management software (FMS) and autopilot capability should consider a comprehensive F&R testing program regardless of weight.)
b. Procedures.
(1) Test Plan. Efforts should begin early in the certification program to provide assistance to the applicant to develop a certification plan and subsequently a test plan to ensure coverage of all certification requirements. The applicant should develop a test plan that includes the required instrumentation.
(2) Instrument Calibration. Test instrumentation (trans ducers, indicators, and other installed instrumentation) should be calibrated (removed from the airplane and bench checked by an approved method in an approved facility) within six months of the tests. When electronic recording devices are used, such as oscillographs, data loggers, and other electronic data acquisition devices, preflight and post-flight parameter recalibrations should be run for each test flight to ensure that none of the parameters have shifted from their initial zero settings. Critical transducers and indicators for critical tests (for example, airspeed indicators and pressure transducers for flight tests to VD) should be calibrated within 60 days of the test in addition to the other requirements mentioned above. The instrument hysteresis should be known; therefore, readings at suitable increments should be taken in both increasing and decreasing directions.
Calibration records, like the one shown in table 3, should be signed by the agent of the repair or overhaul facility doing the work and be available to the test pilot prior to beginning test flying. It should be emphasized that these calibrations must be accomplished at an approved facility. For example, performing a pitot-static/air data system leak check to "calibrate" an airspeed indicator, whether in or out of the airplane, is not acceptable.
Table 3 – Sample Portion of Airspeed Indicator Calibration XYZ INSTRUMENT SERVICE, INC.
ABC CITY AIRPORT APPROVED REPAIR STATION - NO. 1234 8/12/80 P/N 1701DX8-04 S/N AF55-17044 A/S Indicator KNOTS Master Test Ascent Indicator Reads Descent Indicator Reads 40 38.0 39.0 50 49.0 50.5 60 59.5 61.0 70 70.0 71.0 80 80.0 81.0
(3) Use of Ballast.
(i) Loading. Ballast loading of the airplane can be accomplished several ways to achieve a specific weight and c.g. location as long as the loading remains within the physical confines of the airplane. In flight test work, loading problems will occasionally be encountered, making it difficult to obtain the desired c.g. location. Those cases may require loading in engine compartments or other places not designed for load carrying. When this condition is encountered, care should be taken to ensure that local structural stresses are not exceeded or that airplane flight characteristics are not changed due to changes in moments of inertia caused by adding a very long arm (tail post, and so forth).
(ii) Solid and Liquid Ballast. There are basically two t ypes of ballast that may be used in airplane loading: solid or liquid. The solids are usually high-density materials such as lead or sandbags, while the liquid is usually water. In critical tests, the ballast should be loaded in a manner so that disposal in flight, if considered, can be accomplished and be located at a point that will produce a significant c.g. shift to a less critical c.g. location. Applicants should perform a simple structural analysis for complex load installations that have the potential to interfere with safe flight and landing if there was a failure of the supporting structure. In any case, the load should be securely attached in its loaded position. In airplanes with multiple fuel 5
tank arrangements, the fuel load and distribution should be considered for weight and c.g.
control.
(4) Function and Reliability (F&R) Tests for Airplanes over 6,000 pounds Maximum Certificated Weight. While § 21.35(b)(2) does not require F&R testing for airplanes weighing less than 6000 lbs., the Small Airplane Directorate’s policy is to require (via special conditions) a comprehensive Function and Reliability testing program regardless of weight for complex airplanes that include features like FADEC controlled engines, pressurization, and avionics integrated with FMS and autopilot capability.
(i) A comprehensive and systematic check of all aircraft components must be made to assure that they perform their intended function and are reliable.
(ii) F&R certification testing must be accomplished on an aircraft that is in conformity with the approved type design. Furthermore, the number of airplanes used should be limited to meet the intent of F&R testing. F&R certification testing should follow the type certification testing. This is to assure that si gnificant changes resulting from type certification tests can be incorporated on the aircraft and successfully retested prior to the F&R tests.
(a) A certain portion of the F&R test program should emphasize systems, operational conditions, or environments found particularly marginal during type certification tests. Applicants should put the aircraft in as many different environments as possible (for example, cold, hot, rain, etc.) as well as different types of airports to exercise the systems as they will be in the field.
Note: It is important to perform F& R testing with the configuration the manufacturer is planning to deliver to the customer (this may be different than what is configured at the end of initial certification). This is especially important in the case of software and post-TC changes.
(iii) All components of the aircraft should be periodically operated in sequences and combinations likely to occur in service. Ground inspection should be made at appropriate intervals to identify potential failure conditions; however, no special maintenance beyond that described in the Aircraft Maintenance Manual should be allowed.
(iv) A complete record of defects and failures should be maintained along with required servicing of aircraft fluid levels. Results of this record should be consistent with inspection and servicing information provided in the Aircraft Maintenance Manual.
2. § 23.23 Load Distribution Limits.
a. Explanation.
(1) C.G. Envelope. The test tolerance of ±7 percen t of the total c.g. range (given in § 23.21) is intended to allow some practical relief for in-flight c.g. movement. This relief is only acceptable when the test data general scatter is on either side of the limiting c.g. or when c.g.
correction from test c.g. to limit c.g. is acceptable. Sufficient points inside the desired weight and balance envelope should be explored to ensure that the operational pilot will not be placed in an unsafe condition. Should unsatis factory flight characteristics be present, the limits of the envelope should be reduced to ensure safe margins. Where variation in the c.g. position may have a significant effect on the result of a test (for example, spins and VMC), the result should be corrected to the most critical c.g. position within the operating limits to be approved. If such a correction is impractical or may be unreliable, the actual test should ensure that the measured characteristics represent the critical value.
(2) Narrow Utility C.G. Envelope. Some utility category airplanes, for which spin approval is sought, may have a very narrow c.g. range. If a limited fuel load is required to achieve the narrow c.g. envelope, the test pilot should ensure that loading instructions or aids (such as fuel tank tabs) will enable the operational pilot to stay in the approved c.g. envelope.
(3) Gross Weight Effects. The test pilot is expected to determine the effect that gross weight, including low-fuel state, may have on the airplane's flight characteristics. If it is found the flight characteristics would be adversely affected, tests should be performed for trim, stability, and controllability including VMC, stalls, and spins under the most adverse weight condition. Separate loading restrictions may apply to certain flight operations, such as spins.
(4) Lateral Loads. If possible loading conditions can result in a significant variation of the lateral c.g., this lateral range of c.g. must be established by:
(i) The limits selected by the applicant;
(ii) The limits for which the structure has been proven; or
(iii) The limits for which compliance with all the applicable flight requirements has been demonstrated. The demonstrated we ight and c.g. combinations should consider asymmetric loadings. When investigating the effects of asymmetric lateral loads, the following sections in this flight test guide (FTG) represent applicable flights requirements:
(a) § 23.143 Controllability and Maneuverability, General.
(b) § 23.147 Directional and lateral control.
(c) § 23.149 Minimum control speed.
(d) § 23.151 Acrobatic maneuvers.
(e) § 23.157 Rate of roll.
(f) § 23.161 Trim.
(g) § 23.201 Wings level stall.
(h) § 23.203 Turning flight and acc elerated turning stalls.
(i) § 23.221 Spinning.
(j) § 23.233 Directional stab ility and control.
(k) § 23.701 Flap interconnection.
b. Procedures. None.
3. § 23.25 Weight Limits.
a. Explanation.
(1) Maximum Weight Limits. The maximum weight may be limited in three ways: at the election of the applicant, by structural design requirements, or by flight requirements.
(2) Maximum Weight Exceptions. The regulations concerning design maximum weight allow an exception that some of the structural requirements may be met at a lesser weight known as a design landing weight, which is defined in § 23.473. (Refer to Advisory Circular (AC) 23-7 if the airplane is being modified for an increase in maximum weight.) Due to changes in the operational requirements of an owner/operator, in many cases the need arises to modify and substantiate the structure for an increase in maximum weight and maximum landing weight.
Any one of these increases affects the airplane's basic loads and structural integrity, and they could affect the limitations and performance.
(i) If an airplane was certificated with maximum landing weight equal to maximum weight, some applicants take advantage of the five percent difference between design landing and design maximum weight permitted by § 23.473(b). These applicants use the STC process. In these cases, re-substantiation of th e landing gear for landing loads is not required when increasing the maximum weight by as much as five percent. For those programs involving more than a five percent increase in maximum weight, some re-substantiation of the landing gear should be accomplished.
(ii) Other applicants are replacing piston engines with turbopropeller engines, thus requiring that gasoline be replaced with jet fuel, which weighs as much as 17 percent more.
In some cases, the quantity of fuel is being increased at the same time as engine replacement, but the maximum zero fuel weight remains the same.
(iii) All the above types of modifications should be investigated to verify that critical loads have not increased or that those loads that have increased are capable of being carried by the existing or modified structure.
(3) Weight, Altitude, Temperature (WAT). For all airplanes with a maximum takeoff weight exceeding 6,000 pounds and turbine engine airplanes, a WAT chart may be used as a maximum weight limitation.
(4) Ramp Weight. The applicant may elect to use a "ramp weight" provided compliance is shown with each applicable section of part 23. Ramp weight is the takeoff weight at brake release plus an increment of fuel weight consumed during engine start, taxiing, and runup. Generally, this increment of fuel should not exceed one percent of the maximum permissible flight weight up to 125 pounds. The pilot should be provided a means to reasonably determine the airplane gross weight at brake release for takeoff. A fuel totalizer is one way of providing the pilot with fuel on board. Alternately, a mental calculation by the pilo t may be used, if the pilot is provided the information to make the calculation and the calculation is not too complex.
Normally, fuel for engine start and run-up will be sufficiently close to an amount fixed such that taxi can be considered as the only variable. If the pilot is provided with taxi fuel burn rate in 8
pounds per minute, then the resulting mental calculation is acceptable. The pilot will be responsible to ensure that the takeoff gross weight limitation is complied with for each takeoff, whether it is limited by altitude, temperature, or other criteria. The maximum ramp weight should be shown as a limitation on the TC Data Sheet and in the AFM.
(5) Lowest Maximum Weight. Sections 23.25(a)(2)(i) and 23.25(a)(2)(ii) require that each of the two conditions, (i) and (ii), must be considered and that the maximum weight, as established, not be less than the weight under either condition. This has to be shown with the most critical combinations of required equipment for the type of operation for which certification is requested.
(6) Placarding of Seats. When establishing a maximum weight in accordance with § 23.25(a)(2)(i), one or more seats may be placarded to a weight of less than 170 pounds (or less than 190 pounds for utility and acrobatic category airplanes). An associated requirement is § 23.1557(b). The AFM loading instructions, requ ired by § 23.1589(b), should be specific in addressing the use of the placarded seats.
b. Procedures. None.
4. § 23.29 Empty Weight and Co rresponding Center of Gravity.
a. Explanation.
(1) Fixed Ballast. Fixed ballast refers to ballast that is made a permanent part of the airplane as a means of controlling the c.g.
(2) Equipment List. Compliance with § 23.29(b) may be accomplished with an equipment list that defines the installed equipment at the time of weighing and the weight, arm, and moment of the equipment.
b. Procedures. For prototype and modified test airplanes, it is necessary to establish a known basic weight and c.g. position (by weighing) from which the extremes of weight and c.g.
travel required by the test program may be calculated. Normally, the test crew will verify the calculations.
5. § 23.31 Removable Ballast.
a. Explanation. This regulation is associated only with ballast that is installed in certificated airplanes under specified conditions. The ballasting of prototype airplanes so that flight tests can be conducted at certain weight and c.g. conditions is covered under § 23.21, paragraph 3, of this AC.
b. Fluid Cargo. For those airplanes configured to carry fluid cargo (such as agricultural chemical tanks, minnow tanks, slurry tanks, and so forth), airplane handling qualities should be evaluated for controllability and non-exceedance of the limitations at the full and the most critical partial fluid loads. When so equipped, the effects of in-flight jettison or dumping of the 9
fluid load should be evaluated to establish that the pilot is able to exercise sufficient control to prevent unacceptably large flight path excursions or exceedance of operational/structural limits.
6. § 23.33 Propeller Speed and Pitch Limits.
a. Explanation. Section 23.33(a) requires that propeller speed and pitch be limited to values that will ensure safe operation under normal operating conditions.
b. Procedures.
(1) Fixed Pitch Propellers.
(i) Maximum Revolutions per Minute (r.p.m.). The regulation is self-explanatory.
(ii) Static r.p.m. Determine the average static r.p.m. with the airplane stationary and the engine operating at full throttle under a no-wind condition. The mixture setting should be the same as used for maximum r.p.m. determination. If the wind is light (5 knots or less), this static r.p.m. can be the average obtained with a direct crosswind from the left and a direct crosswind from the right.
(iii) Data Sheet r.p.m. Determination. For fixed pitch prope llers, the static r.p.m.
range is listed in the TC Data Sheet: for example, not more than 2200 r.p.m. and not less than 2100 r.p.m. The allowable static r.p.m. range is normally established by adding and subtracting 50 r.p.m. to an average no-wind static r.p.m. Applicants should account for altitude and temperature effects in the TCDS. An applicant may desire to obtain approval for one or more additional propellers and retain only one r.p.m. range statement. An applicant may also choose to extend the propeller's static r.p.m. range.
(a) Lower r.p.m.. The static r.p.m. range may be extended on the low side by obtaining approval for a propeller with a lower static r.p.m. In this case, the approval must be accomplished with due consideration of performance requirements. The airplane with the new propeller installed must be able to meet the minimum climb performance requirements.
(b) Higher r.p.m.. If the static r.p.m. range is to be extended upward, the new propeller would have to be tested to ensure that it did not cause an engine speed above 110 percent of maximum continuous speed in a closed throttle dive at the never-exceed speed. It must not exceed the rated takeoff r.p.m. of the engine up to and including the best rate of climb speed of the airplane. An engine cooling climb test may also be required due to the additional power produced by the faster turning propeller.
(2) Controllable Pitch Propellers Without Constant Speed Controls.
(i) Climb r.p.m. With the propeller in full low pitch, determine that the maximum r.p.m. during a climb using maximum power at the all-engine(s)-operating climb speed does not exceed the rated takeoff r.p.m. of the engine.
(ii) Dive r.p.m. With the propeller in full high pitch, determine that the closed throttle r.p.m. in a dive at the never-exceed speed is not greater than 110 percent of the rated maximum continuous r.p.m. of the engine.
(3) Controllable Pitch Propellers With Constant Speed Controls.
(i) Climb r.p.m. With the propeller governor operative and propeller control in full high r.p.m. position, determine that the maximum power r.p.m. does not exceed the rated takeoff r.p.m. of the engine during takeoff and climb at the all-engine(s)-operating climb speed.
(ii) Static r.p.m. With the propeller governor made inoperative by mechanical means, obtain a no-wind static r.p.m.
(a) Reciprocating Engines. Determine that the maximum power static r.p.m., with the propeller blade operating against the low pitch stop, does not exceed 103 percent of the rated takeoff r.p.m. of the engine.
(b) Turbopropeller Engines. Although this rule references manifold pressure, it has been considered to be applicable to turbopropeller installations. With the governor inoperative, the propeller blades at the lowest possible pitch, with takeoff power, the airplane stationary, and no wind, ensure that the propeller speed does not exceed the maximum approved engine and propeller r.p.m. limits. Propellers that go to feather when the governor is made inoperative need not be tested.
(iii) Safe Operation Under Normal Operating Conditions.
(a) Reciprocating Engines. Descent at VNE with full power, although within the normal operating range, is not a normal operating procedure. Engine r.p.m., with the propeller on the high pitch blade stops, that can be controlled by retarding the throttle may be considered as acceptable in showing compliance with § 23.33(a).
(b) Turbopropeller Engines. Perform a maximum r.p. m. at maximum torque (or power) descent at VMO to ensure that normal operating limits for the propeller are not exceeded.
(4) Data Acquisition and Reduction. Outside air temperature and altitude needs to be considered for the ground tests. The observed r.p.m. data in each case must be corrected for tachometer error. The airspeed system error must also be taken into consideration to determine the proper calibrated airspeed. True airspeed may also need to be considered because propeller angle of attack is a function of true airspeed.
Section 2. Performance
1. § 23.45 General.
a. Explanation.
(1) Atmospheric Standards. The purpose of § 23.45(a) is to set the atmospheric standards in which the performance requirements should be met. The air should be smooth with no temperature inversions, mountain waves, and so forth. This is essential to obtaining good data and repeatable results. N on-standard conditions of temperature, pressure, and so forth, can be corrected to standard, but there are no corrections to compensate for poor quality data due to turbulence or poor pilot technique. A thorough knowledge of the limitations of the testing procedures and data reduction methods is essential so that good engineering judgment may be used to determine the acceptability of any tests.
(i) Reciprocating Engine-Powered Airplanes Below 6,000 pounds (2730 kg) Maximum Weight. Performance tests will normally be conducted in non-standard atmospheric conditions, but ideally for accuracy in data reduction and expansion, tests should be conducted in atmospheric conditions as near those of a standard atmosphere as possible. Accounting for winds and non-standard conditions requires testing procedures and data reduction methods that reduce the data to standard atmospheric conditions.
(ii) Reciprocating Engine-Powered Airplanes of More Than 6,000 pounds (2730 kg) Maximum Weight and Turbine-engine Powered Airplanes. Performance tests should be conducted in the range of atmospheric conditions that will show compliance with the selected weight, altitude, and temperature limits. Refer to § 23.53 of this AC for guidance on extrapolation of takeoff data and § 23.75 for extrapolation of landing data.
(2) Standard Atmosphere. The Standard Atmosphere is identical to the International Civil Aviation Organization (ICAO) Standard Atmosphere for altitudes below 65,000 feet.
Appendix 7, figure 1, gives properties of the Standard Atmosphere in an abbreviated format.
(3) Installed Power. The installed propulsive horsepow er/thrust of the test engine(s) may be determined using the applicable method described in appendix 1, based on the power approved during airplane certification. The meth ods in appendix 1 account for installation losses and the power absorbed by accessories and services. Consideration should also be given to the accuracy of the power setting instruments/systems, and the pilot's ability to accurately set the power/thrust.
(4) Propeller Cutoff. In general, if the airplane will be approved with an allowable cutoff for the propeller, then the performance flight testing should be done using the most critical propeller diameter. In most cases, this is expected to be the minimum diameter propeller allowed.
(i) For normal, utility, and acrobatic category airplanes only, a two percent margin was the allowed reduction in propeller diameter. Historically, the two percent margin was selected as being the maximum permissible reduction in diameter of a given propeller that will not noticeably reduce performance. Service history has shown this to be an acceptable margin without additional flight testing.
(5) Flight Procedures. The flight procedures must not be unduly sensitive to less than ideal atmospheric conditions. The atmosphe ric conditions "reasonably expected to be encountered in service" may be different depending on the class of aircraft, but should cover at least the maximum demonstrated crosswind component established in compliance with § 23.233(a).
(6) Flight Test Data. For calibrated engines, test day power is the calibrated test day power. For uncalibrated engines, an acceptable met hod is to assume that the test day power is the upper tolerance chart brake horsepower; however, the upper tolerance chart brake horsepower for reciprocating engines using 14 CFR part 33 tolerances of +5 percent, -0 percent may lead to unrealistic performance. Historically, the two major reciprocating engine manufacturer’s engines have tolerances closer to +2 percent, -0 percent. Refer to appendix 1 for further information. The performance data required by § 23.1587 is dependent on the horsepower assumed for the various temperature and altitude conditions. Refer to appendix 1, which deals both with test data reduction and expansion.
(7) Humidity Correction. Refer to appendix 1.
b. Procedures. Refer to appendix 1.
c. Operation on Unpaved Runways.
(1) Small airplane operations from grass runways. For airplanes less than 6,000 pounds (2730 kg) maximum weight, the factors given below may be quoted in the flight manual, as an alternative to the scheduling of data derived from testing or calculation. It should be noted that these factors are intended to cover the range of airplane types in this category and are necessarily conservative. Manuf acturers are, therefore, encouraged to produce and schedule their own data in accordance with (1) and (2) above to obtain optimized performance for their airplane.
Table 4 - Smooth, Flat Runway Takeoff Landing Dry Grass 1.2 1.2 Wet Grass 1.3 1.6 Note: If the runway is not smooth, the grass is very long or very short, higher factors may be warranted.
Very short grass = golf course length Very long grass = unmowed field or high interstate highway median grass length
(2) Airplanes with 6,000 pounds (2730 kg) maximum weight or more. The following guidance was developed to include high performance multiengine airplanes. The level of detail should be scaled back to that appropriate for the airplane. Simple, single-engine turbine bush airplanes, for example, do not need the level of detail discussed here.
Operations on other than smooth, dry, hard runway surfaces may require specific approval and the scheduling of information on the effect of those surfaces on takeoff and landing distances in the flight manual. Applicants may use the factors for grass runways in paragraph
(1). These factors would, in many cases, result in such conservative distances that operational utility may be compromised. Applicants may propose their own factors if accompanied with supporting data. To obtain approval for take off and landing operations on unpaved runway surfaces, compliance with the following should be shown:
(i) Each type of surface must be defined so that it can be recognized in operations in service including conditions such as day or wet.
(ii) It must be determined that the airplane can be operated on each defined surface without hazard from likely impingement or engine ingestion of any foreign objects that constitute parts of the surface.
(iii) If any special procedures or techniques are found to be necessary, these must also be determined and scheduled.
(iv) The takeoff and landing performance on each defined surface must be determined in accordance with §§ 23.53 and 23.75, as modified in (3) below.
(3) Takeoff and Landing Data. Takeoff and landing data must be determined for each type of unpaved surface for which approval is requested.
(i) The test runways on which the takeoff and landing distance measurements are conducted should be chosen to be representative of the worst characteristics (that is, high rolling friction, low braking friction) of each of the types of runway under consideration.
(ii) In establishing the operation limitations for a particular type of unpaved runway, the runway’s load bearing characteristics, rolling and braking friction, and impingement and ingestion characteristics, should be considered.
e. Wet and Contaminated Runways. Reserved.
2. § 23.49 Stalling Speed.
a. Explanation.
(1) 61 Knot Stall Speed. The 61 knot stalling speed applies to the maximum takeoff weight for which the airplane is to be certificated.
(2) Background. Since many of the regulations pertain to performance, handling qualities, airspeed indicator markings, and other variables that are functions of stall speeds, it is desirable to accomplish the stall speed testing early in the program so the data are available for subsequent testing. Because of this interrelationship between the stall speeds and other critical performance parameters, it is essential that accurate measurement methods and careful piloting techniques be used. Most standard airplane pitot-static systems have not been found to be acceptable for stall speed determination. These tests require the use of properly calibrated instruments and usually require a separate test airspeed system, such as a trailing bomb, a trailing cone, or an acceptable nose or wing boom. The stall speed determinations necessary for marking the airspeed indicator are in terms of indicated airspeed (IAS) corrected for instrument error.
The other stall speeds are in terms of calibrated airspeed (CAS). Thus, a production airspeed system should be available during stall speed measurements to determine stall speeds in terms of IAS.
(3) Stall Definition. Section 23.49(d) requires the V S0 and VS1 speeds to be determined using the procedures specified in § 23.201. Refer to 14 CFR part 1 and § 23.49 for definitions of VS0 and VS1. “Landing configuration” in part 23 is gear down and full flaps.
Section 23.201(b) defines when the airplane can be considered stalled for airplane certification purposes when one of three conditions occurs, whichever occurs first. The conditions are:
(i) Uncontrollable downward pitching motion;
(ii) Downward pitching motion resulting from the activation of a device (for example, stick pusher); or,
(iii) The control reaches the stop.
(a) For those airplanes where the control reaches the stop, VS is considered to be the minimum speed obtained while the control is held against the stop. Elevator limited airplanes may or may not develop a minimum steady flight speed. Refer to figure 1 for a graphic representation of stall speed time histories for various configurations. The time the control is held against the stop for stall speed determination should be a minimum of two seconds and consistent with the time against the stop for stall characteristics testing in § 23.201.