Cessna 150/Lycoming O-320-E2D Limited Performance Evaluation

TABLE OF CONTENTS Page No. PREFACE ui EXECUTIVE SUMMARY v LIST OF ILLUSTRATIONS » LIST OF TABLES ™ INTRODUCTION 1 Background * Test Objective 1 Test Item Description * TEST AND EVALUATION 3 General 3 Cruise Performance 3 Test Objectives 3 Test Procedures 3 Test Results 3 Flight Test Data Reduction 3 RPM Model Generation 5 Predicted Aircraft Performance 5 Pitot-Static Calibration 7 Test Objectives 7 Test Procedures 8 GPS Speed Course Method 8 GPS Ground Speed Method 8 Test Results 8 Climb Performance ^ Test Objectives 10 Test Procedures 10 Test Results 10 Descent Performance 12 Test Objectives 12 Test Procedures 12 Test Results 12 vii Takeoff Performance 13 Test Objectives 13 Test Procedures 13 Test Results 14 CONCLUSIONS AND RECOMMENDATIONS 15 REFERENCES 17 APPENDIX A - TEST DATA 19 APPENDIX B-FLIGHT MANUAL INPUTS 103 APPENDIX C - AIRCRAFT MODELING 117 Matching RPMModd to Flight Test Data 118 Propeller Model Adjustment 118 Engine Model Adjustment 118 Fuel Flow Adjustment 119 Full Throttle Modeling 120 Rate of Climb Adjustment 120 Aircraft Model File C150150.ACF 127 Engine Model File LO320A.ENG 130 Propeller Model File C150150.PRP 134 APPENDIX D - FLIGHT TEST TECHNIQUE AND DATA REDUCTION DETAILED DESCRIPTION 135 Cruise Performance 136 Test Procedures 136 Data Reduction Methods 137 Data Presentation 139 True Airspeed and RPM at Non-Standard Conditions 139 Fuel Flow at Non-Standard Conditions 140 Range and Endurance at Non-Standard Conditions 141 Pitot-Static Calibration 141 Test Procedures 141 GPS Speed Course Method 141 Data Reduction Methods .' 142 GPS Ground Speed Method 144 Data Reduction Methods 144 Climb Performance 147 Test Procedures 147 Data Reduction Methods 147 Data Presentation 147 Vlll r Descent Performance 148 Test Procedures 148 Data Reduction Methods 148 Data Presentation 1*0 Takeoff Performance 15° Test Procedures 15° Data Reduction Methods 15° Data Presentation 152 LIST OF ABBREVIATIONS AND SYMBOLS 158 IX This page intentionally left blank. Figure 1 LIST OF ILLUSTRATIONS Title Page No. Cessna 150 General Arrangement 2 APPENDIX A Al Engine Horsepower Determination Methods Comparison 20 A2 Drag Polar Curve Fit 20 A3 N557TH Drag Results Compared to Aircraft Drag Polar 21 A4 N557AW Drag Results Compared to Aircraft Drag Polar 21 A5 N557SH Drag Results Compared to Aircraft Drag Polar 22 A6 Brake Horsepower Required Curve Fit 22 A7 N557TH Brake Horsepower Required Results Compared to Aircraft Brake Horsepower Required " A8 N557AW Brake Horsepower Required Results Compared to Aircraft Brake Horsepower Required z A9 N557SH Brake Horsepower Required Results Compared to Aircraft Brake Horsepower Required 24 A10 Brake Specific Fuel Consumption Results 24 All Specific Air Range Results 25 A12 Specific Endurance Results 2^ A13 Cruise Airspeed Performance 26 A14 Cruise RPM Performance 26 A15 Cruise Fuel Flow Performance 27 A16 Dual Constant Airspeed Cruise Range Performance 27 A17 Dual Constant Airspeed Cruise Range Performance With 45 Minute Reserve 28 A18 Dual Constant Airspeed Cruise Endurance Performance 28 A19 Dual Constant Airspeed Cruise Endurance Performance With 45 Minute Reserve 29 A20 Solo Constant Airspeed Cruise Range Performance 29 A21 Solo Constant Airspeed Cruise Range Performance With 45 Minute Reserve 30 A22 Solo Constant Airspeed Cruise Endurance Performance 30 A23 Solo Constant Airspeed Cruise Endurance Performance With 45 Minute Reserve 31 A24 Aircraft Brake Horsepower Required and Available 31 A25 Aircraft Thrust Horsepower Required and Available 32 A26 Aircraft Thrust Required and Available 32 A27 Airspeed Pitot-Static Position Correction 33 XI A28 Altitude Pitot-Static Position Correction 33 A29 Airspeed Pitot-Static Position Correction, N557SH 34 A30 Standard Day Rate of Climb Performance (Indicated Airspeed) 34 A31 Standard Day Rate of Climb Performance (Calibrated Airspeed) 35 A32 Test Day Rate of Climb Matching 35 A33 Test Day Rate of Climb Matching 36 A34 Nonstandard Day Rate of Climb Performance at 65 KIAS 36 A35 Nonstandard Day Time and Fuel to Climb at 65 KIAS 37 A36 Nonstandard Day Distance to Climb at 65 KIAS 37 A37 Nonstandard Day Rate of Climb Performance at 80 KIAS 38 A38 Nonstandard Day Time and Fuel to Climb at 80 KIAS 38 A39 Nonstandard Day Distance to Climb at 80 KIAS 39 A40 Idle Descent Drag Polar Curve Fit 39 A41 Idle Descent Drag Polar 40 A42 Idle Descent Penetration Chart 40 A43 Idle Descent Polar Chart by Indicated Airspeed 41 A44 Idle Descent Polar Chart by True Airspeed 41 A45 Nonstandard Day Idle Rate of Descent Performance at 65 KIAS 42 A46 Nonstandard Day Time and Fuel to Descend at Idle at 65 KIAS 42 A47 Nonstandard Day Distance to Descend at Idle at 65 KIAS 43 A48 Nonstandard Day Idle Rate of Descent Performance at 107 KIAS, 2250 RPM 43 A49 Nonstandard Day Time and Fuel to Descend at 107 KIAS, 2250 RPM 44 A50 Nonstandard Day Distance to Descend at 107 KIAS, 2250 RPM 44 A51 Standardized Takeoff Ground Roll Performance 45 A52 Mean Takeoff Ground Run 45 A53 Takeoff 95th Percentile Dispersion 46 A54 Takeoff 99th Percentile Dispersion 46 APPENDIX C Cl RPMEngine, Model For Lycoming O-320-E2D 121 C2 RPM Propeller Model For McCauley TM7458/1C172; Thrust Coefficient 122 C3 ÄPA/Propeller Model For McCauley TM7458/1C172; Torque Coefficient 122 C4 RPM Propeller Model For McCauley TM7458/1C172; Power Coefficient 123 C5 ÄPA/Propeller Model For McCauley TM7458/1C172; Propeller Efficiency 123 C6 ÄPA/Model RPM Matching 124 XI1 C7 RPMModd Manifold Pressure Matching 124 C8 RPMModd Fuel Flow Matching (in gal/hr) 125 C9 RPMModd Fuel Flow Matching (in percent) 125 CIO RPMModd Full Throttle Manifold Pressure Matching 126 Cll Slipstream Effects on Rate of Climb I26 APPENDIX D Dl Determining Fuel Burn Amount for Cruise Test Points 154 D2 GPS Speed Course Distance Arcs 154 D3 GPS Speed Course Wind Drift Error 155 D4 Determining Leg Length for GPS Speed Course Test Points 155 D5 Ground Speed Variation in a Turn 156 D6 GPS Ground Speed Method Vector Diagram 156 D7 Finding Test Day Rate of Climb 157 D8 Finding Test Day Rate of Descent 157 xm This page intentionally left blank. xiv LIST OF TABLES Table Title Page No. 1 Range and Endurance Scenarios 6 2 Climb Speed Comparison (in KIAS (KCAS)) 10 3 Maximum Performance Climb Results 10 APPENDIX A Al Tabulated Cruise Data by Indicated Airspeed 47 A2 Tabulated Cruise Data by RPM 77 A3 Maximum Range Airspeed 101 A4 Range Results by Airspeed 101 APPENDIX D Dl GPS Ground Speed Method Example Data 145 xv This page intentionally left blank. xvi INTRODUCTION BACKGROUND This test program was requested by the 94th Flying Training Squadron (FTS) to collect flight manual performance data for the Cessna 150 flown by the United States Air Force Academy (USAFA) Cadet Competition Flying Team (CCFT). The Air Force Academy airfield elevation is 6572 feet. To improve performance for operating from the high elevation airfield at the Air Force Academy, the original 100 HP Continental O-200 engine was replaced with a 150 HP Lycoming O-320-E2D engine. As a result, the manufacturer's flight manual performance data were no longer applicable. The primary operational requirement for this test was determining engine fuel consumption. These data were necessary for determining range and endurance for flight planning. These data were also necessary for CCFT competitions, where fuel burn must be predicted within 10 percent for certain events. These tests were conducted by members of the USAFA Department of Aeronautics. Aircraft scheduling and maintenance was performed by the 94th FTS. This program consisted of 24 flights totalling 41.0 flight hours during the period of 3 July 1995 to 16 May 1996. Primary flight testing was conducted in the local area around the Air Force Academy. Additional flight testing to verify results at lower altitudes was accomplished between the Air Force Academy and Hays, Kansas. TEST OBJECTIVE The general objective of this evaluation was to determine the modified Cessna 150 performance characteristics for purposes of generating flight manual performance charts. Areas included were pitot-static calibration, and cruise, climb, descent, and takeoff performance. All objectives were met. TEST ITEM DESCRIPTION The Cessna 150, as operated by the USAFA CCFT, is a two-place general aviation airplane. A three-view drawing of the aircraft is shown in Figure 1. It is powered by one normally aspirated, carbureted, 4-cylinder, 150 horsepower Lycoming O-320-E2D engine driving a MacCauley TM7458/IC172 fixed- pitch propeller. The high wing has an area is 160 square feet and an aspect ratio is 7.0. The maximum takeoff gross weight was 1760 lbs. The flight control system is a reversible flight control system. Each Cessna 150 operated by the USAFA CCFT is considered representative of the other two. These Cessna 150s are not considered production representative of unmodified Cessna 150s. Reference 1 has a more complete description of the CCFT Cessna 150. nDDnnniiiiiinn1111111 dvsn 8' - 6" MAX I°DnniriiiiiniiiIIjIIIIIp| -T 7 1/4 Figure 1 Cessna 150 General Arrangement TEST AND EVALUATION GENERAL Pitot-static calibrations and takeoff, cruise, climb and descent performance tests were conducted. Production instrumentation was used for measuring airspeed, altitude, outside air temperature (OAT), and RPM. A manifold pressure guage was added for flight test in place of the VOR 2 head. A fuel flow/used indicator was installed for use in normal operations. Time was recorded using a digital wristwatch and a Hewlett Packard 48SX calculator. Position was determined from a Garmin GPS 55 handheld Global Positioning System (GPS) receiver. No additional calibration of instruments was accomplished beyond normal Federal Aviation Administration (FAA) Instrument Flight Rules (IFR) requirements. Flight testing was accomplished in the USAFA local flying areas at pressure altitudes of 8,000 to 12,000 feet. Additional data were collected at lower altitudes near La Junta, CO (6,000 feet), Hays, KS (3,000 feet), and on routes between USAFA, CO and Hays, KS (4,500 - 7,500 feet). Flight test data were used to create and verify a computer model of the Cessna 150 aircraft, engine, and propeller using the Reciprocating Engine and Propeller Modeling Program (RPM). (Reference 2) This computer model was used to expand standardized data to create the flight manual data. Details of the construction of the aircraft model are contained in Appendix C. CRUISE PERFORMANCE Test Objectives; The test objectives for cruise performance were: 1. Determine power settings, fuel flow, range, and endurance as functions of airspeed and altitude. 2. Determine power required and power available as a function of airspeed. 3. Determine the aircraft drag polar. 4. Determine airspeeds for maximum range and for maximum endurance. 5. Create charts and tabulated data for Flight Manual cruise data. Test Procedures: Cruise data were collected using steady state trim shots at constant pressure altitude (PA) and airspeeds of 50, 60, 70, 80, 90, and 100 knots. Trim shots were also recorded at the airspeed for full throttle. Data recorded included the indicated altitude (hi), indicated airspeed (Vi), outside air temperature (OAT), manifold pressure (MAP), engine RPM, start and end time, initial fuel used and final fuel used. Fuel used was measured by a Hoskins FT101A Fuel Totalizer. Fuel used was reported to the nearest tenth of a gallon. This indicator normally displayed fuel flow, which varied too much to be usable for this test. This variation arose primarily from the actual variation in fuel flow as the carburetor float opened and closed the fuel inlet valve to the carburetor bowl. Fuel used could be read by pressing a button on the indicator. After displaying the fuel used for a few seconds, the display would revert to fuel flow. To improve the accuracy of the fuel used measurement, the flight test engineer pressed the button on the indicator each time the display reverted to fuel flow. This resulted in a reasonably constant display of fuel used. Timing for each test point was begun or ended as the tenths digit changed. Test points were flown long enough to burn exactly 0.5 gallons. Further information on fuel measurement can be found in Appendix D. Cruise data were reduced using the Piw - Viw method and other cruise data reduction methods shown in Appendix D. Test Results: Flight Test Data Reduction. Cruise data were collected as described above. Additionally, MAP and RPM data were used from Pitot-static test points. These data increased the amount of data available for determining drag and power required, but did not include fuel flow data. The drag and power required were determined using engine horsepower and propeller efficiency. The engine horsepower was determined from MAP, RPM, hi, and OAT, using the engine chart as described in Appendix C. Propeller efficiency was determined from RPM and true airspeed (V), using the propeller chart as described in Appendix C. N557TH was the only aircraft to have a MAP gauge installed, and not for all flights. On flights when a MAP gauge was not available, engine horsepower was determined using the propeller power coefficient. Figure Al compares the horsepower calculated by each method for test points where a MAP gauge was available. Ideally, all points would lie on the line with a slope of 1:1. The match between the two methods is satisfactory, especially considering that the propeller model is fairly simple, with only inputs of blade shape, diameter, and pitch. The drag polar was determined by a linear least squares curve fit to the drag coefficient (CD) versus the square of the lift coefficient (CL), as shown in Figure A2. This technique assumed a drag polar of the form CD = CD + KCL o with no linear CL term. The aircraft drag polar was CD = 0.042696 + 0.068861 CL2 While this drag polar reports more significant figures than can be justified from the flight test data, this drag polar is reported as the drag polar used in the RPM model. Given the aspect ratio of 7.0, this drag polar indicates an Oswald's efficiency factor of 0.66. The parasite drag coefficient is also higher than normally seen for general aviation aircraft. This difference is suspected to be caused by separation drag from the rear window. This separation is suspected again later when explaining the climb results. Figure A3 compares the drag results from aircraft N557TH, the primary test aircraft, with the drag polar. These data are in agreement with the drag polar. Drag data from two flights in N557AW are shown in Figure A4, again with good agreement. Drag data from N557SH were not originally in good agreement with the drag polar. After applying a different position correction to the N557SH data, as suggested from the Pitot-static tests, the agreement was improved to an acceptable level, as shown in Figure A5. Additional testing should be conducted to verify the validity of the drag polar for all three aircraft. (Rl)1 The brake horsepower required by a linear least squares curve fit to the product of standardized brake horsepower and standardized equivalent airspeed (BHPiwViw) versus standardized equivalent airspeed raised to the fourth power (Vjw4), as shown in Figure A6. This technique also assumes a drag polar of the form shown earlier and a negligible change in the propeller efficiency between the test data point and the standardized data point. Using this curve fit, the brake horsepower required was calculated, as shown in Figure A7 through Figure A9. These figures also show that the brake horsepower required results from each aircraft agree with the brake horsepower required curve in the same manner as the drag results with the drag polar. Figure A10 shows the fuel consumption results as brake specific fuel consumption (BSFC) plotted against brake horsepower. BSFC is normally considered as a constant with respect to brake horsepower. Thus, Figure A10 shows BSFC results for all cruise points, from pressure altitudes of 3,000 to 9,000 feet. The fairings are derived from RPM model data at cruise conditions at altitudes from sea level to 15,000 feet. According to the RPM model, the BSFC for a given brake horsepower will change slightly with altitude. However, all of these fairings are well within the boundaries of the scatter of the data, and are therefore considered reasonable. Figure All shows the results for specific air range (SAR) for all cruise data collected at all altitudes. Specific air range can be expressed as SAR = - 1 BSFC CD W V wf Numerals preceded by an "R" in parentheses at the end of a paragraph correspond to the recommendation numbers tabulated in the Conclusions and Recommendations section of this report. If the propeller efficiency (TIP) and the BSFC are considered constant with altitude for a given equivalent airspeed, then SAR should be independant of altitude. In practice and according to the RPM model, propeller efficiency does remain constant, but BSFC will vary slightly, as shown in Figure A10. Thus, the fairings derived from RPM model data at cruise conditions at altitudes from sea level to 15,000 feet show a variation similiar to the variation seen in the BSFC data. Again, all of the SAR fairings are within the boundaries of the scatter of the data, and are therefore considered reasonable. Figure A12 shows the results for specific endurance (SE) for all cruise data collected at all altitudes. Specific endurance can be expressed as SE = - c^ BSFC CD w% 1 Wf Since density appears explicitly in this equation, SE will be a function of altitude. Figure A12 does not attempt to break out the SE data by altitude, as the variation with altitude, shown by the fairings from ÄPA/data, is smaller than the scatter of the data. Any adjustments to improve the fit of SE data would be accomplished by improving the fit of the fuel flow data, which also affects BSFC. If the modeling of the fuel flow data is satisfactory, then the modeling of SE and BSFC will be satisfactory. Again, all of the SE fairings are within the boundaries of the scatter of the data, and are therefore considered reasonable. RPM Model Generation. A computer model of the performance of the aircraft was generated using the Reciprocating Engine and Propeller Modeling Program (RPM). The airframe was modeled using the drag polar derived from flight test data. The engine model and propeller model were adjusted until a satisfactory fit was obtained with flight test MAP, RPM, and fuel flow data. The process of this adjustment and the resulting model data files are described in detail in Appendix C. Predicted Aircraft Performance. The RPM model was used to create performance charts similar to those seen in the Flight Manual for a general aviation aircraft. Figure A13 shows the cruise true airspeed as a function of density altitude and power setting. To use this chart, start with the OAT, go straight up to the pressure altitude, go straight across to the power setting, then straight down to read the true airspeed. For the example shown: OAT: 40° F Pressure Altitude: 8,000 feet Power Setting: 60% True Airspeed: 102 KTAS Figure A14 relates RPM to power setting as a function of density altitude. To use this chart, start with the OAT, go straight up to the pressure altitude, go straight across to the power setting, then straight down to read the RPM. For the example shown: OAT: 40° F Pressure Altitude: 8,000 feet Power Setting: 60% RPM: 2345 Appendix D shows that true airspeed and RPM remain the same for an aircraft in cruise flight at the same density altitude, power setting, and weight. Thus, the density altitude can be used to account for non- standard conditions. Figure Al5 shows the cruise fuel flow as a function of pressure altitude, power setting, and OAT. To use this chart, start with the pressure altitude, go straight across to the power setting, then straight down to the zero temperature deviation line. Follow the guidelines (up for OAT above standard, down for OAT below standard) by the amount of temperature deviation from the standard temperature for the pressure altitude. Then go straight down to read the fuel flow. For the example shown: Pressure Altitude: 8,000 feet Power Setting: 60% OAT: Std + 60° F Fuel Flow: 8.6 gal/hr Appendix D discusses the correction to fuel flow for non-standard temperatures. Table Al tabulates the cruise performance for various altitudes, airspeeds, and temperatures. Values for manifold pressure, percent power, RPM, true airspeed, and fuel flow are given for each flight condition. This table was reproduced on 5x8" cards for the flight crews for flight planning purposes, without the manifold pressure and percent power information. Power settings above 75 percent were flagged. Cards were produced with data at each 1000 feet of altitude. Cards were also produced corresponding to Visual Flight Rules (VFR) hemispheric altitudes (each 1000 feet + 500 feet). Images of these cards are shown in Appendix B. Table A2 shows the same information as Table Al, except that data is arranged by RPM, not indicated airspeed. This format is similar to that used in Cessna Flight Manuals. Figure A16 through Figure A23 show the range and endurance for cruise at a constant indicated airspeed. Data are shown for dual and solo flight. Both conditions assume takeoff at maximum gross weight. Takeoff at less than maximum gross weight (with the same amount of fuel) would result in longer range and endurance. Table 1 details the assumptions for the dual and solo scenarios. Table 1 RANGE AND ENDURANCE SCENARIOS Parameter Dual Solo Empty Weight 1249 lbs 1249 lbs Aircrew 2 (340 lbs) 1 (170 lbs) Baggage 15 lbs 113 lbs Unusable Fuel 3 gal 3 gal Useable Fuel 23 gal 35 gal Startup, Taxi, Takeoff, and Climb Fuel 2 gal 2 gal Climb Distance 10 nm 10 nm Climb Time 8 min 8 min The range and endurance charts were created using the RPM model. After setting the aircraft weight and useable fuel, the fuel consumption was computed in 10 minute time intervals. Between each interval, the aircraft and engine were retrimmed to account for reduction in drag arising from the reduction in weight. This process was continued until all of the useable fuel was consumed. Figure A16 through Figure A23 include the climb distance and climb time shown in Table 1. No distance or time for descent were included in these charts. To use the range charts, start with the OAT, go straight up to the pressure altitude, go straight across to the indicated airspeed, then straight down to the zero temperature deviation line. Follow the guidelines (up for OAT above standard, down for OAT below standard) by the amount of temperature deviation from the standard temperature for the pressure altitude. Then go straight down to read the range. For the example shown: OAT: 80° F Std + 35° F Pressure Altitude: 4,000 feet Indicated Airspeed: 80 KIAS Range: 315 Appendix D explains the corrections to range for non- standard conditions. To use the endurance charts, start with the pressure altitude, go straight across to the indicated airspeed, then straight down to the zero temperature deviation line. Follow the guidelines (up for OAT above standard, down for OAT below standard) by the amount of temperature deviation from the standard temperature for the pressure altitude. Then go straight down to read the endurance. For the example shown: Pressure Altitude: 4,000 feet Indicated Airspeed: 80 KIAS OAT: Std + 35° F Endurance: 3.7 hr Appendix D discusses the correction to endurance for non-standard temperatures. Figure A24 shows the brake horsepower required and available at altitudes from sea level to 15,000 feet. This figure was created using the RPM model. To determine other performance parameters, propeller efficiency was applied to the curves of Figure A24 to calculate thrust horsepower required and available, shown in Figure A25. From this chart, thrust required and available were calculated, shown in Figure A26. Table A3 shows the airspeed for maximum range as determined by three methods at sea level and 10,000 feet. Assuming constant propeller efficiency and BSFC, maximum range for a propeller driven aircraft occurs at the airspeed for maximum L/D. (Reference 5) Thus, the airspeed for maximum range would be found at the minimum of the thrust required curve (Figure A26) or at the tangent from the origin to the thrust horsepower required curve (Figure A25). This method gives an indicated airspeed of 56 KIAS. Using the SAR shown in Figure All, the SAR is a maximum at 73 KIAS at sea level and 68 KIAS at 10,000 feet. These values of SAR are for maximum gross weight at a given standardized airspeed. As the weight decreases, the standardized airspeed corresponding to a constant indicated airspeed will increase. Increasing standardized airspeed from the airspeed for maximum SAR will reduce the SAR. Therefore, for an overall maximum range, the indicated airspeed would be less than the values indicated on this chart. Additionally, changes in propeller efficiency or BSFC as weight decreased would change the values of SAR. According to the range charts (Figure A16 and Figure A20), the maximum range occurs at 65 KIAS for both dual and solo flight. The range at 56 KIAS is not shown in Figure A16 since it was less than the maximum, but the range at 56 KIAS was only 8 to 11 nautical miles less than the range at 65 KIAS, depending on altitude. Assuming an approximate range of 300 nautical miles, this difference would be 3 to 4 percent. At 73 KIAS, Figure A16 suggests the difference in range to be 0 to 5 nautical miles shorter than the range at 65 KIAS. This difference would be under 2 percent. Thus, each method results in an airspeed giving a range within 4 percent of the other methods. Since this method used to create Figure A16 accounts for the changes in weight during cruise, 65 KIAS was probably the most accurate airspeed for maximum range, even though the change in range is very small with airspeed around 65 KIAS. Therefore, 65 KIAS was chosen as the airspeed for maximum range. Flying at maximum range airspeed is typically too slow for operational considerations. Table A4 shows the effect on range of flying at higher speeds at typical cruise altitudes of 5,000 and 10,000 feet. Reference 6 suggests that the airspeed for maximizing airspeed per amount of fuel burned, and thus the most efficient cruise speed considering time and fuel use, is found at the tangent from the origin to the thrust required line. Figure A26 shows this airspeed to be 85 KCAS. This airspeed corresponds to 86.5 KIAS. However, for ease of reading the range charts and operational simplicity (the airspeed indicator has a mark at 85 KIAS), this airspeed was investigated at 85 KIAS. Flying at 85 KIAS increases the airspeed by 20 KIAS, with only a 10 to 20 percent reduction in range. Table A4 also shows range performance for the typical operational technique practiced at the 94th FTS. Flight time has an operational cost since the maximum flight time allowable per day is limited, thus limiting the total range available per day. Additionally, flight time has a monetary cost in per diem payments for TOY aircrew. Since fuel cost is typically negligible compared to the cost associated with flight time, flights are typically conducted at maximum airspeed, either at 75 percent power or full throttle if 75 percent power is not attainable. At 5,000 feet density altitude, 75 percent power yields an indicated airspeed of 105 KIAS and a 34 to 40 percent reduction in range. At 10,000 feet density altitude, full throttle yields 90 KIAS and a 23 to 25 percent reduction in range. The airspeed for maximum endurance can also be determined three ways: the minimum power required, the maximum specific endurance, and the largest calculated endurance from the RPM model. In each of these cases (Figure A12, Figure A18, Figure A22, and Figure A25) the maximum endurance airspeed is shown to be the minimum speed tested, or 50 KIAS. Although slightly more endurance would probably be possible at a slower speed, 50 KIAS is the minimum practical endurance speed for holding, considering the Flight Manual reported stall speed of 47 KIAS. Flight data, such as RPM, fuel flow, fuel used, indicated airspeed, true airspeed, range, and endurance should be collected on 94th FTS deployments and CCFT practice and competition flights and compared to the performance data presented in this report for further verification of these performance data. (R2) PITOT-STATIC CALIBRATION Test Objectives; The test objectives for Pitot-static calibration were: 1. Complete a calibration of the production Pitot- static system. 2. Verify Pitot-static corrections given in the Flight Manual. Test Procedures: GPS Speed Course Method. This Pitot-static calibration method was an adaptation of the traditional ground speed course method (Reference 3). Instead of using landmarks to determine distance, GPS distance-to-go readings were used. These distance-to-go readings were based on a waypoint at least 30 tun away. This waypoint was chosen such that the heading directly toward or away from the waypoint would be approximately perpindicular to the wind. The aircraft was flown on a heading directly toward and away from the waypoint with no wind drift correction. For each airspeed tested, the time to fly four nautical miles (ground distance) was recorded in each direction. Additionally, hi, V;, OAT, MAP, RPM, and fuel used were recorded. The true airspeed was assumed to be equal to the average ground speed for runs toward and away from the waypoint. From this true airspeed the position correction was determined. For this testing, airspeed and altitude instrument errors were assumed to be negligible. A more complete description of this technique and the data reduction are shown in Appendix D. GPS Ground Speed Method. The GPS ground speed method was developed at the USAF Test Pilot School (USAF TPS), and became known to the test team during the flight test phase of this project (Reference 4). Additional Pitot-static testing was completed to compare the relative position errors of different CCFT aircraft, and at the request of USAF TPS for further development of this method. In this method, the aircraft true airspeed was estimated based on indicated airspeed, estimated position correction, pressure altitude, and outside air temperature. Starting on a heading with an expected headwind or tailwind, a slow turn was initiated. The turn was continued until the GPS ground speed matched the calculated true airspeed. At this point the aircraft should be heading perpindicular to the wind. The aircraft was then turned 180 degrees to confirm the same ground speed. These headings were then used for the data collection. The aircraft was flown at the aim airspeed and altitude on the crosswind heading. The primary data collected were V„ heading, GPS ground speed, and GPS track angle. Additionally, h;, OAT, MAP, RPM, and fuel used were collected. The primary data were recorded multiple times for approximately one minute to detect any variations from outside effects such as wind gradients. The same data were collected for the same flight conditions on the opposite heading. The true airspeed was determined by multiplying the GPS ground speed by the cosine of the angle difference between the heading angle and the GPS track angle (i.e. the drift angle). For this testing, airspeed and altitude instrument errors were assumed to be negligible. A more complete description of this method and the data reduction are shown in Appendix D. Test Results: Figure A27 shows the flight test derived position correction curve compared with the flight test data and the Flight Manual position correction curve. The flight test derived curve seems reasonable, as it follows the general trend of the Flight Manual curve. The flight test data shown in this figure were all collected in N557TH. Pitot-static data were collected in this aircraft on flights 1, 5, 6, 7, 10, 22, and 23. The flight 1 data are not shown as they were significantly different from all later flights and did not pass the reasonableness test. On flight 1, Pitot-static data were collected using the GPS Speed Course method, but with legs only one nautical mile in length. Prior to flight 5, the test team decided that legs of at least four nautical miles in length were necessary to reduce possible errors to an acceptable level. (See Appendix D) Pitot-static data was collected using the GPS Speed Course method on flights 5, 7, and 10. The data shown for flight 5 have a similar slope to the final position correction curve, but were displaced down from the curve by two to four knots. These data were weighted less heavily than the rest, since the leg times implied that the legs were not being flown perpindicular to the wind. Thus, the data was suspected to have been corrupted by wind effects. Starting with flight 7, more care was taken to ensure that legs were flown perpindicular to the wind. The data from flights 7 and 10 agreed well, and were used to create the position correction curve. The leg times showed that wind effects were minimal for these flights. No Pitot-static data points had been collected to this time at SO KIAS, so these were picked up on flights 22 and 23. These data were collected using the GPS Ground Speed method, which had become known to the test team by this point These two data points agreed very well with the slope implied by the previous data points at 60 and 70 KIAS. Further confidence was gained in the flight test derived position correction curve when drag polar data and power required data fell into the shapes expected from theory. Using prior position correction curves, such as the curves used based on flight 1 or flight 5 data, the drag polar and power required data did not follow the generally linear trend seen in Figure A2 and Figure A6. When the final position correction was used, the data lined up as shown in these figures with no further compensation. The CCFT normally plans its competition navigation legs at 90 KIAS. At this airspeed, the flight test derived position correction is only one knot different from the Flight Manual position correction, well within the scatter of the data. The negative value also correlates with historical experience by the CCFT of seeing "higher than expected headwinds" in practice and competition. Failure to correct for the Pitot-static position error would result in the aircraft flying slower than was planned for. Additionally, most cross country flights by the 94 FTS are flown at full throttle or 75 percent power, whichever is lower. In this range of airspeeds, the flight test and Flight Manual position corrections are within a knot of each other. However, at low airspeeds, the flight test position correction is about 6 knots higher than the Flight manual position correction. This will result in a conservative error, with the aircraft on the proper approach indicated airspeed actually flying at a higher calibrated airspeed than predicted by the flight manual. Since operations at the 94 FTS have been successful over the years, there is no reason to change the Flight Manual takeoff or approach speeds. Figure A28 shows the altitude position correction curve at sea level from flight test and the Flight Manual. These curves were derived from the airspeed position correction curve, using the equation from Reference 3: AHpc = -AP. Pg 2ag rrffV While this correction varies slightly with altitude, the small values of ±30 feet are not operationally significant, and can be ignored for normal operations. All preceding Pitot-static data were collected on N557TH. Additional investigation was conducted to determine if noticeable differences existed in the Pitot- static errors between the three CCFT aircraft. One test point was flown using the GPS Ground Speed method in N557AW, but was rejected for excessive wind error. However, all drag polar and cruise data collected in N557AW (Figure A4 and Figure A8) matched the data of N557TH within the experimental scatter using the same position correction curve. Therefore, the flight test derived position correction curve was considered valid for both N557TH and N557AW. The flight test derived position correction curve did not work as well for N557SH. Figure A29 shows the curve along with Pitot-static data collected in N557SH. These data were collected using the GPS Ground Speed method. While the flight test data do not match the curve, the data do have the same basic shape. For N557SH, an acceptable position correction could be found by subtracting 4 knots from the flight test derived position correction curve, as shown in Figure A29. This finding correlates with operational experience that N557SH flying side by side with either of the two other aircraft would show a higher indicated airspeed. Investigate the Pitot-static position corrections for N557SH and N557AW. (R3) The test team found the GPS Ground Speed method to be superior to the GPS Speed Course method in both test efficiency and data quality. Test points could be accomplished much faster using the GPS Ground Speed method, and did not require maintaining a stable airspeed as long as in the GPS Speed Course method. Since the legs were shorter, it was easier to avoid local air disturbances such as thermals or upslope winds. Since the method includes a technique for approximating a crosswind heading, wind effects from incorrect winds aloft forecasts are minimized Additionally, the data can be evaluated by inspection for wind effects such as not being on a crosswind heading or wind gradients. GPS Speed Course data required calculations to determine effects of not being on crosswind heading, and did not indicate wind gradients in any way. CAS-* •IAS2) CLIMB PERFORMANCE Test Objectives: The test objectives for climb performance were: 1. Determine maximum rate of climb at full throttle. 2. Determine the airspeed for maximum rate of climb at full throttle. 3. Determine rate of climb as a function of airspeed at full throttle. 4. Determine best angle of climb at full throttle. 5. Determine the airspeed for best angle of climb at full throttle. 6. Determine time to climb, distance to climb, and fuel to climb as a function of altitude. 7. Create charts for Flight Manual climb data. Test Procedures; Climb data were collected using the sawtooth climb FTT (Reference 3). Full throttle constant airspeed climbs were conducted at SO, 60, 65, 70, 80, and 90 KIAS. For each test point, two climbs were flown on opposite headings perpendicular to the wind. Times were recorded every 100 feet of pressure altitude using the time function of the Hewlett Packard 48SX calculator. Climb data were reduced as shown in Appendix D. Test Results; Figure A30 and Figure A31 show the standard day rate of climb performance in terms of indicated and calibrated airspeed. Each figure shows lines indicating the best rate of climb airspeed and best angle of climb airspeed as they vary with altitude. While the values in indicated airspeed are more useful operationally, the values in calibrated airspeed are shown to justify the best angle of climb airspeed. The best angle of climb airspeed can be found at the tangent line from the origin to the rate of climb curve. This determination can be done on a rate of climb chart plotted against calibrated, equivalent, or true airspeed, since in each case the entire line for a given altitude is multiplied by the same factor regardless of airspeed. However, this determination cannot be performed on a rate of climb chart plotted against indicated airspeed. The shape of the curve changes since the conversion from calibrated to indicated airspeed is non-linear and dependent upon airspeed. The best angle of climb airspeeds were found using the chart plotted against calibrated airspeed, and these calibrated airspeeds were converted to indicated airspeeds and plotted on Figure A30. The flight test derived climb speeds compare to the Flight Manual climb speeds as shown in Table 2. In converting between indicated and calibrated airspeed, the flight test derived Pitot-static position correction was used for flight test data, and the Flight Manual correction was used for Flight Manual data. The resulting rates and angles of climb are shown in Table 3. Table 2 CLIMB SPEED COMPARISON (IN KIAS (KCAS)) Best Angle Best Rate Altitude Flight Test Flight Manual Flight Test Flight Manual Sea Level 54 (60) 56 (56) 71 (72) 68 (66) 10,000 ft 50(58) 56 (56) 59 (64) 62 (61) Table 3 MAXIMUM PERFORMANCE CLIMB RESULTS (Standard Day, Standard Weight) Best Angle Best Rate Altitude Airspeed (KIAS) Angle ('leg) Airspeed (KIAS) Rate (fl/min) Sea Level 56 7.5 65 865 10,000 ft 56 2.6 65 330 Comparing climb speeds in calibrated airspeed, the flight test results for best angle of climb are 2 to 4 knots faster than recommended in the Flight Manual. Flight test best rate airspeeds are 3 to 6 knots faster than recommended in the Flight Manual. Flying at the Flight Manual recommended speed for best angle will result in a climb angle of 7.2 degrees at sea level for a 0.3 degree (4 percent) loss of climb angle. The Flight Manual recommended speed for best rate will result in a rate of climb of 850 ft/min at sea level for a 15 ft/min 10 (less than 2 percent) loss of climb rate. These differences are small enough that no changes in the Flight Manual are warranted. Climb data is also presented for a cruise climb at 80 KIAS, which increases distance flown at a small loss of climb rate (2 percent at sea level, increasing to 26 percent at 10,000 feet) for situations where the mavitniim rate of climb is not required. Additionally, climbing at 80 KIAS improves the pilot's forward visibility by lowering the pitch angle. The climb data were analyzed using the RPM model. Since the RPM model will simulate non- standard atmospheric conditions, it was assumed that if the model could be made to match the flight test data at several non-standard conditions, then the model would be considered good and valid for any atmospheric conditions. Figure A32 shows climb data for two flight conditions, one at 8,000 feet and one at 12,000 feet pressure altitude and temperatures close to standard day temperatures. The RPM model was adjusted to closely match the 8,000 foot data, and then compared to the 12,000 foot data. The fairings in Figure A32 represent the RPM model prediction The RPM model data were considered to be in reasonable agreement with the 12,000 foot data. Figure A33 shows climb data for two flight conditions, one at 6,000 feet and one at 8,000 feet pressure altitudes and temperatures significantly above standard day temperatures. Climb data from the same RPM model is shown to be in reasonably good agreement with the flight test data. The maximum deviation from the flight test data is SO ft/min, which is only 1/2 a division on a Vertical Velocity Indicator (WI). The RPM model produced a valid representation of the aircraft climb performance. To get the RPM model data to match the climb data, two additional compensations were made within the computer program. The first was to account for expanding pressure contours on non-standard days. On a hotter than standard day, 1000 feet of pressure altitude is greater than 1000 feet of tapeline altitude. Therefore, a rate of climb expressed in terms of pressure altitude will be less than the same rate of climb expressed in terms of tapeline altitude. This compensation was merely an application of a principle normally used in climb data reduction. The second compensation was to account for an apparent increase in aircraft drag in climbs over that seen in cruise flight. This difference in drag was more noticeable at low speeds and less noticeable at high speeds. This result was hypothesized to be a result of the interaction of the slipstream and the separation drag from the cockpit rear window. The steeply sloping rear window is known to cause separated flow and thus increase the aircraft drag. Additionally, this window is fully engulfed in the propeller slipstream. At low speeds, the difference between the induced velocity of the propeller at full power and cruise power is the greatest, reducing to no difference at maximum airspeed. Therefore, the slipstream velocity over the rear window would be much higher in a slow speed climb than in cruise flight at the same airspeed A relationship was developed and applied to the model data to account for this extra drag. This relationship and the method for accounting for non-standard day pressure altitude variations are further described in Appendix C. Because the sawtooth climbs were relatively short compared to the amount of fuel burned, fuel used during the climb was not recorded. The fuel flows were calculated by the RPM model using the same fuel flow calculation method from cruise flight as a function of engine MAP and RPM. Figure A34 through Figure A39 show rate of climb, fuel flow, time to climb, fuel to climb, and distance to climb at 65 KIAS and 80 KIAS. These charts are also submitted for Flight Manual inputs in Appendix B. These charts represent the RPM model and will give values for non-standard conditions. To determine rate of climb, enter the bottom left side of the chart at the appropriate OAT, go up to the pressure altitude, across to the rate of climb line, and straight down to read the rate of climb. The variation of fuel flow with non-standard temperature and pressure are almost identical to the rate of climb variation, so both of these values are plotted on the same chart. For the example shown in Figure A34: OAT: 80° F Pressure Altitude: 6,000 feet Rate of Climb: 450 ft/min Fuel flow is found using the same procedure with the fuel flow line. 11 To determine time to climb, fuel to climb, or distance to climb, the chart must be used twice. Enter with the initial OAT, go up to the initial pressure altitude, over to the appropriate line, and straight down to read the value. Repeat this process with the final conditions. The difference between the two values will be the time, fuel, or distance expected to be seen in the climb. For the example shown in Figure A3 5: Start OAT: Start Pressure Altitude: Start Time: End OAT: End Pressure Altitude: End Time: Time to Climb: 80° F 6,000 feet 16 min 66° F 10,000 feet 27 min 11 min Fuel to climb and distance to climb are found using the same procedure with the appropriate line. There is a substantial difference between the variation of time, fuel, and distance to climb and the variation of rate of climb with non-standard conditions. Thus, these are plotted on separate charts. However, the difference in the variation of time and fuel to climb and the variation of distance to climb with non- standard conditions is small; on the order of 5 percent To reduce the number of charts in the pilot's checklist, the time, fuel, and distance to climb are presented on the same chart in the Flight Manual inputs in Appendix B. This is consistent with the data presentation format used by some general avaiation manufacturers. The climb data presented were based on results from N557TH. The CCFT suspects that differences may exist between the climb performance of the three aircraft Further testing should determine if differences exist in the climb performance of the three CCFT aircraft. (R4) DESCENT PERFORMANCE Test Objectives: The test objectives for descent performance were: 1. Determine the best no wind glide ratio with throttle idle. 2. Determine airspeed for best glide ratio with throttle idle. 3. Determine the minimum sink rate with throttle idle. 4. Determine airspeed for minimum sink rate with throttle idle. 5. Determine time to descend, distance to descend, and fuel to descend at the best glide ratio airspeed as a function of altitude. 6. Determine time to descend, distance to descend, and fuel to descend at maximum structural cruising speed (VN0 - top of green arc on airspeed indicator) as a function of altitude. 7. Create charts for Right Manual descent data. Test Procedures: Descent data were collected using the sawtooth descent FTT (Reference 3). Idle power constant airspeed descents were conducted at 50, 60, 65, 70, 80, and 90 KIAS. Descents were also flown at 107 KIAS and 2250 RPM to simulate enroute descents. For each test point, two descents were flown on opposite headings perpendicular to the wind. Times were recorded every 100 feet of pressure altitude using the time function of the Hewlett Packard 48SX calculator. Descent data were reduced using the methods described in Appendix D. Test Results: Descents were analyzed by considering the aircraft as a glider, i.e. counting any windmilling drag from the propeller against the airframe, and finding a drag polar which would represent the descent performance. This drag polar was determined by fitting a straight line to values of the drag coefficient plotted against the square of the lift coefficient, as was done for cruise data. This curve fit is shown in Figure A40. The resulting drag polar is shown in Figure A41. For reference, these figures also show the cruise drag polar. The idle descent drag polar is unusual in that it is less than the cruise flight drag polar. Generally a windmilling drag polar is greater than the cruise drag polar due to the additional drag from the windmilling propeller. However, in this case it was suspected that the reduction in separation drag over the cockpit rear 12 window from the reduced slipstream velocity was larger than any increase in drag arising from the windmilling propeller. The fact that both drag polars have the same parasite drag coefficient was suspected to be strictly coincidental. Aircraft without a rear window like the Cessna 150, and thus without the separation drag, would see a different relationship between the cruise and idle descent drag polars. The idle descent drag polar was CD = 0.0427 + 0.0477CL2 Using this drag polar, the descent performance for the aircraft was analyzed. Figure A42 shows the penetration chart (L/D vs. Indicated Airspeed). The maximum glide ratio was 11 at 50 KIAS. At the Flight Manual recommended glide speed of 65 KIAS, the glide ratio was 10.5, or a reduction of 5 percent. Either airspeed should be operationally acceptable. The Flight Manual speed has the advantage of being the same as the climb speed, and thus one less airspeed for the pilot to remember. Figure A43 shows the polar chart (Rate of Descent vs Indicated Airspeed). Figure A44 shows the same data presented against true airspeed. These charts show a small variation in rate of descent with altitude. The minimum sink rate at sea level is 530 ft/min at 50 KIAS. Theory states that the minimum sink rate should occur at a slower airspeed than the best glide ratio. The true minimum sink rate probably occurs at a slower speed than 50 KIAS, and possibly the minimum sink rate is at just above the stall speed, and not at the minimum of the power required curve. For the airspeeds tested, the minimum sink rate occured at 50 KIAS. Figure A40 through Figure A44 also show values for a penetration descent at VN0 (107 KIAS) and 2250 RPM The tachometer was placarded to avoid descending in the range of 1850 - 2250 RPM. Idle RPM would be below this range, and the descent rate would be too high for a normal penetration descent. Flying at full throttle and 107 KIAS would overspeed the engine at high altitudes and have too slow of a descent rate. An RPM of 2250 was chosen as being easy to remember, and the top end of the caution range. A CL of 0.28 and a CD of 0.022 were used to predict descent performance for this flight condition, as shown by the two labeled data points on Figure A41. The resulting descent rate of 900 ft/min is probably still too high for a penetration descent in operational conditions. Investigate descents at 107 KIAS and RPM greater than 2250 to find the optimum throttle setting for a penetration descent. (R5) Fuel burn was not measured during the descents because of the short duration of the descents. By observing the fluctuating fuel flow indications, the test team estimated a fuel flow of 1.5 gal/hr for idle descents, and a fuel flow of 5.5 gal/hr for descents at 107 KIAS and 2250 RPM. Figure A45 through Figure A47 show the descent performance at idle power and 65 KIAS for non- standard conditions. Figure A48 through Figure A50 show the same data for descents at 107 KIAS and 2250 RPM These charts are also submitted for Flight Manual inputs in Appendix B. These charts are used in the same manner as the corresponding climb charts. In this case, the variation of distance with non- standard conditions was sufficiently different from that of time and fuel that distance is presented as a separate chart. TAKEOFF PERFORMANCE Test Objectives: The test objectives for takeoff performance were: 1. Determine takeoff ground roll using the Flight Manual takeoff procedure. 2. Create charts for Flight Manual takeoff data. Test Procedures; Takeoff data were collected using the Flight Manual procedure. This procedure consisted of Maintain directional control by use of nosewheel steering. Hold the elevator slightly aft of neutral to keep weight off the nose gear and hold aileron into the wind. At 50 KIAS, raise the nose smoothly to takeoff attitude. Maintain this attitude and allow the aircraft to fly off the ground which will occur between SO and 60 knots. (Reference 1) 13 All takeoffs were done with flaps fully retracted. The fuel mixture was leaned at fields above 5000 feet elevation. Below 5000 feet elevation, takeoffs were done with the mixture at full rich. Pressure altitude, outside air temperature, fuel used, wind direction and wind velocity were recorded prior to takeoff. The time from brake release to liftoff and the liftoff airspeed were recorded during the takeoff. If available, runway lights were used to estimate the takeoff distance. The data were reduced and corrected to a common liftoff speed to determine takeoff distance. These distances were standardized to produce a chart for predicting takeoff distance by the methods shown in Appendix D. Test Results: chart is of the same form used by several general aviation manufacturers. To use this chart, enter at the field OAT. Go up to the current field pressure altitude. Go across to the Weight Reference Line. From here, follow the guidelines down until reaching the vertical line for the takeoff weight Go across to the Wind Reference Line. Follow the guidelines to the wind component down the runway (down for headwinds, up for tailwinds). Go across to the right side to read the mean takeoff ground roll in feet. For the example shown: OAT: 80° F Pressure Altitude: 6,500 feet Weight: 1600 lbs Headwind: 10 knots Ground Roll: 1170 feet Twenty four takeoffs were accomplished at pressure altitudes ranging from 1490 to 6780 feet. The Flight Manual procedure specifies a rotation airspeed, not a liftoff airspeed. The liftoff airspeeds varied from 52 to 65 KIAS, with an average of 57 KIAS. All takeoff data were standardized to a liftoff airspeed of 57 KIAS, zero wind, standard weight of 1760 pounds, and sea level density. These results are shown in Figure A51. This chart was also included in the Flight Manual inputs in Appendix B. The mean ground roll distance was 1000 feet at a mean liftoff airspeed of 57 KIAS. The 95 percent confidence interval for ground roll distance (one-tailed test; shorter ground rolls are not an operational concern) was bounded at 1165 feet The 99 percent confidence interval for ground roll distance was bounded at 1234 feet The 95 percent confidence interval (two-tailed test) for liftoff airspeed was bounded at 51 KIAS and 63 KIAS. Figure A53 and Figure A54 are included to show the effects of dispersion on ground roll distance. These figures show how much additional distance should be added to find the distance at the limit of the 95 percent and 99 percent confidence intervals. Only the effects of weight and density altitude are shown. A headwind will always shorten the takeoff roll, and takeoffs should not be attempted in anything above a very small tailwind. To use these charts, enter with the takeoff weight go up to the appropriate density altitude, and to the left to read the dispersion distance. Add this distance to the mean takeoff ground roll to get the maximum expected ground roll. Note that normally Figure A53 and Figure A54 would not be needed by the operational pilot Additional runway length allowed for stopping after an engine failure on the runway will normally greatly exceed the additional distance from dispersion. Using the methods shown in Appendix D, the mean ground roll distance was expanded for non- standard conditions as shown in Figure A52. This 14 CONCLUSIONS AND RECOMMENDATIONS Performance data were collected on the USAF Academy Cadet Competition Flying Team (CCFT) Cessna 150 in the areas of cruise, Pitot-statics, climb, descent, and takeoff. These data were used to develop a computer model of the aircraft using the Reciprocating Engine and Propeller Modeling Program (RPM). This computer model was then used to create performance charts and tabulated data for the operational flight envelope for inclusion in the next update of the Flight Manual. All test objectives were met Cruise flight was characterized by the aircraft drag polar, CD = 0.042696 + 0.068861 CL2 derived from flight test Based on limited flights in two of the aircraft, drag and power required data for all three aircraft were in satisfactory agreement. 1. Additional testing should be conducted to verify the validity of the drag polar for all three aircraft. (Page 4)' The fuel flow data was modeled with satisfactory agreement with the flight test data. Tabulated cruise data were created on 5x8" cards for flight planning purposes, with power settings above 75 percent flagged. The TTiayimiim range airspeed was 65 KIAS. An airspeed of 85 KIAS gave the maximum airspeed per pound of fuel burned, with a reduction in range of 10 to 20 percent Typical 94 FTS operational procedures of flying at roa*i"""" speed resulted in a reduction in range of 23 to 40 percent The maximum endurance for airspeeds tested occurred at 50 KIAS, the minimum airspeed tested. 2. Flight data, such as RPM, fuel flow, fuel used, indicated airspeed, true airspeed, range, and endurance should be collected on 94th FTS deployments and CCFT practice and competition flights and compared to the 2 Page numbers in parentheses refer to the page number in the Test and Evaluation section of this report where the recommendation is made. performance data presented in this reportfor further verification of these performance data. (Page 7) The Pitot-static position correction curve was derived from flight test using the GPS Speed Course and GPS Ground Speed methods. The flight test derived curve followed the general trend of the Flight Manual curve. At airspeeds normally seen during competition or cross country flight the flight test derived curve was within one knot of the Flight Manual curve. At low speeds, the flight test curve is about 6 knots higher than the Flight Manual curve, resulting in higher calibrated airspeeds at Flight Manual takeoff and approach speeds. Based on good operational experience, there is no reason to change the Flight Manual takeoff and approach speeds. The flight test derived position correction curve was considered valid for N557TH and N557AW. An acceptable position correction curve for N557SH could be found by subtracting 4 knots from the flight test derived position correction curve. 3. Investigate the Pitot-static position corrections for N557SH and NS57AW. (Page 9) Climb data were analyzed using the RPM model to match flight test data at several non-standard conditions. The model was then considered valid at all flight conditions. The flight test results for best angle of climb airspeed and best rate of climb airspeed were in good agreement with the Flight Manual, and do not warrant any changes to the Flight Manual airspeeds. The maximum rate of climb was 865 ft/min at sea level and 330 ft/min at 10,000 feet. The maximum angle of climb was 7.5 degrees at sea level and 2.6 degrees at 10,000 feet. Climb charts are presented for maximum rate of climb at 65 KIAS and for cruise climb at 80 KIAS. The climb data presented were based on results from N557TH. The CCFT suspects that differences may exist between the climb performance of the three aircraft. 4. Further testing should determine if differences exist in the climb performance of the three CCFT aircraft. (Page 12) 15 Descent data were analyzed by considering the aircraft as a glider, counting the windmilling drag of the propeller against the airframe. The idle descent drag polar was CD = 0.0427 + 0.0477 CL2 which led to the curious conclusion that the idle descent drag was less than the cruise drag. This was suspected to result from the interaction between the slipstream and separation drag from the cockpit rear window. The maximum glide ratio was 11 at SO KIAS, and 10.S at the Flight Manual recommended glide speed of 63 KIAS. This small difference does not warrant a change to the Flight Manual, since 65 KIAS is easier to remember as the same airspeed for best rate of climb. The minimum sink rate for the airspeeds tested was 530 ft/min at 50 KIAS. Descent performance was also investigated for a penetration descent at 107 KIAS (VN0) and 2250 RPM. At these conditions, the lift coefficient was 0.28 and the drag coefficient was 0.022. The resulting descent rate of 900 ft/min is probably still too high for a penetration descent in operational conditions. 5. Investigate descents at 107 KIAS and RPM greater than 2250 to find the optimum throttle setting for a penetration descent (Page 13) Takeoff data were standardized to zero wind, standard weight of 1760 pounds, and sea level density. The mean ground roll distance was 978 feet at a mean liftoff airspeed of 57 KIAS. The 95 percent confidence interval for ground roll distance was bounded at 1140 feet The 99 percent confidence interval for ground roll distance was bounded at 1208 feet. The 95 percent confidence interval for liftoff airspeed was bounded at 51 KIAS and 63 KIAS. The performance of the CCFT Cessna 150 was satisfactorily characterized. Further testing should address the recommendations of this report, and the results of this testing should be incorporated in the Flight Manual. 16 REFERENCES 1. Operating Instruction 51-150, 94th FTS, USAF Academy, Colorado, 15 September 1993. 2. Erb, Russell E., Reciprocating Engine and Propeller Modeling Program, computer software, Erb Engineering, Arlington Texas, yet to be published. 3. Payne, James M, Flight Test Handbook, JP Aviation, USAF Academy, Colorado, 1989. 4 Bailey William D, Captain, USAF, et al, Investigation of Using Global Positioning System for Air Data ' System Calibration of General Aviation Aircraft (HAVE PACER 11), AFFTC-TR-95-76, AFFTC, Edwards AFB, California, January 1996. 5. Anderson, John D., Jr., Introduction to Flight, 3rd ed., McGraw-Hill Book Company, New York, 1989. 6. Carson, B. H., Fuel Efficiency of Small Aircraft, AIAA-80-1847, AIAA Aircraft Systems Meeting, Anaheim, CA, 4-6 August 1980. 7. von Mises, Richard, Theory ofFlight, Dover Publications, Inc., New York, 1959. 17 This page intentionally left blank. 18 APPENDIX A TEST DATA 19 U 3AFA CCFT CMIM 150/150HP Englm: Lyconing O-H0-E2D Pnptttr. HcCautoy TH745M1C172 HtaKun: Lamad CarbHaat OFF 140 c © o E o o 58 0.13 0.12 0.11 0.10 0.09 0.08 0.07 0.06 0.05 0.04 0.03 0.02 0.01 0.00 0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 Test Brake Horsepower From MAP/RPM Figure Al Engine Horsepower Determination Methods Comparison USAFA CCFT Ctnni 150/150HP Engine Lycoming O-320-E2O Propallar: McCaultyTM745»1C172 Mixtur»: Luntd Weight 17601b« CarbHaat OFF Flaps: UP Data Bute: Flight Tast CD = 0.042696 + 0.068861 CL2 + ■ 111 11 i i i i i 11 11111 i i i i i i 111 111 i i i 0.0 0.2 0.4 0.6 0.8 Lift Coefficient Squared, CL2 1.0 Figure A2 Drag Polar Curve Fit 20 0.14 0.12 0.04" 0.02 0.00 USAFA CCFT Cassna 130/150HP NS57TH Engine Lycomlng O-320-E2D Propallan McCauleyTM7458/1C172 Mixtur«: Laanad Weight 1780 lb* Car* Heat OFF Flip«: UP Data Bade Flight Tast CD = 0.042696 + 0.068861 CLZ ■ ■■'■' -H- ■ i ■■■••«* * ' I ■ ' ■ ' ' —H 0.0 0.2 0.4 0.6 0.8 Lift Coefficient, CL Figure A3 N557TH Drag Results Compared to Aircraft Drag Polar USAFA CCFT Cassna 180/150HP NSS7AW Engine Lycomlng O-320-E2D Propallac McCaulty TM74SW1C172 Mixture: LaaiMd Waight 17601b* Cart) Haat OFF Flap* UP Data Basis: Flight Tast 1.0 c 0) o o O Lift Coefficient, CL Figure A4 N557AW Drag Results Compared to Aircraft Drag Polar 21 0.14 0.12 USAFACCFT Cessna 150/150HP NSS7SH Bigine: Lycomlng C-320-E2O Propeller McCauleyTM74S8/1C172 Mixtur«: Leaned Weight 176011» Can Heat OFF FUp«: UP Data Basis: Flight Test O 0.10 c | 0.08 1o 0.06 O) £ 0.04 CD = 0.042696 + 0.068861 CL2 0.02 0.00 ■ 0.0 I I I I I I I I I I I I I I I I I I I I I I I I 0.2 0.4 0.6 Lift Coefficient, CL 0.8 1.0 Figure A5 N557SH Drag Results Compared to Aircraft Drag Polar USAFA CCFT Cessna 150/150HP Engine: Lyeoming O-320-E2O Propeller McCauley TM7468V1C172 Mixture: Leaned Weight 1760 lbs Cart Heat OFF Flaps: UP Flight Test BHP^V^ = 7.950E-05 V^4 + 1714 J 1 I I 1 1 L_ _J I 1 L_ 40000000 80000000 120000000 160000000 Viw4 (knots4) Figure A6 Brake Horsepower Required Curve Fit 22 a * a. z m iV 12 o X E m .a ■E n ■o e to 130 120 110 100 90 80 70 60 50 40 30 20 10 0 USAFACCFT Cessna 13V150HP N997TH Engine Lycomlng O-320-E2D Propallan McCaul«yTM7458/1C172 Mixtur.: LMntd W.lght 1760 lbs CarbHsat OFF Flaps: UP Data Basis: Flight Tast , ■ I ■ ■ ■ ■ i i ■. ■ i i ■ ■ ■ I i i i i I i i i i I i i i i 40 50 60 70 80 90 100 110 120 Standardized Equivalent Airspeed, Viw(KEAS) Figure A7 N557TH Brake Horsepower Required Results Compared to Aircraft Brake Horsepower Required 2 a. ia eo E CD 1"E 8 to 130 120 110 100 90 80 70 60 50 40 30 20 10 0 USAFA CCFT Cassna «0/1S0HP NS97AW Englna: Lycomlng O-320-E2O Propallar McCaukyTM7458/1C172 Mixtur.: Laaned Walght 1760 lbs CarbHaat OFF FI*P«: UP Data Basis: Flight Tast _l I I 1 L I I | I I I I | I I I I | ' ' ' ' | I''' | 40 50 60 70 80 90 100 110 120 Standardized Equivalent Airspeed, Viw(KEAS) Figure A8 N557AW Brake Horsepower Required Results Compared to Aircraft Brake Horsepower Required 23 130 _120 afl10 CQ 100 USAFACCFTCannalMSISOHP NS67SH Engine lycomkig O-320-E2D ProptUan McCaulay TM74SB/1C172 Mixtura: Lsanad Wdght 17601b« CarbHut OFF Flaps: UP DM« Basic FllsMTttt J—I—1—i—i—i—I—i i I i I i i i i I i I i 1 I i i i i I i i i i I • • i i H" 40 50 60 70 80 90 100 110 120 Standardized Equivalent Airspeed, Vjw (KEAS) Figure A9 N557SH Brake Horsepower Required Results Compared to Aircraft Brake Horsepower Required 0.8 | 0.7 4 K 0.6 S I 0.5 i 0.4 4 sÜ 1 0.3 u. I °'2 CO | 0.1 ffl t: 8000ft 10000 ft' 16000 ft 0.0 USAFA CCFT Cessna 15W1SOHP Englm: Lycenlng 0-33&C20 PropaMar McCaulayTM7438MC172 Mbitura: Laanad Walght 1760 lbs Carb Haat OFF Flaps; UP Data Basis: Flight Test Fairing Basis: RPU Modal .Saa Laval T I IIIIIII|IIIIMIII|III MM II II II I II llll II II I II II |l II I II I II li i Ulli iii.i ii I in i I [ I I I I I I I I I I I 30 40 50 60 70 80 90 100 110 120 130 140 Brake Horsepower, BHP (hp) Figure A10 Brake Specific Fuel Consumption Results 24 USAFA CCFT Casana «0/160HP Engln«: Lycomlng O-320-E20 Mbcturt: LMiwd CarbHaat OFF DataBaala: Flight Taat Propallan McCaul«yTM74««/1C172 Wtlght 17601b« Flap«: UP Falling Basic RPM Modal 50 60 70 80 90 100 110 Standardized Equivalent Airspeed, Vh* (KEAS) 0.25 T (0 £o.20 + HI CO ©0.15 + o c 10.10 tu o §0.05 + Q. CO 0.00 Figure All Specific Air Range Results USAFA CCFT Cessna 180/1S0HP Engine Lycomlng O-320-E20 Prep«ltar McCaulayTM7458/1C172 Mixture: Uanad Wslght 1760 lbs Can Hast OFF naps: UP Data Basis: night Tast Fairing Baals: RPU Modal Saa Laval 18,000 ft 10.000 ft 8.000 ft _i i i—i—i—i_ _j i i—i——i—i—i—i- 50 60 70 80 90 100 110 120 Standardized Equivalent Airspeed, Viw (KEAS) Figure A12 Specific Endurance Results 25 USAFA CCFT Cassna ISOflSOHP Engliw: Lyeomlng O-320-E20 Mixtur«: Luntd Carli Hut OFF Data Basis: fUH» Modal Propallar: McCaulty TM7468/1C172 W.lght 17Mlb« Flap«: UP 40% Parcant Braks Horsapowar 0 20 40 60 80 100 60 70 80 90 100 110 120 Outside Air Temperature, OAT (°F) True Airspeed, V (KTAS) Figure A13 Cruise Airspeed Performance USAFA CCFT Cassna 150/150HP Engine Lyeomlng O-320-E2O PropaUar McCaulayTM745V1C172 Mixture: Laanad Walght 17601b* CarbHaat OFF Flap«: UP Data Basis: /PMModal 40% Parcant Braks Horsapowar 0 20 40 60 80 100 1700 1900 2100 2300 2500 2700 Outside Air Temperature, OAT (°F) RPM Figure A14 Cruise RPM Performance 26 USAFA CCFT Caaana 16O/160HP Engina: Lyeomlng O-S20-E20 Propallar: MeCaul«yTM7«S»71C172 Mixtun: Laanad Walgbt: 1760 lb« CartoH.lt OFF 40% &**'■ up Data Baal«: SHU Modal Pareant Brak« Horaapowar | r ■ ■■ [ | ....|.... |. ■■ ■ | ■ ■ii|ii i i |i i i i | i ITI | i i i i |i i i i | i i i i | i i i i | i i n | 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 Fuel Flow (gal/hr) Figure A15 Cruise Fuel Flow Performance USAFA CCFT Caaana 160/150HP Emma: Lycomhg O-M0-E2D Propallan McCaulay TM746W1C172 Rial at Startup: 2« gal MMura: Laanad Walghfc 1760 fc« Start, Taxi, Takaoff, Climb Fuat: 2 gal CarbHaat OFF Rap« UP Climb Olatanea: 10 nm N 7llt 0 50 100 11 £+80 r Outside Air Temperature, s e'S n OATCF) |=| ü IIa -80 200 250 300 350 Dual Range (nm) Figure A16 Dual Constant Airspeed Cruise Range Performance 27 USAFA CCFT Caaan* 150/150HP Englna: Lycomlng O-320-E2D Propattar McCauWy TM7458/1C172 Fual at Startup: 26 gal Mbctura: Uanad Walghfc 1760 lb« Start, Taxi, Takaoff, Climb Fuat 2 gal CarfaHaat OFF Flap« UP Clanb OiaUnca: 10 ran DataBaata: M>*f Modal ■0. 75. 65, Knot* Indlcatad Alrpapaad 100 0 50 Outside Air Temperature, OAT CF) 150 200 250 300 Dual Range with 45 min reserve (nm) Figure A17 Dual Constant Airspeed Cruise Range Performance With 45 Minute Reserve USAFA CCFT Caaana 1S0/150HP Englna: Lycomlng O-330-C2D Prop.lan McCauUy TM7458/1C172 Fual at Startup: 26 gal Mixtur* Laanad CaitHaafc OFF Data Skala: «P* Modal Walght: 17601 Flapa: UP 14000 -r £ 12000 4 f 10000 I 8000 £ 6000 f 3 (0 (0 I 2000 -E 0 JJE+80-T til o-; *ll -80 -t SUrt, TaxL Takaoff, Climb Fual: 2 gal Climb Tima: tmlnuta* 75 70 68 60 55 g 4000 Knota Indlcatad Alrpapaad 100 ** I I I I I I I + + + I I I I I I I I I I I I I I I I I I I I I I I I I H 0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 Dual Endurance (hr) Figure A18 Dual Constant Airspeed Cruise Endurance Performance 28 Engin«: Lycom log O-320-E20 Mixtur*: Laanad CarbHaat OFF Data Baal«: RMI Modal USAFA CCFTCaaana 150/150HP PropaUar: McCaulay TM745S71C172 W.lght: 17M UM Flapa: UP Knot« Indicated Alrpapaad Fual at Startup: 26 gal Start, Taxi, Takaoff, Climb Fual: 2 gal Climb Tim« t minute« 70 65 60 65 50 I I I I I I I I I I I I + I 1 I I I I I I I I 0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 Dual Endurance with 45 min reserve (hr) Figure A19 Dual Constant Airspeed Cruise Endurance Performance With 45 Minute Reserve USAFA CCFT Caaana 1S0S150HP Engln«: tycoxlng O420-E2D Propalten McCaulay TM745S/1C172 Fual at Startup: 36 gal Manure: Laanad Walght: 1760 fca Start Taxi, TakaofT, Cllnib Fual: 2 gal CarbHaat OFF Flap«: UP Carnb Olatenc«: 10 nm DateBaaiK MPafModai •0 75 65 Knot« Indicated ■* Airpapaad 9« 90 50 Outside Air Temperature, OAT (*F) 100 +80T lit o 1 §"2 &i 5 -80 HS I I I I ■ I ■ I ■ t I I I ■ I I ■ I I I I I I I I I 300 350 400 450 500 550 600 Solo Range (nm) Figure A20 Solo Constant Airspeed Cruise Range Performance 29 USAFA CCFT dura 150/1SOHP Englna: Lycomlng O-320-E2D PropaBar McCaulay TM745S/1C172 FIMI at Startup: M gal Mbtturo: Laanad Waighfc 17M Iba Start, Taxi, Takaoff, Climb Fuat 2 gal Cart) Haat OFF Flapa: UP Climb Diatanca: 10 mi Data Bast« «P* Modal •0 76 6» Knots Indlcatad 85 Alrpapaad 50 Outside Air Temperature, OAT (°F) 250 300 350 400 450 500 Solo Range with 45 min reserve (nm) Figure A21 Solo Constant Airspeed Cruise Range Performance With 45 Minute Reserve Englna: Lycomlng O-320-E2D Mixture: Laanad CarbHaat OFF Date Baakt: RPM Modal USAFA CCFT Caaana 1S0/150HP Prooalan McCaulay TM74SS/1C172 Fual at Startup: 3« gal Wakjht 17S0t» Start, TaxL Takaoff, Climb Fual: 2 gal Flapa: UP Climb Tim« Smlnutaa SO 7S 70 IS SO SS SO 0 12 3 4 5 6 7 Solo Endurance (hr) Figure A22 Solo Constant Airspeed Cruise Endurance Performance 8 30 Engin«: Lycom log O-J20-E2D Mbrtura: Uamd CaitoHaat OFF Data Buh: RPIf Modal USAFA CCFT Cnm 15W150HP Propall«: McCauUy TH745W1C172 W.lght 17S0U» Flap«: UP Knot« liuMcatad AkpapMd Fud at Startup: 36 gal Start, Taxi, Takaoff, Climb Fual: 2 gal CUmbTima: Smlnutaa 70 65 60 SS 50 | I I I I | I I I I | I I I I | ' I ' ' | I I I I I I I I 01 2345678 Solo Endurance with 45 min reserve (hr) Figure A23 Solo Constant Airspeed Cruise Endurance Performance With 45 Minute Reserve USAFA CCFT Caasna 1507180HP Engine Lycomlng O-320-E2D Propallar McCaul«yTM745671C172 Mixture: Laanad Walght 17601b» CartHaat OFF Flap«: UP Data Basic RPM Modal 1ÖU - 140 |l20- |100 o 8" 80- 12 x 60 - o 2 40 J m 20- . 0 - !_.—i_.—| ■ ■ ■ + + 0 20 40 60 80 100 True Airspeed, V (KTAS) Figure A24 Aircraft Brake Horsepower Required and Available 120 31 140 USAFA CCFT Csssna 160/160HP Englm: Lycomlng O-320-E2D Propslkr: McCaulsy TM74SW1C172 Mixture Uansd Wslght: 1760 lb« CvbHut OFF Flap«: UP Data Basic RPMModsl 20 40 60 80 True Airspeed, V (KTAS) 100 120 Figure A25 Aircraft Thrust Horsepower Required and Available 600 USAFA CCFT Cessna 160/1S0HP Englns: Lyeomlng O-320-E2D Propsllsr: McCaulsy TM746I/1C172 Mbttura: Lsaiwd Wslght 1760 lbs CarbHsat OFF Flaps: UP DaU Basis: ftPMMooM (0 £500 20 40 60 80 True Airspeed, V (KTAS) 100 120 Figure A26 Aircraft Thrust Required and Available 32 USAFA CCFT Caaana 160/150HP Engine Lycomlng C-S20-G2D Proper McCaulayTM74«8/1C172 Weight 17MltM Flapa: UP DataBase FHgMTM« 10 3f 8 0 e 6 Jt 4> < 2e 0 '■6 0 e -2 o o -4 e o -6 M o -8Q. 1a. 12 -10 -12 < -14 o ■'S eo o o HS «A O Q. « < 40 x 30 I 20 10 0 -10 t -20 : -30 | -40 a Flights A Right 6 « Right 7 o Right 10 ■ Rt 22/23 sFlight Manual I ■ ... | ■■■■ | ■ ■ ■ ■ | i i i i | i i i i | i i i 40 50 60 70 80 90 100 110 Indicated Airspeed, V, (KIAS) Figure A27 Airspeed Pitot-Static Position Correction USAFA CCFT Caaana 160/1SOHP Englna: Lycomlng O-J20-E2D PropaBar: MeCaulayTM7458/1C172 Walght 17NltM Flap«: UP AkKuda: Saa Laval Data Baal«: Flight Ta»t Flight Test SFlight Manual i ■ ■ ■ | ■ ' ' ' | i '—i i | i i ' ' | ' ' ■ ' | ■ ' ■ ' | ' 40 50 60 70 80 90 100 Indicated Airspeed, V, (KIAS) Figure A28 Altitude Pitot-Static Position Correction 110 33 Ie < o o c I o a. a. Ü2 < E g. O 2 xf E o a "S3 of U3AFACCFTC«una150/1S0HP N6673H Englna: Lycomlng O-320-E2O Propallar McCaultyTM7468/1C172 Wtlsht 17*0 lb* Flaps: UP Data Bui« Plight Tast 10 T 8 6 t 4 2 f 0 -2 f -4 -6 -8 -10 f -12 -14 N557TH, NS57AW a «18 * FK20 N667SH -I I I I I—I I u + + 40 900 -r 800 4 700? 600-1 500 4 400 4 300 4 200 -\ 100 4 50 60 70 80 90 100 110 Indicated Airspeed, Vi (KIAS) Figure A29 Airspeed Pitot-Static Position Correction, N557SH USAFA CCFT Cnana 1M/160HP Englm: Lyeomlng O-420-E2O PTopasai: McCsulayTM7468/1C172 Wxtun: Laamd Walght 17M KM CarbHaat OFF Flaps: UP Tamparatui«: Mandard Data Baal«: RPM Modal Best Rate of Climb Airspeed Best Angle of Climb Airspeed i i i i i 11 i i i i i i i i i i i i i i i i i i i i i i i | i i i i | i i i i | i i i i | i i i i | i i i i 0 10 20 30 40 50 60 70 80 90 100 110 120 Indicated Airspeed, V; (KIAS) Figure A30 Standard Day Rate of Climb Performance (Indicated Airspeed) 34 USAFA CCFT Ca«n 1SO/150HP c E £ ü g JO E ü 'S Ia: Englno: Lycomlng O-320-E2D Mbrtura: Lnrad CarbHsat OFF Tamparatura: Standard 900 T Best Rate of . Climb Airspeed. 800 - 700 Best Angle of /V Climb Airspeed ^//V 600 - 500 - 400- 300- r 200- ~ 100 0 |""|'" ' I " " I " " I " " I' ProH"^ McC«ul*y TM74SW1C172 Walght 1780 fea Flaps: UP Data Basis: RP*f Modal | ■ ■ i ■ | ■ ' ■ ■ | " ' ' | " ■ ' | " " | 10 20 30 40 50 60 70 80 90 100 110 120 Calibrated Airspeed, Vc (KCAS) Figure A31 Standard Day Rate of Climb Performance (Calibrated Airspeed) USAFA CCFT Csssna 1S0/180HP Englna: Lycomlng O-S20-62D Propslsr: ItoCaulty TM748W1C172 MUtura- Lssnsd Walght («,000 ft): 1708-18«« lbs CarbHsat OFF Wslght (12.000 ft): 1717-1M2 It» Flapa: UP Data Basis: Flight Tart Fairing Basis: RPM Modal c E g, o 2 | o 'S IS 500 450 f 400- \ 350 j 300 4 250 200 f 150-1 100 4 50 0 O 8,000 ft, 39* F A UOOttlL 34' F ! t , | .... | .... [ ■ ■ ■ ■ [ i i i i | I I I I | I I I I | I ' '' | I' '' | 45 50 55 60 65 70 75 80 85 90 Indicated Airspeed, V, (KIAS) Figure A32 Test Day Rate of Climb Matching 35 E g. ü O OHn n | O 'S 2 500 450 400 -E 350 300 | 250 4 200 150-t 100 50-E USAFA CCFT C*s«na 160/1MHP Engkw: Lycomlng O-320-E2D Propallar: McCauby TM74CV1C172 Mbrtura: Laanad Walght (6,000 ft): 173«-1707 lb« Cart) Hut: OFF Walght (1,000 ft): 1701 -1880 lb* Flap«: UP Data Baal«: Flight T««« Fairing Baala: RPM Modal o 6,0OOft69*F 0 8,000ft69*F + + 45 50 55 60 65 70 75 80 Indicated Airspeed, V, (KIAS) Figure A33 Test Day Rate of Climb Matching USAFA CCFT Caaana 160/160HP Engine Lycomlng O-320-E2D Propallar: McCaul*yTM74S8/1C172 Mbrtura: Laanad Walght 17*0 lb« Cart) Hast OFF Flapa: UP Alrapaad: M MAS Data Baris: RPM Modal 85 90 _»;; \ \ r^*>&# \ \ -^ \ f^ \ y ^HT \ \" -&p»Z^ ^ W <- -- X ^ as* ^ v4l ^- ¥ \ % tV -«* *> -"'i i *> v> \ K \ \ '\ ^ ~ii* \ > ^ --* \ \ \ \ \ ^ «1Pir \ > V , *- ^ \, - •i \ \ S* ^ A *i»}5» \ \ ^ ^ ^ V y \ ^- ^ \ ' \ \ 20 40 60 80 Outside Air Temperature (°F) 0 200 400 600 800 1000 1200 Rate of Climb (ft/mln) and Fuel Flow(gal/hr*100) Figure A34 Nonstandard Day Rate of Climb Performance at 65 KIAS 36 USAFA CCFT Ciunl 150/1 »HP Enoln«: Lycomlng O-S20-E2D Prepallai: MeCaul«yTM74««/1C172 Mbrtura: Laanad W.lght 1«0lb. CarbHaaC OFF Ftapt: UP DataBase: RP»f Modal lO^-- ~7^ ^ _^^u.,ggr^_-^- -*' ^- -J&L^^ 7 Z ^"~\\^\^'^ —"""" ./ _2 — *§^p2>—~ ^-—' -# ^/ -~~\v\ ■—""'' ^j;-— _»Z] / ^^^KgB^"'' "^,-- _J _/ -~'~~\ ■~~~'~~ -= x Z .--"^ Vpp""^ , - t_< . —"""" USS1^"""""" X ——^~v —— 2_2 , \ i-t«!— — ' / —" y~7 :: it ' jL 0 20 40 60 80 0 Outside Air Temperature (*F) 10 20 30 40 50 60 Time (min) and Fuel (galMO) 70 Figure A35 Nonstandard Day Time and Fuel to Climb at 65 KIAS USAFA CCFT Catena 160/1 «OHP Bigk»: Lycomlng O-120-E2D Propeller MeCauleyTM7«S«/1C172 Mlxtura: Leaned Weight 1700 fee Cart Heat OFF Flap«: UP Data Baal«: WWW Modal 7p Uj I—- gsi -2*«-" Ö»25« <<fh f> k • 3L>J t 4 «Sjkj- / t\ .. j ' jw. *S5A" / V J " VUjj i . \ 1 -* =** j. oyji. / — " / — •* AE2 I , / y / 10 20 30 40 50 60 70 Distance (nm) Figure A36 Nonstandard Day Distance to Climb at 65 KIAS 0 20 40 60 80 0 Outside Air Temperature (°F) 37 USAFA CCFTCHI» 150/150HP Engin«: Ly coming O-320-E20 Prop«ll«c McCaulay TM74M/1C173 Mixtur.: Laanad W.lght: 1780 lb« CarbHaat' OFF Flap«: UP Ainpaad: M MAS D«ta Ba.la: RPM Modal \ \ -—- "\ liS ""l ^JO« ^^ \ ) ^-, *> wPi \\ -- **■ Yi \ \11 -- 4 «•"•I- 5 \ L k 2\ k> A ■ >" \ \ -"* 1 \ £5l V V \ ^■i \ \ ?5^i \ v1 ) «1 S \ ^ ^, V --- ^ \ \ *"* '\ »• \ ^ r^ >* \ \ \ 0 20 40 60 80 0 200 400 600 800 1000 1200 Outside Air Temperature (°F) Rate of Climb (ftrniln) and Fuel Flow (gal/hr*100) Figure A37 Nonstandard Day Rate of Climb Performance at 80 KIAS USAFACCI-T C«««IM 150/150HP Engln«: Lycomlng O-320-E2O Prop.»«; McCaulay TM745V1C173 Mixtur«: U«n«d W.lght 1760 lb« CarbHaat OFF Flap«: UP Airepaad: MNM Data Baal«: RPH Modd ^ w»*! ■0. i*^ / <**' IssütJÄ 4t 50J J*' / wy -A\ •a &*- / / < ^ / / r -* \" «0» %~- / / \ / / ■*■> i> n DOOJ / / \ / / \ liäK >! f / / I / / 0 20 40 60 80 0 Outside Air Temperature (°F) 10 20 30 40 50 60 70 Time (min) and Fuel (gal*10) Figure A38 Nonstandard Day Time and Fuel to Climb at 80 KIAS 38 USAFA CCFT Cmu 150/150HP Engln.: Lyeomlng 0-S2M20 Prop«««: HcCutay TM745W1C172 Mixture Lwntd W.lght 17Mlb« Carb H—K OFF Fhp»: UP AirapMd: 80 MAS D«U O.a.: W>M Mod«! ^ r^^^ ^^\ _,^-""V&2&T ^ **>^\ B*»(P^^--^^^^- \'#ä^^ -^^ju-d^T" ^^-^ y. ^,^ \ yp^X^ --"■"^ ^ -""wCI—- — ' ^--""''~ -2. ^"^TU ~"^ 7 ^^^-Vr^S-^ ■=: -V *""" -"" ""^ A „a- "~" X *■"" ^.^ "7}&\—""" ' """ """^ju—■—"^===* IZ "~ —"—zXI^s^^i— Z-——■— ' u 0 20 40 60 80 0 10 20 30 40 50 60 70 Outside Air Temperature ("F) Time (min), Distance (nm), and Fuel (galMO) Figure A3 9 Nonstandard Day Distance to Climb at 80 KIAS USAFA CCFT Cmn 150/1S0HP Engln«: lycoming O-S20-BD PmptttT. HcCautay TH74S8/1C172 Mbclura: Laarnd Walghfc 17Mlb» Carb HMC OFF Flap« UP ThrotoV IOLB Data Bute: FllgMTaa« 0.12 T Idle Descent Drag Polar CD-0.0427 +0.0477 CL2 0.00 I I I I I I I I I I I I I I I I I I I I I ' I I I ''■■ I ll11 I ll11 I 0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 Lift Coefficient Squared, CL2 Figure A40 Idle Descent Drag Polar Curve Fit 39 USAFA CCFT ClMU 1SW150HP Enflln.: Lyeomlng O-320-E2D Propallar McCaulay TM7458/1C172 Mbtura: Laanad Walght: UM lbs CartiHaat: OFF Flap«: UP Thrott«: IDLE Data Basis: Flight Tast 0.12 T 0.10 a O «0.08 2 a Ho.06- 0) ö £0.04 o o O 0.02 0.00 Cruise Drag Polar CD - 0.042696 + 0.068861 CL2 Idle Descent Drag Polar CD - 0.0427 + 0.0477 CL2 §— 107 KIAS, 2250 RPM i i i i i i i i i i i 1111 I + ""I1 -i—i—t—i-4-,, i i... j--„ j- 0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 Lift Coefficient, CL Figure A41 Idle Descent Drag Polar USAFA CCFTCassna ISOftSOHP Engina: Lycomlng O-320-E20 Propalan McCulay TM74S8/1C172 Mixture: Laanad Walght: 17MB» CartaHaat OFF Flaps: UP Throtlla: IDLE Dal* Basis: Flight Tast 12 j 10 -: J.10 * e eQ O AX -i 2 4- 107 KIAS, 2250 RPM ' ' I ' i ' 1 1 " "I " " I ,1 Ill 1 „I, I 1 I -D-l ,i-l ,1 t I I I I I | t I I 0 10 20 30 40 SO 60 70 80 90 100 110 120 Indicated Airspeed, Vi (KIAS) Figure A42 Idle Descent Penetration Chart 40 w USAFA CCFT CKMIU 15<X150HP Engln.: Lycomkig O-i20i2D Prop.ll« McCul.y TM745W1C172 Mbctura: Lootd W.lgM: 17»*» Cub Hut OFF Flap« UP ThroW« IDLE UmpMiui»: SUnd.rd Oat* Bute: Flight TW« -500 .1000 1500 E a o © u & 'S ».2000 + 8. 107 KIAS, 2250 RPM X. Sea Level 5,000 ft 10,000 ft" -2500 111 ■'i " ■ ■ |' ■ ■ ■ |'''' I " " I''''I'''' I " " I'''' I'''' I'' " I'''' 0 10 20 30 40 50 60 70 80 90 100 110 120 Indicated Airspeed, V, (KIAS) Figure A43 Idle Descent Polar Chart by Indicated Airspeed USAFA CCFT Cram 15O/1S0HP Engln.: lycoming O-120-E2O Proprilw: "eCutay TO74S8/1C172 Mbrtur.: LMmd W.lght 17Mb. CwbHMt OFF Ftap« UP Throtd.: IDLE T.mpM»tur.: 8Und.nl DM. Bute: Flight TMI £ -500 g, Q g-1000 *r c o 8-1500 0 Q i*. o 3 -2000(0 -2500 107 KIAS, 2250 RPM / Sea Level 5,000 ft 10,000 ft' I l|l I I I | I I H|l I I I | I I I I | ■ ■ M | " " | " " | " " |" ' l| I I I I | ■ I" | ■ ' 0 10 20 30 40 SO 60 70 80 90 100 110 120 130 140 True Airspeed, V (KTAS) Figure A44 Idle Descent Polar Chart by True Airspeed 41 USAFA CCFT Caaana 1S0/130HP Engina: Lycoming O-320-E2O Mbrturc Laanad CarhHaat OFF Alrapaad: MKIAS Data Bad«: Flight Taa« Propallac McCautey TM74S8/1C172 W.lghfc 1760 Iba Flapa: UP Throttla: Idla 20 40 60 80 Outside Air Temperature (°F) 0 200 400 600 800 1000 Rate of Descent (ft/min) and Fuel Flow (gal/hr*100) Figure A45 Nonstandard Day Idle Rate of Descent Performance at 65 KIAS U3AFA CCFT Caaana 150/150HP Engln« lycoming O-320-E2O PropaNan McCaulay TM743aV1C172 Mbttun: LMitad Walght: 17*0 ba CarbHaat: OFF Flapa: UP Ainpaad: M MAS Throttla: Idla Data Basia: Flight Taat \ \ V MO« \ \ Pr—inU tud* \\ \ I MOM« \\ \ i »Mt It II -| \$ •\ i l\ •O VV I > 20/. \ V \ \ *••(. \' i \ 0 20 40 60 80 100 Outside Air Temperature (°F) 10 20 30 40 50 Time (mln) and Fuel (gal*10) 60 Figure A46 Nonstandard Day Time and Fuel to Descend at Idle at 65 KIAS 42 USAFA CCFT CtMn 15O/150HP Engine (.»coming O-M0*2D ProH^ WeCul^r TM7*S«/1C1« Hbrtura: LMiwd W.lght 17« Kit OftaHMC OFF Flap« UP AlrapMd: UNAS Throw«: Ml« Flight T—t \ \ 13000 ft \ \ 1<™—unMUk*i \ f- V -1Msfn- \ 1 \ \™ -^£\ \% *t «w« \i =4S MJ_ \ \ i —-, > v \- ^L. \ A 4- \ — •*«ssy V \ 0 20 40 60 80 100 Outside Air Temperature (aF) 10 20 30 40 Distance (nm) 50 60 Figure A47 Nonstandard Day Distance to Descend at Idle at 65 KIAS USAFACCFTC«*na15O/1S0HP Enghw: Lycoming O-320-£2D Prop.ll«: McCautoy TM74MY1C172 Hbrtura: Uamd VWght 17«0fe« CarbHwfcOFF Flap« UP AlrapMd: 107 MAS ThrotUo: 22*0 RPM D«H Baal«: Flight T—t 0 20 40 60 80 Outside Air Temperature (°F) 0 200 400 600 800 1000 1200 Rate of Descent (tt/mln) and Fuel Flow (gal/hrMOO) Figure A48 Nonstandard Day Idle Rate of Descent Performance at 107 KIAS, 2250 RPM 43 USAFA CCFT Cmsna 1S<V190HP Englra: Lycomlng O-320-E2D Prop«««: McCauiay TM743S/1C172 Hbrtura: LMiMd W.lght: 17*0 to. Cub HMC OFF Flap« UP AlrapMd: 107 MAS Throttto: 22» RPM D>UBulK Flight TMI V 12000« \ *MunAattua* v 1 10000n \ 1 •000 L 1( ^ •Ma» 1 f i —fc. 1 J» MA \ \ «.., \\ \\ \ \ 0 20 40 60 80 100 0 Outside Air Temperature (*F) 10 20 30 40 50 Time (min) and Fuel (gaPIO) 60 Figure A49 Nonstandard Day Time and Fuel to Descend at 107 KIAS, 2250 RPM USAFA CCFT C*«*na 150/150HP Englm: Lycombig O-320-E2O Propall«: McCaulcy TM745W1C172 Mbrtura: U*iwd W«gM: 17« k>« CrbHMt OFF Flap« UP AlrapMd: 107 MAS ThroW«: 2290 RPM DaUButa: Flight TMI \ \ 12 «0(1 \ \ P"MW» At «ud. V V \ V \ } i -A\ V \ ^\ Vi 4 i v =4ä tj_2_ X _ \ \ ^. K- -«sjj» \ -~ ■—, J\ 1 V 4jy.«w_ \ \ 20 40 60 80 100 Outside Air Temperature (°F) 10 20 30 40 Distance (nm) 50 60 Figure A50 Nonstandard Day Distance to Descend at 107 KIAS, 2250 RPM 44 USAFA CCFT Casana 1S0/1S0HP Englna: Incoming O-J20«D Propallac McC«ul^TII74S»1C172 70 60 50 -f (Kl 40 30 20 LifflO 0 Mixtur« Laanadabova 5,000 ft HSl Liftoff Alrspsad: S7WAS Standard AWtud« IMUHI Flap«: UP Upper 96% CoiaTdam» Ltal* • «J WAS Standard WalghC 17Mlb< Wind: Calm Carb Haate OFF Data Basis: Flight Taat Maan Liftoff Alrsoaad-87 «AS „o o o Lowsr »»% CoofWanca Um« ■ SI WAS 4- 200 400 600 800 1000 1200 Standardized Ground Roll Distance (ft) Figure A51 Standardized Takeoff Ground Roll Performance USAFA CCFT Cessna 150/150HP Engina: Lycomlng O-320-E2D Propaltor McCauleyTM7458/1C172 Mixture: Laanad abova 5,000 ft MSL Cub Hut OFF LmotTAinpMd: 57 WAS Flaps: UP (Ma Basis: Right Tast 1400 3000 0 50 100 Outside Air Temperature, OAT(°F) 1700 1500 1300 Gross Weight, W (lbs) Figure A52 Mean Takeoff Ground Run 0 10 20 Wind (knots) 45 0 u c sin c o I0 a.w 600 500 400 300 USAFA CCFT Caaana 160/1SOHP Englna: Lyeomlng O-320-E20 Propallar McCaul«yTM74S8/1C172 Mixtur«: Laanad abov« 6,000 « MSI CarbHaatOFF LHtofTAIrapaad: 67KIA8 Flap« UP Data Baala: Right Teat ~ 200 c 0 a0 a. in 100 9fJ^ .»»>*• o»«1 .« I—LJ_ »fl»j^ iS5r" s«i2£L: S3 < i i 1200 1300 1400 1500 1600 Gross Weight, W (lbs) 1700 1800 ou c M 5 c o a.(0 Q « c 0 0 a. 600 500 400 300 200 100 Figure A53 Takeoff 95th Percentile Dispersion USAFA CCFT Casana 160« 60HP Englna: Lycomlng O-320-E2D Propallar: McCaulayTM74S8/1C172 Mlxtura: Laanad abova 6,000 n MSi. 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