Climbs - Equivalent Altitude Method Wing Span, b 33.30 Oswold Efficiency Factor, e 0.80 Indicated Pressure Altitude 3750 [recorded] 2 Calibrated Pressure Altitude 3750 [this example assumes that there is no difference between ind. And cal.] OAT. CAT, deg F 68 [recorded OAT. CAT, deg C 20 [from A/S calibration charts] Indicated Airspeed 68 recorded Calibrated Airspeed 70 [from A/S calibration charts] Observed R/C, ft/min 442 [altitude delta / time (minutes)] Std temp @ Alti, deg F (line 2) 46 [59 - .00356616 * line 2] ABS OAT, deg RI 528 [459.7 + line 3] 10 ABS Std. Temp., deg R 505.24 [518.688 - (.00358616* line 2] 11 Temperature correction 1.04 [line 9 / line 10] 12 Density Altitude @ line 2 and line 3 5.187 (1-6.87535 * 10* * line 2) 32561 * (518.688 / (line 3 + 459.7)1:231 - 1]/(-6.57535 * 10* 13 14 Density @ line 12 (1 / Sq. root density ratio) at line 12 0.002121 1.1096 11/(518.888 (1= (8.87535 * 10) * line 2) 82581) / (line 3 + 459.688) °3 0.0023769 * (1-(6.87536 * 104) * line 2) 4.28411 15 Test airplane wt. 1745 [recorded] 16 Weight Correction 0.99 WSTD/line 15] 17 18 Dynamic Pressure True Airspeed 13.81 78 [line 8 + Line 14] 10.5 + line 13 ‡ (line 17 + 1.467)?l 19 Calculation Factor 38471 ¡line 18 ‡ pie + e ‡ b'] 20 Ind. Drag @ Test Wgt 79 [line 15'/ line 19) 21 Del Di due to Wgt -1.35 line 20 ‡ (line 16' - 1.0)] 22 R/C corr. For Wgt and Temp. 488 [line 7 ÷ line 11)/line 16] 23 Del R/C due to del Di -5.36 [(line 21 = line 17 + 88)/W.l 24 Equivalent Altitude 4267 [line 2 - (0.36 ÷ (line 2 - line 12)] 25 R/C [Std. Atm. @ Alt. line 24] ft/min 471 [line 22 - line 23] Standard Day R/C Expanded to Approximate Effects of Altitude and Temperature Example is for 5000 ft at ISA plus 20 degrees C Pressure Altitude 5000 OAI SI deg C 5.15 [(line 1 + (-0.002) + 15.15] OAT S, deg F 41.27 [(line 9 + line 2) / 5) + 32] 4 OAT, deg C 25.15 [This is line 2 plus or minus the temperature value selected. For this example it is +20] OAT, deg F 77 [(9 + line 4) / 5) + 32] ABS OAT SI deg R 501 [line 3 + 459.7] 7 ABS OAT, deg R 537 [line 5 + 459.7] Temp. correction: Pressure Ratio 0.871714 0.93 11- (6.87535 + 10) + line 2) 2587 [line 6 / line 7] 10 Density @ Alt, and OAT in lines 1 and 0.001933 [line 9 ÷ 0.002377 ‡ line 10] 11 Alt. density @ Alt, and OAT in lines 1 and 7280 [Look up from std. Atmosphere tables.] 12 Equivalent Altitude 5821 [Go to the std. Day climb performance plot and use the performance at 5821 ft for the Performance at 5000 ft and 20 deg C above ISA (warmer day)] Expand for the Approximate Effects of Weight: (This is optional for airplanes under 6000 pounds unless they have a turbine engine) 13 Density Ratio 0.813272 [line 11 / 0.002377] 14 1 / (sq root density ratio) 1.108874 [1 / (line 14) °3] 15 Airplane Weight, Selected 1500 [self-explanatory] 16 Weight Correction 1.15 [Gross Weight / Line 16 selected weight] 17 _ V, mph 61.00 [recommended climb speed] 18 19 20 Draginduced @ Selected Weight Dynamic Pressure TAS 104.36 67.64 7.74 [line 18 ÷ line 15] Tine 16' / (3.14 *b? * e * line 20)] 10.5 + line 11 + (line 18 " 1.467"] 21 22 Del Drag Due to Weight Delta Rate of Climb, ft/min 34.45 96.44 I(2 / (3.14 ‡ b' + e + line 11 + (line 19 + 1.487) + (1.737) = (line 18) / 1730) ‡ 80] [line 21 ÷ (line 17) - 1)] Figure 9 - DATA REDUCTION FORM CLIMBS - EQUIVALENT ALTITUDE METHOD 13
2. The essential limitation to this method is that the correction to the observed climb for power, when the outside air temperature is not standard, is predicated upon the assumption that all engine controls are fixed. That is, no correction for variation in throttle setting, mixture, carburetor heat, etc., is made, and a climb at any part throttle setting is corrected for that throttle setting but not to full throttle. This method is referred to as the “equivalent altitude method” since it employs a correction to climb performance test data for atmospheric temperature variations from “standard” which indicates that the performance obtained under any atmospheric temperature and pressure may be obtained at some “equivalent altitude” in the standard atmosphere. The “equivalent altitude” may for the purposes of this AC be expressed as the “pressure” altitude plus 0.36 times the algebraic difference between the “density” altitude and the observed “pressure” altitude (“equivalent” altitude will always lie between the “pressure” altitude and the “density” altitude). As mentioned above, when climb performance is referred to “equivalent” altitude, no further correction for the effect of power changes needs to be made, and the final climb obtained thereby is that which would be obtained in the “standard atmosphere” at a height equal to the “equivalent” altitude when the power plant is operated as it was in the actual test.
(b) Density Altitude . Data reduction procedures using the density altitude method are an acceptable method of compliance for small airplanes with constant speed propellers. This method is outlined in AC 23-8B. The data reduction form shown in figure 10 may be employed.
This method also is applicable to both saw tooth and continuous climb data sets.
FORM ACE-1762 DEPARTMENT OF COMMERCE FORMERLY FAF-17 CLIMB CORRECTION - SHEET 1 CIVIL AERONAUTICS ADMINISTRATION (1-2-47)
AIRPLANE ENGINE PROPELLER PAGE
MODEL MODEL MODEL STD. WT. RATING RPM BLADE PROV. WT. METO RPM WING AREA GEAR RATIO NO. BLADES REPORT NO.
SPAN "E" MIN. OCTANE TYPE "F" ENG. COWL FLAP OPENING CRIT. ALT. DIA FLIGHT NO.
(RATED) CONFIGURATION CRIT. ALT.
(TEST) NO. ITEM CLIMB CLIMB CLIMB CLIMB CLIMB NO. NO. NO. NO. NO.
INDICATED AIRSPEED (MPH)
TRUE INDICATED AIRSPEED (MPH) VP/PO TRUE SPEED = (2) /13) (MPH)
TIME (MIN) OBSERVED CLIMB (FT.)
7 OBSERVED RATE OF CLIMB R/C (FT./MIN.)
AVERAGE PRESSURE ALTITUDE Ho (FT.)
OBSERVED OUTSIDE AIR TEMP. (°F) 10 ABSOLUTE OBS. OUTSIDE AIR TEMP. (°F) 11 STANDARD AIR TEMP. AT ALT. (8) (°F) 12 ABS. STAND. AIR TEMP. AT ALT. (8) (°F) 13 DENSITY ALTITUDE HD (FT.)
14 STAND. AIR TEMP. AT ALT. HD (13) (°F) 15 ABS. STAND. AIR TEMP. AT HO (13) (°F) 16 OBSERVED CARB. AIR TEMP. (°F) 17 ABSOLUTE CARB. AIR TEMP. (°F) 18 CARB. HEAT RISE = (16) - (9) (°F) 19 OBSERVED ENGINE RPM (RPM) 20 OBSERVED MANIFOLD PRESSURE (in. Hg) 21 ACTUAL WEIGHT (LBS.)
Figure 10 - DATA REDUCTION FORM CLIMBS - DENSITY ALTITUDE METHOD (SHEET 1 OF 2) 15
FORM ACE-1762 FORMERLY FAF-17 (1-2-47) CLIMB CORRECTION - SHEET 2 NO. ITEM CLIMB NO. CLIMB NO. CLIMB NO. CLIMB NO. CLIMB NO.
22 ABS. OUTSIDE AIR TEMPERATURE _ (10) ABS. STAND. AIR TEMP. (Hp) 23 ABS. STD. AIR TEMP. (H) = 17) ABS. CARB. AIR TEMP 24 ABS. STD. AIR TEMP. AT Ho V ABS. STD. AIR TEMP. AT HD + CAHR - V (15) + (18) 25 WT. TO BE CORRECTED TO (WE) ACTUAL WEIGHT 26 (WT. TO BE CORRECTED TO? -1-5-1 (ACTUAL WEIGHT)?
27 ENGINE POWER FROM POWER CURVE (HP) 28 ACTUAL BRAKE HORSEPOWER = (27) X (23) (HP) 29 STD. BHP ATTAINABLE AT HO (HP) 30 BHP OBTAINABLE AT HD = (29) X (24) (HP)
ACTUAL RATE OF CLIMB OF TEST AT TEST 31 WT. & STANDARD COND. = (7) X (22) (FT./MIN.)
32 n - PROPULSIVE EFFICIENCY 33 BHPA CHANGE - ACTUAL TO THAT AT HD = [(30) - (28)] 34 THP A CHANGE-ACTUAL TO THAT AT HD = (32) X 1HP) 35 INDUCED POWER REQ. _ AT TEST WT. (HPRI) 3.013eb? × (3) × (2) (21)? (HP)
INCREMENT OF INDUCED POWER 36 REQUIRED CHANGE-FROM TEST WT. TO FINAL WT. AT CONSTANT SPEED = - (35) X (26) (HP)
THP CHANGE DUE TO HP CHANGES OF 37 (34) + (36) AT FINAL WT. = (34) + (36) (HP)
R/C CHANGE DUE TO HP CHANGES 38 OF (37) A T FINAL WT. = (37) x 33,000 (FT./MIN.)
39 TEST R/C COR. TO WF = (31) X (25) (FT./MIN.)
40 FINAL R/C AT WF + HD(STAND) = (38) + (39) (FT./MIN.)
Figure 10 - DATA REDUCTION FORM CLIMBS - DENSITY ALTITUDE METHOD (SHEET 2 OF 2) 16
(c) Engine Power Correction . These methods make no direct engine power corrections for the difference between calibrated and rated power. The determination of the "equivalent" altitude empirically adjusts the data to allow for atmospheric effects on power for fixed pitch propeller installations. If further corrections for engine power (differences between calibrated and rated power) are desired, the following process may be utilized for either the equivalent or density altitude method:
1 . The airplane climbs because a surplu s of power is available in accordance with the relationship, R C BHP Wpe x=× ×η 33000 where, R C = rate of climb, feet/minimum ηp = propulsive efficiency (.75 ma y be used for constant speed propellers and .65 for fixed pitch propellers if no better value is known) BHPex = excess brake horsepower available W = aircraft weight, pounds
2 . To develop a power correction, the above equation is solved for,
() ( )BHP R C Wex p=× ÷ × η 33000
where, R C = sea level rate of climb developed in flight test W = airplane approved gross weight
Also, BHP BHP BHPex avail req= −
where, BHPavail = maximum power available at sea level at the manifold pressure and engine rpm, which can be achieved.
BHPreq = power required for level f light at the climb speed.
The power correction must be applied to the BHPavail term. For example, the calibrated engine produces 101 percent rated power, then the BHPavail must be reduced by 1 percent.
Therefore, ()BHP BHP CORR BHPex avail reqcorr =− −
where, BHPexcorr = excess power remaining following the power correction power correction to available power CORR =
18 Using the corrected excess power, resolve the original equation to obtain the corrected rate of climb,
R C BHP Wcorr pe x corr=× × ÷η 33000
(d) Data Requirements . In order to provide AFM data showing the effect of altitude and temperature as required by § 23.1587, a sufficient number of climbs through a range of speeds must be completed allowing rate of climb to be plotted against equivalent or density altitude (whichever is appropriate). This relationship is then used to produce the AFM performance information. It is also necessary to perform climbs at different speeds in order to determine the best rate of climb speed Vy.
d. Landing .
(1) Regulations Reference : Sections 23.75 and 23.1587.
(2) Discussion .
(a) Measurement Methods . Section 23.75 requires the measurement of the distance required to land over a 50-foot obstacle and come to a stop. The methods described for takeoff measurement are equally applicable for landing performance measurements.
(b) Data Reduction . A number of ways exist to correct the test data obtained above to standard conditions at sea level. One method that has been used is shown in figure 11. These forms are oriented for reduction of test data obtained by photographic methods but may be adapted to other data collection methods.
TITLE PAGE PREPARED BY DATE REPORT NO.
CHECKED BY DATE MODEL
CORRECTION OF LANDING DATA TO STANDARD CONDITIONS LANDING NUMBER 1 4 1 W† = TEST WEIGHT (LBS.)
2 W$ = WEIGHT TO BE CORRECTED TO (LBS.)
W¡/Ws = (1) / (2) 4 Hp = PRESSURE ALTITUDE @ GROUND DURING TEST (FT.)
OAT = TEMPERATURE AT GROUND DURING TEST (°F) HD = DENSITY ALTITUDE (FT.)
7 p = DENSITY @ (4) + (5) (SLUGS/FT.)
8 PF = DENSITY TO BE CORRECTED TO (SLUGS/FT.)
9 (7)/ 8) 10 TA = AIR TIME FROM 50 FT. TO GROUND (CAMERA DATA) (SEC) 11 Vwo = OBSERVED WIND VELOCITY (FT/SEC.)
*12 VWH = 1.376 (11) = WIND VELOCITY AT 50 FT. (FT./SEC.)
13 VHO = OBSERVED VELOCITY @ 50 FT. (CAMERA DATA) (FT./SEC.)
14 VH = (12) + (13) = TRUE AIRSPEED AT 50 FT. (FT./SEC.)
15 Vco = OBSERVED CONTACT VELOCITY (CAMERA DATA) (FT./SEC.)
16 VCA = (11) = (115) = TRUE AIRSPEED AT CONTACT (FT./SEC.)
17 VCA/VCO = (16) / (15) 18 (VCA/Vco) 1.85 = (17) 1:85 = WIND CORRECTION FOR GROUND RUN 19 VVH = OBSERVED VERTICAL VELOCITY @ 50 FT. (CAMERA DATA) (FT./SEC.)
20 1.210 (14) + (16) 21 .00575 (14) x (20)
(19) 22 1.185 (10) 1/7 * Assuming VWH = Vwo Figure 11 - DATA REDUCTION FORM - CORRECTION OF LANDING DATA TO STANDARD CONDITIONS (SHEET 1 OF 2) 19
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: