TITLE PAGE PREPARED BY DATE REPORT NO. _
CHECKED BY DATE MODEL
CORRECTION OF LANDING DATA TO STANDARD CONDITIONS PÕ LANDING NUMBER 1 2 3 P23 (11) (22) + (21) = AIR RUN CORRECTION FOR WIND (FT.)
24 VI = 13 VSTALL. @ HD = 13 x 1.467 VSO v1/(3) x)Po/7 (FT./SEC.)
25 (24) 2- (16)2+3220 26 (14) 3 - (16) + 3220 27 (25) / (26) = AIR RUN CORRECTION FOR APPROACH SPEED 28 (14) 2 + 3220 29 28 - = WEIGHT CORRECTION FOR TOTAL RUN
(3) (14) = + 3220 30 28 = DENSITY CORRECTION FOR TOTAL RUN
(14) / (9) + 3220 31 S 40 = MEASURED AIR DISTANCE (CAMERA DATA) (FT.)
32 Sco = MEASURED GROUND DISTANCE (CAMERA DATA) (FT.)
33 (15) / 1(15) ½ = (32) 2gSIN] = RUNWAY SLOPE CORR.
34 SAI = |(31) + (23)| (29) (30) = CORR. AIR DIST. @ TEST APPROACH SPEED (FT.)
35 SA2 = (34) (27) = CORR. AIR DISTANCE @ 1.3 VSo APPROACH SPEED (FT.)
36 G= (32) (18 (29) (30) (33) = CORRECTED GROUND DISTANCE FT.
37 STI (34) + (36) = CORRECTED TOTAL DISTANCE @ TEST APPROACH SPEED (FT) 38 ST2 (35) + (36) = CORRECTED TOTAL DISTANCE @ 1.3 Vso APPROACH SPEED (FT) Figure 11 - DATA REDUCTION FORMCORRECTION OF LANDING DATA TO STANDARD CONDITIONS (SHEET 2 OF 2)
e. Balked Landing.
(1) Regulation Reference. Section 23.77.
(2) Discussion.
(a) The equivalent altitude method of climb performance data reduction may be used for airplanes with fixed pitch propellers. The discussion under § 23.65 applies.
(b) For reduced testing requirements, the following process may be employed:
21 1 . At a convenient altitude, in smooth air, conduct climbs through the same altitude band in opposite directions at the same speed. These climb segments should be of approximately three minutes duration (subject to engine cooling restraints). Climbs should be performed perpendicular to reported winds at test altitude. Altitude should be recorded at 30-second intervals.
2 . Reduce these climb data by the equiva lent altitude method and average the two rates of climb.
(c) Ensure the climb gradient equals or exceeds 1:30, as required in § 23.77.
f. Longitudinal Control .
(1) Regulations Reference . Sections 23.145, 23.161, and 23.677(b).
(2) Discussion . Sections 23.145(e)(1), 23.161, and 23.677(b) together have been interpreted to require an elevator trim tab system or an elevator trim system totally independent from the elevator control system. To ensure that a tab system is flutter free with the disconnection of a control element in the tab system (as required by § 23.629(f)), a fully redundant tab control system back to the trim control is effectively required. For small airplanes, the following simplified methods are permitted:
(a) A longitudinal control system that pr ovides trim control through a bungee system connected to the elevator horns complies with this rule.
(b) A bungee trim system attached not to the elevator horn, but to a pushrod, or pushrods, co nnected to the elevator also comply with the rule.
g. Static Longitudinal Stability .
(1) Regulation Reference . Section 23.175.
(2) Discussion :
(a) Stick force curve requirements are specifi ed by § 23.175, but need not be m easured with instruments if changes in speed are clearly reflected by changes in stick forces.
(b) Because of the relatively small speed and center of gravity (CG) envelopes, a sim plified matrix of the configurations to be tested is an acceptable method of compliance.
Additional guidance is available in AC 23-8B.
1. Compliance with the rule may be met by a qualitative assessment of stick force gradients by the test pilot for the following matrix of flight conditions. Tests should be conducted at the most critical weight and CG with gear extended. For light aircraft, the most critical condition is usually lightweight and aft CG.
(aa) Section 23.175(a) Climb - flaps up , gear up, 75 percent power, trimm ed to Vy;
22 (bb) Section 23.175(b) Cruise - flaps up, gear up 75 percent powe r, trimmed for level flight; and
(cc) Section 23.175(c) Appr oach/Landing - landing flaps, power idle, trimm ed to the recommended landing approach speed. Repeat with power for a three-degree descent.
2. For any of these four conditions where stability is marginal, as determined by the test pilot, the stick force gradient must be measured.
3. For example, if stability in climb at aft CG were the only marginal stability flight condition, a measurement of stick force gradient with pertinent comments in this condition would comply with the rule.
h. Stall .
(1) Regulations Reference . Sections 23.201 and 23.203.
(2) Discussion . Compliance may be shown by qualitative test pilot assessments. On airplanes without gyros, a flat Plexiglas plate over the glare shield on which the test pilot can mark bank angles with a grease pencil may be used to substantiate bank angles.
(a) Test pilot notation of altitude loss may be from observations of a standard altimeter.
(b) A Plexiglas plate with incremental ba nk angles p laced over the glare shield and a video camera may be used to record bank angles.
(c) Compliance may be shown by rational analysis with testing at forward CG for stall speed determ ination and aft CG for stall characteristics determination.
i. Stall Warning .
(1) Regulation Reference . Section 23.207.
(2) Discussion . Compliance with this rule for small airplanes is found where the stall warning begins at not less than five knots above stall speed. The upper limit for stall warning must not produce a nuisance.
j. Spin .
(1) Regulation Reference . Section 23.221.
(2) Discussion . AC 23-8B, with or without the following modification, is considered an acceptable method of compliance for small airplane spin testing.
(a) Modifications of AC 23-8B, Provisions . The following modifications of AC 23-8B, paragraphs are accepted for application to normal category spin substantiation:
1. b(1) Weight and CG Envelope . For normal category spins, the corners of the approved loading envelope should be tested.
2. b(3) Control Deflections . For airplanes with rigging tolerances specified as +/-1 degree or less, control surface rigging may be set to nominal specifications. For airplanes with rigging tolerances specified as greater than +/-1 degree, tests should be conducted at nominal specifications; and if any spin configuration is found to be critical, then the effects of rigging should be investigated for that condition.
3. b(9) Trimmable Stabilizer . The effect of the trimmable stabilizer on spins should be determined for the critical conditions with the trim set for 1.5 Vs1 in the cruise configuration, 1.2 Vs1 in the takeoff configuration, and 1.3 Vs1 in the landing configuration.
4. c(3) Recovery From Spins Following Abnormal Control Usage During Entry and Recovery. The provisions of this paragraph should be applied as necessary based on the outcome of the testing conducted in paragraph 5c(4).
sentence:
5. c(4) Spin Matrix. Add the following to this paragraph after the present first
• The airplane configurations to be evaluated should correspond to the configurations expected to be used for takeoff, cruise, and landing. For a normal category airplane, the spin test matrix should cover those abnormal control inputs reasonably expected from a pilot experiencing an inadvertent spin entry such as forgetting to reduce power, reflexively applying anti-spin aileron (against the turn) in response to roll-off, or inadvertently reversing recovery rudder and elevator sequencing. The “Normal Category Spin Test Matrix” is an acceptable method of compliance, unless unusual response requires additional exploration (see figure 12).
NORMAL CATEGORY SPIN TEST MATRIX Configuration* Normal Spins Level Entry Left (Lt) Turn Right (Rt) Turn Clean, Power Off 1 Lt, 1 Rt 1 Lt, 1 Rt 1 Lt, 1 Rt Takeoff, Power On 1 Lt, 1 Rt 1 Lt, 1 Rt 1 Lt, 1 Rt Landing, Power Off 1 Lt, 1 Rt 1 Lt, 1 Rt 1 Lt, 1 Rt Abnormal Spins Power On Power Off Power Off Ailerons Ailerons Elevator 1st Against Against Recovery Clean 1 Lt, 1 Rt 1 Lt, 1 Rt 1 Lt, 1 Rt Takeoff 1 Lt, 1 Rt 1 Lt, 1 Rt 1 Lt, 1 Rt Landing 1 Lt, 1 Rt 1 Lt, 1 Rt 1 Lt, 1 Rt *For normal category airplanes, the configurations to be examined are defined as follows:
Clean Flaps up Gear up Cowl flaps closed Takeoff Flaps at maximum approved to takeoff setting Gear down Cowl flaps open Landing Flaps full down Gear down Cowl flaps closed Each of these configurations should include testing at the combination of weight and CG determined to be most critical for the proposed loading envelope.
Figure 12 - NORMAL CATEGORY SPIN TEST MATRIX
6. c(5) Unrecoverable Spin. Any spin that cannot be recovered within the required spin criteria.
(b) Additional Guidance: Spin Resistant Airplanes. The following guidance on the application of the new § 23.221(a) (2) spin resistant provision is provided:
25 1 . In § 23.221(a)(2)(ii), "with the ailerons deflected opposite the direction of turn in the most adverse manner" means anti-spin or "reflexive" ailerons.
2 . In § 23.221(a)(2)(ii), "respond immediat ely" means recovery in approximately one-quarter turn or that amount that the test pilot determines is immediate and repeatable.
3 . The preceeding spin test guidance does not reduce the regulatory burden to show that it is “impossible to obtain” an unrecoverable spin for a given airplane design. The manufacturer is responsible for compliance with the regulations and should perform a thorough spin test evaluation. This AC simply describes a methodology that the FAA has found is adequate for a showing of compliance to the spin requirements for small, simple, low performance airplanes.
Finding literal compliance to the regulation that states “… it must be impossible to obtain unrecoverable spins…” is not practical since this would necessitate an infinite number of test points.
The methods described have been found to be an adequate number of test points to find compliance for these airplanes. This is derived from experience, technical data, and judgment, through many years using AC 23-8A for showing compliance on more complex airplanes. The methodology in this AC is a refinement of the method in AC 23-8A for showing compliance to the spin requirement of 23.221(a)(1)(iii) for these simple airplanes.
k. Emergency Landing Dynamic Conditions .
(1) Regulations Reference . Section 23.562.
(2) Discussion .
(a) See AC 23.562-1 “Dynamic Testing of Pa rt 23 Airplane Seat Restraint Sy stems and Occupant Protection” for information. AC 23.562-1 addresses all part 23 airplanes, and has no specific guidance for “small, simple, low performance airplanes”. However, all of the guidance material may be applied to “small, simple, low performance airplanes”.
(b) See AC 23-11A “14 CFR Part 23 Type Certif ication of an Airplane Originally Certificated to Joint Aviation Regulations - Very Light Airplane (JAR-VLA) Standards”. Some small, simple, low performance airplanes, are similar to the JAR-VLA discussed in AC 23-11A, and therefore, that AC's guidance material may be of interest to the reader.
l. Fabrication Methods .
(1) Regulation Reference . Section 23.605.
(2) Process Specifications Format . The following sample process specification (for welding) provides an acceptable format. An applicant may develop specifications in this format.
(a) 1.0 PURPOSE
This section should cover what the sp ecification will control. Such as,
This specif ication establishes the engineering requirements for nam e process to be used by company name .
Personnel and machines used shall be able to meet or exceed this s pecification.
(b) 2.0 REFERENCES
The fo llowing publications are to be used to clarify this specification or be the ba sis for testing or equipment/personnel certification:
List other co mpany specifications, military specifications (MIL SPEC) or standards."
(c) 3.0 STANDARDS
3.1 Weld word vocabulary reference AWS A3.0-8 for definitions used in weld rela ted work.
3.2 Examples of acceptable and unacceptable weld should be shown in this section.
3.3 Method of certification and recertification of the welder. Weld specime ns and qualification and requalification test conditions.
3.4 Visual and radiographic inspection, metallographic examination and standards should be included in this section as applicable.
(d) 4.0 QUALIFICATION, CERTIFICATION, REQUALIFICATION, AND
RECERTIFICATION
The ability to weld or inspect to the requireme nt of this specification shows the person, process, and equipment are qualified. "Certified," means the weld er, welding procedure, and welding machine is qualified to make welds, and can make welds, to this specification.
4.1 Qualification requireme nt.
4.2 Equipment, machines, and process procedures should consist of such things as requir ed tooling, welding, schedule, pre-heat cycles, post-heat cycles, and required calibration of equipm ent.
(e) 5.0 MATERIALS
5.1 List welding rod and electrodes.
5.2 Gases.
5.3 Storage of rods and electrodes.
(f) 6.0 TECHNICAL
6.1 Inspection and manufacturing requirements.
6.2 Pre-heat and post-heat requirem ents if needed.
6.3 Heat treatment. This paragraph should cover any heat treatm ent required to restore all material in a weldment to a certain temper or tensile strength.
6.4 Detail gap - joint design information.
6.5 Describe weld repairs if applicable.
6.6 Cleaning, pre- and post-welding.
(g) 7.0 MANUFACTURING AND QUALITY CONTROL
7.1 Manufacturing requirements.
7.2 Quality control requirements.
m. Flutter.
(1) Regulation Reference. Section 23.629.
(2) Discussion. For general guidance on this section see AC 23.629-1A.
(3) Regulation Reference. Section 23.629 (b).
(4) Discussion. All airplanes certified to Title 14 CFR part 23, in any category, must perform flight flutter test(s) to show compliance in addition to any analytical methods.
(5) Regulation Reference. Section 23.629(d).
(6) Discussion. Airplanes that are the subject of this AC would typically meet the requirements of § 23.629(d)(1) through (d)(3). However, the applicant should review the criteria to ensure applicability. Based on the review of criteria, Engineering Report No. 45 (Simplified Flutter Prevention Criteria) could then be used to substantiate this portion of the regulations.
(7) Regulation Reference. Section 23.629(f)(2).
AC 23-15A, Change 1 3/15/05
(8) Discussion. For small low performance airplanes, designing the actuating structures to a factor of safety equaling four and providing redundant fastener safety as a means to minimize loss of single fastener joint integrity would be an acceptable method of showing compliance with this request.
n. Proof of Strength (Wings).
(1) Regulation Reference. Section 23.641.
(2) Discussion. Section 23.641 requires proof of strength of a stressed skin wing through tests or by combined structural analysis and load tests.
(a) Proof of strength of conventional aluminum or wood stressed skin wing structures primarily by analysis to limit and test to ultimate loads for the most critical bending and torsional wing load conditions is acceptable.
(b) In order to determine the most critical bending and torsional wing loads, one must complete a loads survey of the structure to determine which load case(s) produce the lowest safety margin against catastrophic failure of the wing. The actual mechanics of how to determine those specific load conditions requires qualified and experienced engineering personnel. The critical load conditions must be determined for the specific structural and airplane configuration, by detailed load and stress analysis. A specific explanation of this procedure that would be applicable for every airplane design doesn't exist.
(c) For conventional aluminum materials, the ultimate strength is typically less than one and a half times the yield strength. This means that as long as the article doesn't fail at the ultimate load, it will not yield at limit load. This rationale is for the worst-case scenario where the part fails in tension. For most structures however, failures occur due to instability first. The allowable ultimate load capability of a member in compression is lower than the tension capability.
This further reduces the difference between the ultimate and yield strength. Structural analysis should still be used to demonstrate that the yield strength is not exceeded at limit loads. If nonconventional aluminum is used, where this yield-to-ultimate strength relationship doesn't exist, then testing to ensure no yielding occurs at limit loads may be in order, at the ACO's discretion.
(d) In any event, part 23.307 is still app licable. Section 23.307 states in part that “Structural analysis may be used only if the structure conforms to those for which experience has shown this method to be reliable.” If in the ACO's judgment, the structure or analysis techniques used by the applicant, will not produce reliable analytical data, then testing may be in order.
(e) For wooden structures, there are no we ll-defined (published) yield strength values. When you observe the stress-strain curves for wood, it is evident that the relationship is not linear up to failure. The mechanism, that appears as yielding on the stress-strain curve, is actually micro cracks that develop within the matrix (lignin) between the fibers (cellulose). [3/15/05] These micro cracks have a negligible effect on the final ultimate strength [or—3/15/05] stiffness, of the material. Experience has shown that the ultimate strength test along with analysis is adequate to show compliance.
o. Landing Gear .
(1) Regulations Reference . Sections 23.723, 23.725, 23.726, and 23.727.
(2) Discussion .
(a) The designer selects the landing load factor, and the applicant is required to dem onstrate that the airplane will not exceed this factor when the aircraft is landing at the descent velocity specified by the regulations (reference § 23.473). For design purposes, this load factor may not be less than 2.67 and must not be exceeded by the certification drop tests specified in §§ 23.723 through 23.727.
(b) Alternatively, a conservative approach may be chosen by using a load factor of
4.2 g's to directly d etermine the static loads to be applied through design or static test to the landing gear. This load factor is considered conservative enough that it is unnecessary to verify it is not exceeded by an instrumented drop test. If this method (load factor of 4.2g's) is used, the only dynamic test required is the most critical drop test for reserve energy either in a fixture or by dropping the airframe, as described in § 23.727, without the need for instrumentation.
p. Brakes .
(1) Regulation Reference . Section 23.735.
(2) Discussion .
(a) A conservative rational analysis (§ 23 .735(a)(1)), or Technic al Standard Order (TSO) type tests (§ 23.735(a)(2)) and takeoff power test (§ 23.735(b)) are adequate for this class of airplane. (A rational analysis may include data such as service history, similarity to other approved brakes, etc., sufficient to increase the FAA's confidence level prior to installation/flight testing.)
(b) Additionally, taxi and landing stop test s to dem onstrate acceptable longitudinal and directional stability and control, and that no unacceptable vibrations, squeal, fade, grabbing, or chatter are present, are required in accordance with §§ 23.75, 23.143(a)(6), 23.231(a), 23.233(c),
23.493(c), 23.1301(d), and 23.1309(a)(1).
NOTE "Acceptable" means controllable by a pilot of any skill leve l.
q. Compartment Interiors/Powerplant Instruments .
(1) Regulations Reference . Sections 23.853 and 23.1337.
(2) Discussion .
(a) Section 23.853(e) does not prohibit direct reading oil pressure gauges, but it must be ". . . adeq uately shielded, isolated, or otherwise protected so that any breakage or failure of such an item would not create a hazard." Additionally, § 23.1337(a)(2)(i) and (ii) state that, "Each line . . . must have restricted orifices or other safety devices at the source of pressure to preven t the
30 escape of excessive fluid if the line fails; and be . . . located so that the escape of (such) fluid would not create a hazard."
(b) A direct oil pressure gauge line located in the cabin and f itted with an orifice of approximately 0.060-inch diameter located at the engine fitting limiting the flow of the oil into the cabin in the event of an oil line rupture inside the cabin is acceptable without a shrouded oil line vented overboard.
r. Fire Protection of Flight Controls, Engine Mounts, and Other Flight Structure.
(1) Regulation Reference . Section 23.865.
(2) Discussion . Two methods are available to validate that flight controls, engine mounts, and other flight structure can satisfy this requirement:
(a) Verify that the proposed components use a known fireproof material (example, steel or stainless steel). Or, as an alternative, a conformity inspection comparing the proposed assem bly to an existing Parts Manufacturer Approval (PMA) or product that is used in an existing type certificated aircraft can be accomplished.
(b) The applicant can validate components not previously approved by performing a fireproof test. Procedures for conducting such tests are contained in AC 20-135. A proposed test plan would need to be reviewed and approved b y the FAA prior to testing unless an existing TSO validation procedure is already published. The test would use a conformed control; a 2,000-degree Fahrenheit flame would be impinged onto the component and held in place for 15 minutes for fireproof compliance. Before removal of the flame at the termination of the test, loads should be applied to the component to show that it can still perform its intended function.
NOTE This portion of the regulations does not address engine controls. See part 23, § 23.1141(f) for requireme nts regarding the fire resistance testing for engine controls.
Also, the reader may refer to AC 23-17A, “Systems and Equipment Guide for Certification of Part 23 Airplanes” for more information.
s. Lightning Protection of Structure, Fuel System Lightning Protection, and S ystems.
(1) Regulations Reference . Sections 23.867, 23.954, and 23.1309(e).
(2) Discussion .
(a) Lightning testing is not required for those airplanes intended for operation under Visual Flight Rules (VFR) only. A general engineering overview of the airplane construction features to determine that any lightning induced hazards are minimized may be all that is required, if it can be determined that sufficient conductivity is available throughout the airplane and that the design employs features that have been found to minimize the hazard to the airplane and occupants.
Refer to FAA report number DOT/FAA/CT-89/22, “Aircraft Lightning
31 Protection Handbook” for best practices. The Type Certificate Data Sheet (TCDS) will contain this statem ent: "This aircraft is not approved for Instrument Flight Rules (IFR) operation under the provisions of § 91.205(d)."
(b) For those airplanes intended for opera tion under IFR a m ore detailed assessment will need to be made as to the vulnerability to lightning related hazards. Sections 23.867, 23.954, and 23.1309(e) of part 23 of the 14 CFR will need to be individually addressed. The depth of the verification should be commensurate with the degree of hazard. Since airplanes that employ conventional service proven construction and systems (i.e., all-aluminum airplanes with gyro flight instruments, magneto type engine ignition systems, mechanical fuel systems, etc.) have demonstrated over many millions of flight hours excellent inherent lightning protection qualities, similar designs may only need a qualitative engineering assessment that all hazards have been reasonably minimized. The Aircraft Lightning Protection Handbook, DOT/FAA/CT-89/22 may be used as a guide to make this assessment.
(c) Airplane designs that employ all composite construction, electronic flight instrum ent displays, electronic ignition systems, electronic fuel systems or electronic engine controls are generally more susceptible to lightning threats and, require further investigation. Therefore, airplanes that employ these designs are not within the scope of this AC. For electronic systems, see AC 20-53A, AC 20-136, and RTCA DO-160D for additional information. AC 23.1309-1C provides an acceptable means of identifying and assessing complex critical and essential systems.
(d) Airplanes using the foregoing means of compliance shall display the following lim itation placard: Not approved or eligible in Instrument Meteorological Conditions (IMC).
(e) See AC 23-17A “Systems and Equipmen t Guide f or Certification of Part 23 Airplanes” for more information.
t. Fuel Tank Tests .
(1) Regulation Reference . Section 23.965.
(2) Discussion . Certain requirements as defined in § 23.965(b) and (c) would be unduly burdensome in cases where certain small aluminum or nonmetallic fuel tanks incorporate side stiffeners or beads and internally attached baffles. The FAA tests described in § 23.965(a) satisfy the requirements if the tank has less than a 10-gallon capacity, or the tank interior has baffles (or equivalent support) at least every 15 inches in both directions.
u. Avionics Installation .
(1) Regulations Reference . Sections 23.1301, 23.1431, 23.1309(a), and 23.1311.
(2) Discussion .
(a) Section 23.1301 requires that all installed equipment perform its intended functions. For comm unication and navigation equipment required for IFR operations, an acceptable means of performance evaluation of these systems is provided in AC 23-8B, chapter 5.