must be demonstrated under any probable operating condition and with the engine(s) running at all allowable power settings. The effects of power changes and sudden engine failure must also be considered. Modest departures from any recommended operating techniques must not cause unsafe flight conditions. Any unusual flying characteristics observed during flight testing must be evaluated. Any significant variations in flight characteristics caused by rain should be determined with the engine running at all allowable power settings. If marginal conditions exist with regard to pilot effort, limits should be shown by quantitative test and must not exceed those given in the following table: Values in pounds force applied to

the relevant control Pitch Roll Yaw

(a) For temporary application: Stick------------------------------------ 60 30 --------------- Wheel (Two hands on rim)---------- Wheel (One hand on rim)-----------

75 50

50 25

---------------- ----------------

Rudder Pedal-------------------------- ----------------- ----------------- 150 (b) For prolonged application 10 5 20 2. General Explanation

The phrase exceptional pilot skill, strength or alertness requires highly qualitative judgements on the part of the test pilot. These judgements should be based on the TPs estimate of the skill and experience of the pilots who will normally fly the type of aircraft under consideration. Exceptional alertness or strength requires additional judgement factors when control forces are deemed to be marginal or when a


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condition exists which requires rapid recognition and reaction to be coped with successfully. Temporary application, as specified in the table, may be defined as the period of time necessary to perform the required pilot actions to relieve the forces, such as trimming or changing the power setting. Prolonged control forces would be for some condition that could not be trimmed out, such as a forward CG landing. The time of application would be for the final approach only. If the aircraft could be flown in trim to that point. Controllability is the ability of the pilot, through proper manipulation of the controls, to establish and maintain the attitude of the aircraft with respect to its flight path. The design of the aircraft should make it possible to control the attitude about the longitudinal, lateral and directional axes. Controllability should be delineated as satisfactory or unsatisfactory. Unsatisfactory controllability would exist if the test pilot finds it to be so inadequate that a dangerous condition might occur. Such characteristics would be unacceptable for showing compliance with the regulations. Manoeuvrability is the ability of the pilot, through proper manipulation of the controls, to alter the direction of the flight path of the aircraft. In order to accomplish this, the aircraft must be controllable, since a change about at least one of the axes is necessary to change a direction of flight. Manoeuvrability is so closely related to controllability as to make them inseparable when considering any real motion of the aircraft. It is also similarly largely qualitative in its nature and should be evaluated in the same manner as controllability. 3. General Procedures

A qualitative determination of the controllability and manoeuvrability characteristics of the aircraft by the test pilot will suffice unless control force limits are considered marginal. In this case, force gauges or instrumentation should be used to measure the forces at each affected control while flying through the required manoeuvres. 4. Requirements Longitudinal Control

With the aircraft as nearly as possible in trim at 1.3 VS1, it must be possible at speeds below the trim speed, down to VS1, to pitch the nose downward so that a speed equal to 1.3 VS1 can be achieved promptly. This must be achievable with the aircraft in all possible configurations and at all engine power settings. It must be possible to lower the nose to maintain a safe flying speed when the engine power is suddenly reduced from take-off power to idle while climbing at the take-off safety speed. It must be possible, throughout the appropriate flight envelope, to change the configuration (landing gear, wing flaps, etc) without control of the aircraft requiring exceptional piloting skill and without exceeding the control forces defined above. It must be possible to raise the nose at VNE at all permitted CG positions and engine powers.


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The pitch control forces during turns or when recovering from manoeuvres must be such that at a constant speed an increase in load factor is associated with an increase in control force. For aircraft certificated to FAR 23 there is also a requirement to demonstrate that the aircraft is safely controllable and able to establish a zero rate of descent at an attitude suitable for a controlled landing without the use of the primary longitudinal control system. 5. Explanation Longitudinal Control

This area requires a series of manoeuvres be conducted to determine the longitudinal controllability during push-overs from low speed, flap extension and retraction, and during speed and power variations. The prime determinations to be made by the test pilot are whether or not there is sufficient elevator power to perform the required manoeuvres and that control forces are not excessive. 6. Procedures Longitudinal Control

The wording of the requirements describes the manoeuvres needed. Special instrumentation should not be required since most assessment is qualitative. However, longitudinal control forces should be measured if the forces are considered marginal or excessive. 7. Requirements Lateral and Directional Control

Using an appropriate combination of controls it must be possible to roll the aeroplane from a steady 30 degree banked turn through an angle of 60 degrees, so as to reverse the direction of turn, within five seconds from initiation of roll with:

a. flaps in the take-off position,

b. landing gear retracted,

c. maximum take-off power, and

d. the aeroplane trimmed at 1.2 VS1, or as nearly as possible in trim for straight flight.

Using an appropriate combination of controls it must be possible to roll the aeroplane from a steady 30 degree banked turn through an angle of 60 degrees, so as to reverse the direction of turn, within four seconds from initiation of roll with:

a. flaps extended,

b. landing gear extended,

c. engine operating at idle power and engine operating at power for level flight, and

d. the aeroplane trimmed at 1.2 VS1.


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For aircraft certificated to FAR 23 there is also a requirement to demonstrate that the aircraft is safely controllable without the use of the primary lateral control system in any all-engine configuration and at any speed or altitude within the approved operating envelope. 8. Explanation Lateral and Directional Control

Meeting this requirement should ensure enough lateral and directional control to provide an acceptable level of manoeuvrability in all phases of flight. 9. Procedures Lateral and Directional Control

The wording of the requirements describes the manoeuvres needed. Control forces should remain acceptable. 10. Requirements Elevator Control Forces in Manoeuvres

The elevator control forces during turns or when manoeuvring must be such that an increase in control force is needed to cause an increase in load factor. 11. Explanation Elevator Control Forces in Manoeuvres

The positive stick force per g levels in a cruise configuration must be of sufficient magnitude to prevent the pilot from inadvertently over-stressing the aeroplane during manoeuvring flight. FAA AC 23-8B provides detailed information. 12. Procedures Elevator Control Forces in Manoeuvres

Compliance may be demonstrated by measuring the normal acceleration and associated elevator stick force in a turn while maintaining the initial level flight trim speed. The local value of control force gradient should not be less than 3 lbs/g for stick-controlled aircraft or 4 lbs/g for those with control wheels.


SECTION 15 TRIM

1. Requirements

In level flight at 0.9 VH or VC (whichever is lower) the aeroplane must remain in trimmed condition around the roll and yaw axes with the respective controls free. (VH is the maximum speed in level flight with maximum continuous power set.) The aeroplane must maintain longitudinal trim in level flight at any speed from 1.4 VS1 to 0.9 VH or VC (whichever is lower). The aeroplane must maintain longitudinal trim during:

a. a climb with maximum continuous power at the best rate of climb speed, VY, with landing gear and wing flaps retracted, and

b. a descent with idle power at a speed of 1.3 VS1 with landing gear extended and wing flaps in the landing position.

For aircraft certificated to FAR 23 there is also a requirement to demonstrate that the aircraft is safely controllable following any probable powered trim system runaway that might be reasonably expected in service. 2. Explanation

The trim requirements ensure the aircraft will not require exceptional pilot skill, strength or alertness to maintain a steady flight condition. The tests require the aircraft to be trimmed for hands-off flight during the conditions specified. 3. Procedures

If installed, trim actuator travel limits should be set to the minimum allowable. Trim tests should be conducted in smooth air. Tests requiring the use of maximum continuous power should be conducted at as low an altitude as practical to ensure the required power is attained. Trim tests should be conducted at the most critical combinations of weight and CG. Forward CG is usually critical at slow speeds and aft CG critical at high speeds.


SECTION 16 STABILITY

1. General

The aircraft must be inherently stable and must show suitable stability and control feel under any conditions normally encountered in service. 2. Requirements - Static Longitudinal Stability

Under the conditions specified below, and with the aircraft trimmed as indicated, the elevator control forces and the friction within the control system must have the following characteristics:

a. A pull must be required to obtain and maintain speeds below the specified trim speed and a push required to obtain and maintain speeds above the specified trim speed.

b. The airspeed must return to within plus or minus 10 percent of the original trim speed when the control force is slowly released from any speed within the specified range.

c. The stick force must vary with speed so that any substantial speed change results in a force clearly perceptible to the pilot.

Climb Condition. The stick force curve must have a stable slope at speeds between the minimum speed for steady unstalled flight and the trim speed plus 20 knots or the flap limiting speed with;

a. flaps in the climb position,

b. landing gear retraced (if applicable),

c. maximum continuous power, and

d. the aircraft trimmed at 1.4 VS1.

Cruise Condition. The stick force curve must have a stable slope at speeds between 1.3 VS1 and VNE with;

a. flaps retracted,

b. landing gear retracted (if applicable),

c. maximum continuous power, and

d. the aircraft trimmed for level flight.

Approach Condition. The stick force curve must have a stable slope at speeds between the minimum speed for steady unstalled flight and the trim speed plus 20 knots (or the maximum flap extended speed if lower) with;

a. flaps in the landing position,


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b. landing gear extended,

c. normal approach power and power off (idle), and

d. the aircraft trimmed at the recommended approach speed.

If the aircraft does not have a longitudinal trim system the trim speeds for the climb, cruise and approach conditions should be those used when determining stall speeds (Section 12). 3. Explanation Static Longitudinal Stability

The requirement for the free return speed to be within 10% of the original trim speed effectively limits the amount of control friction that will be acceptable. For stability testing control cable tensions should be adjusted to the maximum allowable. The stable slope requirement requires judgement on the part of the test pilot as to whether or not the slope of the stick force versus airspeed curve is sufficiently steep to allow safe operation of the aircraft. 4. Procedures Static Longitudinal Stability

The aircraft should be trimmed in smooth air for the conditions required. After observing trim speed, apply a slight pull force and stabilise at a lower speed. Note the new airspeed and the required pull force. Continue this process in increments of five to ten knots, depending on the speed range being investigated, until reaching the specified minimum speed. The pull force should then be gradually relaxed to allow the aircraft to return toward the trim speed and zero stick force. Depending on the amount of friction in the control system the eventual speed at which the aircraft stabilises will be somewhat less than the original trim speed. The new speed, called the free return speed, must be within 10% of the original trim speed. Starting again at the trim speed push forces should be applied gradually in the same manner as described above for speeds in five or ten knot increments up to the specified maximum speed. The push force should then be smoothly released and, as before, the free return speed should be within 10% of the original trim speed. Tests should be conducted at the critical combination of weight and CG. Aft CG is normally critical. Both the light and heavy weight conditions should be checked. Force measurements can be made with a hand-held force gauge, fish scales or through electronic means, and plotted against calibrated airspeed to assess compliance. Test data should be obtained within a reasonable band of the trim point, no more than +/- 2000 ft. 5. Requirements Static Directional and Lateral Stability

The static directional stability, as shown by the tendency to recover from a skid rudder free, must be positive in the take-off, cruise and landing configurations. This must be demonstrated at all engine powers and at speeds from 1.3 VS1 up to the maximum allowable for the configuration.


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The static lateral stability, as shown by the tendency to raise the low wing from a sideslip, must be positive in the take-off, cruise and landing configurations. This must be demonstrated at engine powers up to 75% MCP at speeds above 1.3 VS1 and up to the maximum allowable for the configuration. The static lateral stability may not be negative at 1.3 VS1. In straight steady sideslips at 1.3 VS1, in the take-off, cruise and landing configurations, at up to 50% MCP, aileron and rudder forces and displacements must increase with sideslip angle. At large angles of skid, up to that at which full rudder is used or the rudder force exceeds 150 lbf, rudder forces may lighten but they must not reverse. This must be shown at all speeds between 1.3 VS1 and VA or the limiting speed for the configuration. 6. Explanation Static Directional and Lateral Stability

The requirements of this section demonstrate that the aircraft has positive directional and lateral stability and verify the absence of rudder lock. The skid angle required to assess compliance with the basic static directional stability requirements outlined above should be the maximum skid angle expected in service. This judgement should be based on aircraft manoeuvrability and control forces. The sideslip angle required to assess compliance with the basic static lateral stability should be the angle required to maintain a steady heading with a bank angle of 10 degrees. If the aircraft cannot maintain a steady heading with 10 degrees of bank applied then the heading may be allowed to drift. 7. Procedures Static Directional and Lateral Stability

Testing should be conducted at the highest altitude at which the required engine power can be achieved. The aircraft should be loaded to the aft CG limit and maximum take-off weight. Directional Stability. With the aircraft in the desired configuration and stabilised at the trim speed, slowly yaw the aircraft in one direction keeping the wings level with the lateral control. When the rudder is released the aircraft should tend to return to straight flight. The test should be repeated yawing the aircraft in the opposite direction. The amount of yaw should be appropriate to aircraft type. Rudder Lock. Continue to increase the rudder deflection beyond that used above until full deflection or the rudder force limit is reached. In this region rudder forces may lighten but may not reverse. Lateral Stability. With the aircraft in the desired configuration and stabilised at the trim speed conduct a sideslip by maintaining a steady heading with the rudder while banking at least 10 degrees with the ailerons. When the ailerons are released the low wing should tend to return to level. The pilot should not assist the ailerons during this demonstration but should hold full rudder (either up to the deflection limit or to the force limit, whichever occurs first). Repeat in the opposite direction.


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8. Requirements Dynamic Stability

Any short period oscillation occurring between the stalling speed and the maximum allowable speed appropriate to the configuration must be heavily damped with the controls;

a. fixed, and

b. free.

This requirement must be met at all allowable engine powers. 9. Explanation Longitudinal Dynamic Stability

Without instrumentation the short period oscillation requires a qualitative evaluation. The short period mode is the first response experienced after disturbing the aeroplane from its trimmed condition with the elevator control. It involves a succession of pitch acceleration, pitch rate and pitch attitude changes that occur so rapidly the airspeed does not change significantly. Qualitative evaluation of the short period mode should reveal it to be deadbeat or, if perceptible at all, damped to no more than one overshoot. If damping is any less a flight with appropriate instrumentation installed may be necessary. The motion should be damped within two cycles after input. If the disturbance from trim conditions is sustained long enough for the airspeed to change significantly, and if the pitch attitude excursions are not constrained by the pilot, the long period (or phugoid) oscillation will be excited with large but relatively slow changes in pitch attitude, airspeed and altitude. FAA AC 23-8B provides further information regarding dynamic stability. 10. Procedures Longitudinal Dynamic Stability

The tests for longitudinal short period dynamic stability are accomplished by a movement or pulse of the longitudinal control at a rate and degree to obtain a short period pitch response from the aircraft. Initial inputs should be small and conservatively slow until more is learned about the aircrafts response. Control inputs can then be made gradually large enough to properly evaluate the short period mode. The doublet input excites the short period oscillation while suppressing the phugoid. The doublet is performed, after trimming the aircraft at the required flight conditions, by applying a smooth, but fairly rapid, movement of the longitudinal control. Apply forward stick to decrease pitch attitude a few degrees then reverse the input to bring the pitch attitude back to trim. As pitch attitude reaches trim return the cockpit control to its original position and release it (controls free evaluation) or restrain it in the trim position (controls fixed). Both methods should be utilised. At the end of the doublet input, pitch attitude should be at the trim position (or oscillating about it) and airspeed should be approximately trim airspeed.


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The frequency of the doublet input depends on the response characteristics of the aircraft. The test pilot should adjust the frequency until the maximum response is generated. The short period mode should be investigated at selected points covering all flight conditions. AC 23-8B provides information on evaluating the aircrafts phugoid characteristics. 11. Explanation Lateral/Directional Dynamic Stability

Characteristic lateral/directional motions normally involve three modes:

a. A highly damped convergence, called the roll mode, through which the pilot controls roll rate and, hence, bank angle.

b. A slow acting mode, called the spiral, which may be stable but is often neutral or even mildly divergent in roll and yaw.

c. An oscillatory mode, called the Dutch roll, that involves combined rolling and yawing motions and may be excited by either rudder or aileron inputs or by gust encounters.

12. Procedures Lateral/Directional Dynamic Stability

Various techniques are available for assessing the lateral/directional dynamic stabilities. Some are discussed in AC 23-8B. Rudder pulsing, or doublets, can be used to excite the Dutch roll motion while suppressing the spiral mode.


SECTION 17 SPINNING

1. Requirements

Single engine aeroplanes not cleared for aerobatics must be able to recover from a one turn spin or a three second spin, whichever takes longer, in not more than one additional turn after initiation of the first control action for recovery. An alternative option is to demonstrate that the aeroplane is inherently resistant to spinning. In addition, an aeroplane that is to be cleared for aerobatics must also be able to recover from any point in a spin up to and including six turns in not more than one and a half additional turns after initiation of the first control action for recovery. 2. Explanation

A spin is a sustained autorotation at angles of attack above the stall. The rotary motions of the spin may have oscillations in pitch, roll and yaw superimposed upon them. The fully developed spin is attained when the trajectory has become vertical and the spin characteristics are approximately repeatable from turn to turn. Some aeroplanes can autorotate for several turns, repeating the body motions at some interval, and never stabilise. Most aeroplanes will not attain a fully developed spin in one turn. Some are reluctant to spin at all and may prefer to enter spiral dives. 3. Procedures

Conducting a thorough assessment of an aeroplanes spinning characteristics is a complex exercise. It should not be undertaken lightly and is only required for initial certification of the type design. Any advice from the aircraft designer that the aircraft is not cleared for intentional spinning should be heeded. FAA AC 23-8B provides detailed information on test techniques and procedures. In addition, FAA AC 23-15A provides an abbreviated spin test matrix that can be used to satisfy the requirements for light, simple aircraft.


SECTION 18 VIBRATION AND BUFFETING 1. Requirement

Each part of the aeroplane must be free from excessive vibration at all speeds up to VNE. In addition, in any normal flight condition, there must be no buffeting severe enough to interfere with the satisfactory control of the aeroplane, cause excessive fatigue in the crew, or result in structural damage. Stall warning buffeting within these limits is allowable. This requirement must be met with the engine running at all allowable power settings. 2. Explanation

The tests required under this section should not be confused with flutter tests. No attempt is made to excite flutter, but this does not guarantee against encountering it. Therefore the tests should be carefully planned and conducted. The indicated airspeeds for the tests should be determined from the airspeed calibration data. Careful study of the aeroplanes airspeed calibration is required with respect to the characteristics of the slope at the high speed end and how the airspeed calibration was conducted. This is necessary to determine the adequacy of the airspeed calibration for extrapolating to VNE. 3. Procedures

In the clean configuration at the gross weight, most critical CG (probably most aft) and the altitude selected for the start of the test, the aeroplane should be trimmed in level flight at maximum continuous power. Speed is gained in gradual increments in a dive until VNE is obtained. The aeroplane should be trimmed, if possible, throughout the manoeuvre. Remain at the maximum speed only long enough to determine the absence of excessive buffet, vibration or controllability problems. With flaps extended and the aeroplane trimmed for level flight at a speed below VFE, commence a shallow dive to stabilise at VFE and make the same determination as above.


SECTION 19 FLUTTER

1. Requirements

The aeroplane must be free from flutter, aerofoil divergence and control reversal in each configuration and for any condition of operation within the V-n diagram and at all speeds up to the applicable limiting speed. Flight flutter tests must be conducted. 2. Explanation

In addition to analytical methods compliance with the requirement for the airframe and control system to remain free from flutter can be shown by systematic flight tests designed to induce flutter at speeds up to VDF or the appropriate limiting speed. These tests must show that there is no rapid reduction of damping as the limiting speed is approached. Additionally,

a. control effectiveness around all three axes should not decrease in an unusually rapid manner, and

b. there should be no signs of approaching aerofoil divergence of wings, tailplane and fuselage indicated from the trend of the static stabilities and trim conditions.

3. Procedures

Conducting a thorough assessment of an aeroplanes flutter characteristics is a complex exercise and has the potential to be dangerous. Assistance from engineers experienced in flutter analysis is strongly recommended. FAA AC 23-629-1B Means of Compliance with Title 14 CFR, Part 23, 23.629 Flutter (Appendix 2 Flight Flutter Testing), provides detailed information on test techniques and procedures. FAA AC 23-15A and AC 90-89A provide additional and qualifying information applicable to light, simple aircraft.


SECTION 20 TAKE-OFF DISTANCE 1. Requirements

The take-off distance should be established from a smooth, dry, hard-surfaced runway and/or from a short dry grass surface, as applicable to the appropriate airworthiness standard. It is the distance required to reach a height of 50 feet from a standing start under the following conditions;

a. the engine(s) operating within maximum take-off power limitations,

b. the aircraft reaching a height of 50 feet at an airspeed not less than the take-off safety speed (VTOSS),

c. the landing gear extended throughout,

d. wing flaps in the take-off position,

e. maximum take-off weight and most forward CG, and

f. sea level ISA conditions.

A VTOSS should be established for each flap setting for which take-off distance information is to be provided. The VTOSS shall be an airspeed not less than 1.2 VS1 or VS1 plus 10 knots, whichever is the greater, at which adequate control is available in the event of sudden complete engine failure during the climb following take-off. Take-off charts, when included in the aircraft flight manual, shall schedule distances established in accordance with the above provisions, factored by 1.15. 2. Explanation

Take-off distance tests should be conducted in steady wind conditions, preferably nil wind. Gusty conditions will probably produce unacceptably inconsistent results. Suitable measuring techniques should be employed to measure the total take-off distance from a standing start to a height of 50 feet. The air and ground run segments of the total take-off distance must also be recorded for data reduction purposes. At least five take-offs should be flown with the final distance being the mean of the corrected (reduced) distances. The following quantities should be recorded for data reduction purposes:

a. Weight. Testing should commence with the aircraft loaded to the maximum take-off weight. Weights for subsequent test runs may be derived by recording the times and fuel burns at completion. The CG should be in its most forward position.

b. Wind. Wind velocity and direction should be measured adjacent to the runway during the time interval for each test run.


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c. Temperature and Pressure. Airfield temperature and pressure altitude should be recorded in conjunction with each test run.

d. Slope. Runway gradient can have a significant effect on the ground run distance for low thrust to weight ratio aircraft. The gradient of the take-off surface should be accounted for. Information on runway gradients should be available in the airfield survey documents.

3. Procedures

To reduce the results of take-off tests to ISA sea level surfaces corrections should be made for weight, wind, temperature, pressure altitude and runway slope. Take-off techniques should produce consistent results and not require undue skill or strength on the part of the pilot. FAA AC 23-8B provides detailed information on test techniques and data reduction procedures while FAA AC 23-15 provides additional information applicable to simple, light aircraft.


SECTION 21 CLIMB PERFORMANCE 1. Requirement Take-Off Climb

The gradient of climb after take-off should not be less than 8.3 percent5 under the following conditions:

a. sea level ISA,

b. maximum take-off weight and most forward CG,

c. airspeed not less than 1.2 VS1,

d. landing gear retracted

e. wing flaps in the take-off position,

f. engine operating at not more than maximum continuous power, and

g. in still air out of ground effect.

2. Procedures Take-Off Climb

With the altimeter adjusted to a setting of 1013.2 mb, a series of climbs, initiated at the lowest practical altitude, should be conducted. Stabilise airspeed and power prior to recording data. The time at the beginning of each run should be recorded for weight accounting purposes. Each stabilised climb should be continued for at least three minutes or 3000 feet while holding the airspeed essentially constant. Climbs should be conducted in sawtooth pairs on reciprocal headings to minimise the effects of windshear. Orientation of the climbs should be at 90 degrees to the prevailing wind direction, if known. Precise altimeter readings should be recorded at precise time intervals of not more than 30 seconds. Airspeed, ambient temperature and engine parameters should also be recorded although this can be done at longer intervals. A running plot of altitude versus time can be maintained to assess the acceptability of test data. It is essential to conduct climb tests in smooth air to obtain accurate results. The presence of an atmospheric temperature inversion will also produce unacceptable climb test results if the climbs are conducted through the inversion. The results of reciprocal heading climbs should be averaged before data reduction. Climb data is plotted using altitude and time as the horizontal and vertical axes. The slope of the tangent drawn against the curve at the mean test altitude produces the observed rate of climb. An acceptable method for correcting the test data for temperature, pressure altitude and weight should be used to reduce the data to standard sea level conditions. FAA AC 23-8B provides detailed information. 5 For FAR 23 aircraft. Check the appropriate airworthiness standard different requirements are applicable for different standards.


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3. Requirements Enroute Climb

The steady gradient and rate of climb should be determined at each weight, altitude and ambient temperature within the operational limits with:

a. not more than maximum continuous power,

b. the landing gear retracted,

c. wing flaps retracted, and

d. a climb speed not less than 1.3 VS1.

4. Procedures Enroute Climb

Procedures for establishing enroute climb performance are the same as those described above for the take-off climb. 5. Requirements Baulked Landing Climb

The steady gradient of climb should not be less than four percent at sea level ISA conditions with the aircraft configured as follows:

a. maximum landing weight,

b. climb speed not to exceed the approach speed,

c. engine(s) operating within take-off power limitations,

d. landing gear extended, and

e. wing flaps in the landing position, except that, if safe and rapid retraction to at least the take-off position is possible without causing excessive change in angle of attack or loss of altitude and without requiring exceptional piloting skill, the flaps may be retracted.

If compliance with the requirement is conditional on wing flaps being retracted a statement should be included in the aircraft flight manual detailing the procedure that must be followed in the event of a baulked landing manoeuvre. It must be possible to make a safe transition to the enroute climb configuration and speed from the recommended approach speed. 6. Procedures Baulked Landing Climb

Procedures for establishing baulked landing climb performance are the same as those described above for the take-off climb. The additional requirement is for the test pilot to make a qualitative judgement as to whether the transition from approach to climb configurations can be achieved without requiring exceptional piloting effort or skill.


SECTION 22 GLIDE PERFORMANCE 1. Requirements

For single engine aeroplanes the maximum horizontal distance travelled in still air, in nautical miles per 1000 feet of altitude lost in a glide, and the speed necessary to achieve this, should be determined with:

a. the engine inoperative,

b. its propeller in the minimum drag position, and

c. landing gear and flaps in the most favourable available positions.

2. Explanation

The primary purpose of this requirement is to provide the pilot with information about the aircrafts gliding performance. Such data can be used as a guide to the gliding range that can be achieved. Some reasonable approximation in its derivation is acceptable. 3. Procedures

A method of determ

testing generic

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