wvvw.chinook-helicopter.comwvvw.chinook-helicopter.com/...Profile_Performance_Handbook_CH-… · DISCLAIMER The findings in this report are not to be construed as an official Department

ktVE'LTRASANA

TECHNICAL REPORT NO. 3-78

SFLIGHT PROFILE PERFORMANCE HANDBOOK

VOLUME VIIC-CH-47C (CHINOOK)

APRIL1979

C:4APPROVED FOR PUBLIC RELEASE;DISTRIBUTION UNLIMITED

DEPARTMENT OF THE ARMY

.US--ARI-YRADOC SYSTEMS-ANAL-YSI-S- A-CTiVITYWHITESAt -MISSILERANGE

NEW MEXICO 88002

C=-A~

C=2

DISCLAIMER

The findings in this report are not to be construed as an officialDepartment of the Army position.

WARNING

Information and data contained in this document are based on the inputavailable at the time of preparation. The results may be subject tochange and should not be construed as representing TRADOC positionunless o0 specified.

4,

DEPARTMENT OF IHE ARMYUS ARMY TRADOC SYSTEMS ANALYSIS ACTIVITY

WHITE SANDS MISSILE RANGENEW MEXICO 88002

iri1N1

1-7

TRASANAECHNICAL ýE3-78

FLIGHT PROFILE PERFORMANCE HANDBOOK,*

..OLUME VIIC *,,CH-47C (CHINOOKý,,

•k • Alan JthaM -fe

- -

APRIL 1979

TRASANA -TR-3-78 -VOL-7G

Z DEPARTMENT OF THE ARMYUS ARMY TRADOC SYSTEMS ANALYSIS ACTIVITY

f• WHITE SANDS MISSILE RANGENEW MEXICO 88002

• Lg1 '

ACKNOWLEDGMENT

At AVRAbCOM, Mr. Harold Sell, Mr. James O'Mailey and Mr. Dala Pitt

provided and validated the data in the Handbook. They also assisted in

devising the formats to assure clarity in the data presentation and dis-

cussion.

At TRASANA, Mr. Frank Gonzalez provided help and guidance during

the preparation of the Handbook.

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TABLE OF CONTENTS

Page

Acknowledgment iv

Index of Tables and Figures vi

Chapter 1 - Introduction 1

Chapter 2 - Flight Profile Example 5

Chapter 3 - Performance Data Table Descriptions 11

Chapter 4 - CH-47C Performance Data Tables (235 RPM) 25

Chapter 5 - CH-47C Performance Data Tables (245 RPM) 117

Appendix A - Functions for Calculating Basic Fuel Flow 209

Appendix B - Functions for Calculating Delta Fuel Flow for Drag 217

Appendix C - Functions for Calculating "round Idle Fuel Flow 223

Appendix D - Functions for Calculating Gross Weight Limits 225for Takeoff

Appendix E - Short Description of CH-47C Data Source 231

Accession For

NTiS G&•,A

DDC TABUnn..cn. edJu.Aification_ _

Dist - Upeia/

V

.7•-

i i . .

INDEX OF TABLES AND FIGURES

Page

CH-47 (CHINOOK) vii

Illustration 2-1 - Mission Example 5

Table 2-1 - Flight Plan Example 6

Table 2-2 - Ground Idle Fuel Flow Table 7

Table 2-3 - Basic Fuel Flow 8

Table 2-4 - Completed Flight Plan Example 9

Table 3-1 - Basic Fuel Flow 13

- Table 3-2 - Basic Fuel Flow 14

Table 3-3 - Delta Fuel Flow for Drag 15

Figure 3-1 - Takeoff Criteria 18

Table 3-4 - Gross Weight L uit for Takeoff 20

Table 3-5 - Gross Weight Limit for Takeoff 21

Table 3-6 - Velocity Limits Table 23

Table 3-7 - Expanded Flight Plan Example 24

Tables 4-1 to 4-24 - Basic Fuel Flow Data (235 RPM) 27

Tables 4-25 to 4-48 - Delta Fuel Flow for Drag Data (235 RPM) 53

Table 4-49 - Groupd Idle Fuel Flow Data 81

Table 4-50 to 4-55 - Gross Weight Limits Data (235 RPM) 83

Tables 4-56 to 4-79 - Velocity Limits Data (235 RPM) 91

Tables 5-1 to 5-24 - Basic Fuel Flow Data (245 RPM) 119

Tables 5-25 to 5-48 - Delta Fuel Flow for Drag Data (245 RPM) 145

Table 5-49 - Ground idle Fuel Flow Data 173

Tables 5-50 to 5-55 - Gross Weight Limits Data (245 RPM) 175

Tables 5-56 to 5-79 - Velocity Limits Data (245 RPM) 183

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CHAPTER 1

INTRODUCTION

1. PURPOSE

The purpose for preparing this handbook series is fourfold: (a) tovalidate CHINOOK performance data quickly, (b) to reduce the manpowerand time to prepare accurate flight profiles, (c) to standardize per-formance data so that the analysis community can benefit from a singlereference in conducting studies and (d) to provide a handbook that canbe used for training in the mission profile planning area.

2. BACKGROUND

The CHINOOK performance data contained in this Flight Profile PerformanceHandbook (FPPH) series was originally acquired as a data base for theAircraft Mission Processing Simulation (AMPS) model. AMPS is a computerprogram developed by the Aviation Systems Analysis Branch of the US ArmyTRADOC Systems Analysis Activity (TRASANA) to support Cost and OperationalEffectiveness Analyses (COEAs). AMPS generates detailed flight profilesfor a wide variety of helicopter missions. The data was provided TRASANAby the Army Aviation Research and Development Command (AVRADCOM) and wasthe most accurate data available to AVRADCOM at the time of handbookpublication. !n structuring the data base for AMPS it was noted thatthe data, when properly organized, could provide a method of doing quickand simple flight profile simulations. This volume presents theCHINOOK data and explains how it can be used.

3. OBJECTIVES OF THE HANDBOOK

a. Data Validation. This volume of the handbook contains tableswith the precise performance data and format required to develop flightprofiles for computer simulations. Using the handbooks as a reference,the individual project manager (PM) will be able to quickly validate orupdate as required all associated data contained in the different tabIes.If this procedure is followed by the various PMs, support of HelicopterCOEAs and other analyses can be efficiently implemented.

b. Flight Profile Development. Much of the manpower and time -pentin preparing flight profiles for supporting aircraft COEAs is dedicatedto look-up, correlation and validation of performance data. Once theprocedure contained in this handbook is implemented, flight profiles canbe easily prepared. What normally took one man 4 to 5 days to preparecan now be prepared in 3 to 4 hours.

S, • 1

c. Standardization of Performance Data. Each of the PMs has beencontacted by AVRADCOM to validate the performance data contained ineach handbook in this series. Once each handbook is published, thedata contained will be kept current as of the publication date. Sincethe requests for current information are constantly being forwarded tothe PMs by analysis groups, this handbook can be a reference and assurea commonality in studies within the community.

d. Training for Planning Missions and Flight Profiles. For trainingpurposes each handbook can stand alone. It is only a matter of followingthe example provided and applying the proper data to fit the flightprofile desired. Although the example shown is simplistic, the methodo-logy may he expanded to apply to any flight profile no matter how complex.

4. OTHER VOLUMES

This handbook is one of a series that covers the helicopters in the USArmy inventory. The complete set of handbooks and their subjects are:

Volume I - FPPH Description

Volume II - UH-60A (BLACKHAWK)

Volume III - AH-IG ,COBRA)

Volume IV - AH-lS (COBRA)

Volume V - YAH-64 (Advanced Attack Helicopter [AAH])

Volume VI - OH-58C (KIOWA)

Volume ViT - CH-47 (CHINOOK)

Volume VIII - CH-54 (TARHE)

Volume ix - UH-lH (HUEY)

5. GENERAL HANDBOOK DESCRIPTION

a. Performance Data. The data contained in these volumes is CHINOOKperformance data compiled from the results of actual experimnents. It isnot engineering data and is noi intended to serve as a base for futurehelicopter construction or acquisition. The more mature the helicopterbecomes, the less likely there will be a change in the basic performancedata.

2

b. Handbook Organization. This volume is one of a series of volumesas identified in paragraph 4 above. Volume I is a description of themethodology used to develop the tables for each of the other volumes.This volume and all other volumes except Volume I provides a simplifiedflight profile example in Chapter 2. Chapter 3 provides an explanationof each of the five types of data tables contained in the handbook.The five types of tables deal with: (1) Basic Fuel Flow Data, (2) DeltaFuel Flow for Drag Data, (3) Ground Idle Fuel Flow Data, (4) Gross WeightLimits Data and, (5) Velocity Limits data. Chapter 4 contains the actualtables to be used for developing flight profiles.

c. Volume VII Organization. The US Army has four differentversions of the CH-47 CHINOOK. Due to the large amount of data forthese four versions and to allow for easier reference, there isa separate section of Volume VII for each. Volume VIIA contains datafor the CH-47A. In the same manner, Volume VIIB contains CH-47B data,Volume VIIC contains CH-47C data, and Volume VIID contains CH-47D data,

6. CH-47C OPERATION RATES

The CH-47C engine operates at two different rates which are dependen.ton the aircraft's gross weight. At gross weights of 40,000 lbs or lessthe engine runs at 235 RPM, above 40,000 lbs the rate is 245 RPM. Con-seqUently, separate tables are provided in this volume for the differentRPMs. The tables for 235 RPM are in Chapter 4 of this volume, whileChapter 5 contains the tables for 245 RPM.

3

CHAPTER 2

FLIGHT PROFILE EXAMPLE

1. GENERAL

This chapter provides an example of how to develop a flight profile,albeit simple, that can be extended to cover any number of stops, loadsand distances all depending on helicopter capability and fuel available.

2. DISCUSSION

a. The main question this example of a flight profile will answeris, "Do I have enough fuel to fly the proposed mission?"

b. Suppose a pilot is to fly a simple resupply mission in a CH-47C

CHINOOK helicopter that calls for f'ying (as shown in illustration2-1) from point A (the air base), to point B (the pick up area) to pointC (the drop off area) and return to A.

8 N.M.

A

Illustration 2-1

c. The other information given is airspeed (AS) from A to B whichis to be 70 knots (kts), from B to C 40 kts, and from C to A 70 kts.The CHINOOK helicopter is to be flown, at 4,000 ft for all legs at anambient temperature of 15'C, and an idle altitude for take off, pick-upand drop off areas (ground level) of 2000 ft*. The mission plan alsoshows 10 minutes idle at A before take off, 20 minutes idle at B whileloading, 20 minutes idle at C while unloading and 10 minutes idle onreturn to A before shut down. The CHINOOK will be flown empty at agross weight (GW) of 20,000 lbs from A to B and from C to A, while thecargo from B to C will be 16,000 lbs.

"MAll attitudes are in reference to sea level.

5

CHAPTER 2FLIGHT PROFILE EXAMPLE

1. GENERAL

This chapter provides an example of how to develop a flight profile,albeit simple, that can be extended to cover any number oi stops, loadsand distances all depending on helicopter capability and fuel available.

2. DISCUSSION

a. The main question this example of a flight profile will answeris, "Do I have enough fuel to fly the proposed mission?"

b. Suppose a pilot is to fly a simple resupply mission in a CH-47CCHINOOK helicopter that calls for flying (as shown in illustration2-1) from point A (the air base), to point B (the pick up area) to pointC (the drop off area) and return to A.

80 N..

A

Illustration 2-1

c. The other information given is airspeed (AS) from A to B whichis to be 70 knots (kts), from B to C 40 kts, and from C to A 70 kts.The CHINOOK helicopter is to be flown, at 4,000 ft for all legs at anambient temperature of 150 C, and an idle altitude for take off, pick-upand drop off areas (ground level) of 2000 ft*. The mission plan alsoshows 10 minutes idle at A before take off, 20 minutes idle at B whileloading, 20 minutes idle at C while unloading and 10 minutes idle onreturn to A before shut down. The CHINOOK will be flown empty at agross weight (GW) of 20,000 lbs from A to B and from C to A, while thecargo from B to C will be 16,000 lbs.

•All altitudes are in reference to sea level.

k3

d. The flight plan is prepared by drawirg up a table similar toTable 2-1 below. By filling in the blanks under fuel, it can be deter-mined if the total is too large for the helicopter.

TABLE 2-1

Helicopter: CHINOOK (CH-47C)

Altitude: 4000 ft flight/2000 ft idle

Temperature: 15%C

LEG DISTANCE AS TIME GW (Ibs) FUEL

Idle @ A - - l0 min -

A-B 70 N.M. 70 kts I hr 20,000

Idle @ B - 20 min -

B-C 80N.M. 40 kts 2 hr 36,000

Idle @ C - - 20 min -

C-A 140 N.M. 70 kts 2 hr 20,000

Idle @ A - - 10min

e. FirsL fill in Idle @ A, Idle @ B, Idle @ C and 2nd Idle @ Asince they will all come from Table 2-2. In each case the idle is at2000 ft and a temperature of 15°C. Consulting the ground idle fuelshown in Table 2-2, the value of 1374 lbs/hr is at the intersection of2000 ft and 150 C.

1st Idle @ A = 1/6 X 1374 = 229 lbs

Idle 1 B = 1/3 X 1374 = 458 1bsIdle @ C = 1/3 X 1374 = 458 lbs

1 2nd Idle @ A = 1/6 X 1374 = 229 lbs

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Notice the conversion from minutes to hours. These values must beused because fuel flow is in lbs/hr.

f. The fuel flow for" the thret" legs of the mission are calculatednext. The heading on Table 2-1 shows a need for the Basic Fuel Flowdata chart for the CHINOOK helicopter flying at 4000 ft and at 15%Cambient temperature. Table 2-3 contains the necessary information.

(1) leg A-B is at 70 kts and 20,000 lbs. This is not one of thevalues given but 60 kts is 1362 lb/hr and 80 kts is 1321 lb/hr. Interpolationgives the valu- of 1342 lb/hr for a 70 kts airspeed. Since the leg isone hour long:

Leg A-B = 1 X 1342 = 1342 lbs

(2) Leg B-C is at 40 kts and 36,000 lbs. This value is in thetable; 2259 lbs/hr. Since the leg is two hours long:

Leg B-C = 2 X 2259 = 4518 lbs

(3) Leg C-A is at 70 kts and 20,000 lbs. This fuel flow rate wascomputed above to be 1342 lbs/hr. Since the leg is two hours long:

Leg C-A = 2 X 1342 - 2684 lbs.

g. The flight profile can be finished by filling in Table 2-I asshown in Table 2-4.

TABLE 2-4

Helicopter: CHINOOK (CH-47C)Altitude: 4000 ft flight/2000 ft IdleTemperature: 15%C

LEG DISTANCE AS TIME GW (Ibs) FUEL

Idle @A -0 ri - 229 Tbs

A-B 70 N.M. 70 kts 1 hr 20,000 1342 Tbs

Idle ( 1 - - 20 mrin - 458 Ibs

B-C 80 N.M. 40 kts 2 hr 36,000 4518 lbs

Idle @ C - - 20 min - 458 lbs

C-1 140 N.M. 70 kts 2 hr 20,000 2684 Ibs

j L e- -A -0O 10 rain - 229 lbs

Total 9918 Ibs

9

h. Although only two look-up tables were used for this example,each type of table has several conditions that are changed so that awide band of perfomance parameters can be addressed. The discussion c..each of the five types of tables is contained in Chapter 3. A succinctdescription of each of these five types of tables is:

(1) Basic Fuel Flow Dat,,: Gives the rate the aircraft uses fueldependent on the given flight conditions.

(2) Delta Fuel Flow for Drag Data: Gives the additional rate offuel flow to be added to the basic rate for external drag.

(3) Ground Idle Fuel Flow Data: Gives the rate fuel is used whenthe aircraft is or, the ground with its engine running.

(4) Gross Weight Limits Data: A check on whether or not the aircrafthas enough lift to take off with a given weight.

(5) Velocity Limits Data: Gives the optimum (long range) speed andmaximum rates of speed.

}10

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CHAPTER 3

PERFORMANCE DATA TABLE DESCRIPTIONS

I. GENERAL

This chapter describes each of the five basic type tables used fordeveloping flight profiles. The variables within each type of tanle aredescribed as well as how the specific data required can be extracted.

2. BASIC FUEL FLOW DATA

a. The basic rate u: fuel flow* is detemined by five variables:

(1) Type of aircraft

(2) Altitude (Air Pressure)**

(3) Temperature***

(4) Gross Weight****

(5) Flight Mode

b. In each table (see Table 3-1) within the basic type, the firstthree variables are held constant for the whole table, i.e., (a) Type ofAircraft, (b) Altitude (Air Pressure) above sea level, and (c) Tempera-ture. These vari-ables are stated at the top of each table.

c. There are six rows of fixed gross weights for 235 RPM: 20,000lbs, 24,000 lbs, 28,000 lbs, 32,000 lbs, 36,000 lbs and 40,000 lbs(Table 3-I) There are four rows of fixed gross weights for 245 RPM:40,000 lbs, 42,000 lbs, 44,000 lbs, and 46,000 lbs 'Table 3-2). Theten columns are fixed flight modes.

(1) The first column is Hover In Ground Effect (HIGE). HIGE isused for hovers at a height of 10 feet or less and a component of forwardflight 10 kts or less.

(2) The second column is Hover Out of Ground Effect (HOGE). Thisis used for hovers at a height of more than 10 feet.

7h,1w ba.'lm ic fui* flow , dati c,-ree,,ntes a olean d•a•;. Tiaia',- t 'n *on tha~'; iooi's o-osee,, no 1,1ine stores, ,MI no cxterna? soia l .•Ids.

* %ij; ,titud,,os or. alir pIJssuW'Ce am,? feet zCbove S,30 ! 'el.""For simpi n ty., al'217 teterat:ures arte considered to be the Z.oerauW

tc.-neratwc n which the he"* colter -s oa(Prees Centignade)

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(3) The third column is Nap of the Earth (NOE). This is definedas all flight for variable speeds from 0 to 40 kts and variable altitudes.

(4) The remaining seven columns are for given airspeeds* (in kts) asthe flight mode.

d. There are 24 of these basic fuel flow charts. Each chart is fora different combination of Air Pressure (Altitude) and temperature.

e. The Basic Fuel Flow Data is the main table used in simulatinga flight profile. For example, assume a pilot's flight path will require30 minutes of flight at 80 kts airspeed, 4000 ft. altitude, 15%C and agross weight of 28000 lbs in a CH-47C helicopter. Using Table 3-1 ata gross weight of 28000 lbs and an airspeed of 80 kts, the helicopterwill use 1601 lbs/hr fuel, i.e., for 30 minutes, 801 lbs of fuel will beused.

f. The gross weight values selected provide the basic range ofload carrying capability for the ten flight modes of the CHINOOK heli-copter. Within the gross weight band shown, linear interpolation** isquite accurate for estimating the fuel flow rates.

g. For example, using Table 3-1, if the helicopter's gross weightwas 30,000 lbs and if the flight mode was 60 kts, the fuel flow cannotbe found directly. But by interpolating between 60 kts, 28,000 lbs -1639 lbs/hr and 32,000 lbs - 1811 lbs/hr, the basic fuel flow rate for30,000 lbs is 1725 lbs/hr. In this example, if the helicopter flies inthis mode for 30 minutes, 863 lbs of fuel will be used.

h. As altitude and/or temperature changes occur, different tablesare used to look up the aircraft's basic fuel flow rate for each leg ofthe flight path. Care must be taken that the proper table is used.

i. Appendix A contains a set of functions that will give a goodapproximation of the basic rate of fuel flow.3. DELTA FUEL FLOW FOR DRAG DATA

a. The delta fuel flow for drag is also determined by five variables:

(1) Type of Aircraft

(2) Altitude (Air Pressure)

(3) Temperature

(4) Drag Surface (Equivalent Square Footage)

(5) Air Speed

*All references to airspeeds are to true airspeeds.

**AZZl references to interpolation are lineci interpolations. See FPPHI,Volwne I, Chapter 3 for a discussion on the accuracy of interpolation.

12

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b. Like the basic fuel flow tables, there are 24 tables for deltafuel flow for drag.

c. There are four fixed rows of equivalent square feet of drag:50 equivalent sq ft thru 200 equivalent sq ft.

d. The seven columns are for airspeeds in kts of: 40 kts, 60 kts,80 kts, 100 kts, 120 kts, 140 kts, and 160 kts.

e. When an external load is placed on the helicopter, the imount offuel consumed per hour increases. The delta Fuel flow for drag tablesindicate how much extra fuel consumption to add to the basic fuel flowrate.

f. In the example given earlier, a 30 minute flight at 80 kts airspeed,4000 ft altitude, 150C and a gross weight of 28,000 lbs was used. Usingthe basic fuel flow tables, the basic fuel flow rate was 1601 lbs/hr.Assuming for this new example that part of the load is external andinducing a 100 equivalent sq ft external drag, the delta fuel flow fordrag (Table 3-3) shows 208 lbs/hr should be added to the basic fuel flowrate. Thus the basic fuel flow rate becomes 1601 + 208 or 1809 lbs perhour and for a half-hour flight, 905 lbs of fuel will be used instead ofthe 801 lbs figured without an external load.

g. Appendix B contains a function that will give a gooc approximation

of the delta fuel flow for drag.

4. GROUND IDLE FUEL FLOW DATA

a. The ground idle fuel flow rate is determined by only threevariables:

(1) Type of Aircraft


(3) Temperature

b. There is only one ground idle fuel flow table (shown as Table 2-2).The table has four rows of temperatures: -25°C, -5*C, 15%C and 350C,and six columns of altitudes: Sea Level, 2000 ft, 4000 ft., 6000 ft.,8000 ft., and 10000 ft.

c. The ground idle fuel flow table is used as discussed in theexample flight profile in Chapter 2 (Table 2-2). The CH-47C helicop.teridling for 20 minutes at 2000 ft. altitude and 15°C, (across the rowlabeled 15%C and down the column labeled 2000) find the intersection at1374. Thus, the CH-47C uses 1374 lbs/hr at these conditions and since itis idling for 20 minutes or 1/3 of an hour, it will use 458 lbs of fuel.

16

d. If the hel i copter hdd only been 1000 ft. above sea level, the con-sumpt ion rate would be found by interpolatiiui betwveen the sea level rateof 1454 lbs/hr and the 2000 ft. rate of 1374 lbs/hr which would be 1414lbs/hr. In 1/3 of an hour 471 lbs of fuel would be used.

e. Appendix (C contains a function that will give a good approxima-t ion of the tiround idle fuel flow.

1). tlt'OSS w IGtIl LIMITS DATA

a. (iross we ight limits tables are intenided to show whether or nlotthe aircraft can safely take off for four sets of criteria. Thesecriteria are defi ned in the following paragraphs:

(I) C-r'teria #1 is based on the helicopter using 100% of MaximumPower for take oft and having enough power to lift straight tip and aboveqround effect (See Figure 3-1). Once it is in hovering above grOundeffect level the helicopter begins forward flight until it acquires,transitional lift and is able to climb at 450 ft/min (a desired standardrate of climb) to the desired altitude. This criteria has some risksirce the pilot has no reserve power. It has less risk than Criteria #3but more than Criteria #2 thus it is considered to be "Middle of theRoad" risk.

(2) Criteria #2 (Figure 3-1) is based on the helicopter using 951of ,1aximumi Power for take off and enough power to ininediately begin toclimb at a rate of 450 ft/min. This is the least risky criteria sincethle pilot has power in reserve and is still able to climb at a sat i sfactoryratCe.

(3) Criteria #3 (Figure 3-1) has the miost risk. Using 1001% of MaximumPower tile helicopter will only hover in ground effect. Therefore, atan altitude of 10 feet or less, the pilot nust tegi in forward flightand gradually inc rease airspeed to acquire transitional lift to climb.The reasons for its high risk are readily apparent. First. there is nopower in reserve. Second. the pilot must begin forward flight at avery low altitude.

(4) Criteria #4. Structural Gross Weight Limits is the total upperl imit of gross weight the helicopter can carry under any take off criteria.

b. Gross Weight Limits are determined by four variables:

(I) Type oi Aircraft

(2) Criteria Chesen


(4) Tei:nperature

17

CRITERIA #1(MIDDLE OF THE ROAD)

100% MAX POWER, HOGE,FLIGHT

TRANSITIONJAL LIFT

HOGE

GROUND

NOTHING TO SPARE.

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(MOST RISKY)

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j FLIGHT

TRANSITIONAL LIFT

HI.GE

NOTHING TO SPARE.

Figure 3-1

18

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c. Additionally, Criteria #P, #2, and #3 differ due to engine powerlimits or transmission power limits of the aircraft. Thus there are sixtables:

(1) Criteria #1 (Due to engine)

(2) Criteria #1 (Due to transmission)


(4) Criter;1 #2 (Due to transmission)


(6) Criteria #3 (Due to transmission)

d. The structural gross weight limit is a single value for eachhelicopter and is only depe.,dent on the type helicopter. The CH-47Cstructural gross weight lim',t is given as 46,000 lbs and is listed atthe bottom of each table. As the name implies, it is simply not safe toexpect the CH-47C structure to maneuver normally when the tota'Fweightis larger than that value.

e. In simulating inflight profile, the gross weight limits tablesare used to check whether the aircraft is going to be too heavy to takeoff under the given conditions. As an example, assume the pilot of aCH-47C planned a mission that calleo for using take off criteria #1 and thetake off was to be at 8000 ft., 150 C, and a gross weight of 38,200.Three checks would be required: First, does this gross weight exceedthe structural gross weight limit? Second, does it exceed Criteria #1(due to transmission)? Third, does it exceed Criteria #1 (due to engine)?in the example given, the answer to all three questions is "No", thetake off will not exceed aircraft limits. (Tables 3-4 and 3-5)

f. If the assigned gross weight had been 42,000 lbs, it would haveexceeded the value given for 8,000 ft. and 150C at Criteria #1 (Dueto engine). (Table 3-4) The mission could not be flown as planned.The plan could be changed, for example to take off at 6000 ft. (whichmight not be practical) or change to take off Criteria #3 (which is morerisky but has higher limits).

g. If the assigned gross weight had been 46,300 lbs., it would haveexceeded the structural limits. To perform the mission the only choiceswould be to lighten the load or get another type helicopter.

h. Appendix 0 contains a set of functions that will give a goodapproximation of the gross weight limits for takeoft.

19

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6. VELOCITY LIMITS DATA

a. There are various types of data given in these tables but likethe gross weight limits tables, they are primarily restraints on whatcan be expected of a helicopter in planning a mission profile. Velocitylimits tables are influenced by five variables:

(1) Type of aircraft

(2) Air pressure (altitude)

(3) Temperature

(4) Gross weight

(5) Condition or limit

b. Items (1) through (4) are self-explanatory. There are five typesof information that can be listed under (5):

(1) Long range

(2) Maximum continuous power

(3) Maximum power (due to engine limits)

(4) Transmission limits

(5) Vne (velocity never exceed)

c. For each aircraft, there are 24 Velocity Limits Tables dependingon air pressure and temperature combination. Table 3-6 is an example ofthe content of the Velocity Limits Tabi,.

d. The two columns under Long Range (Table 3-6) give the optimumspeed and fuel flow for each set of variables #1 through #4 above. Thusthe CH-47C operating at 2000 ft., temperature 15°C, and having a grossweight of 28,000 lbs will fly a longer distance if the velocity iskept at 125 kts and will use 2072 lbs/hr of fuel at that velocity.

e. Maximum continuous puwer gives the fastest speed at which ahelicopter can fly for long periods (30 minutes or more) and the associatedfuel flow rate. Ar. example fran Table 3-5 would be a CH-47C at 2000ft. and 150 weighing 28,000 lbs could fly 167 kts with a fuel usageof 3334 lbs/hr.

22

.1 3.4 0 - - .

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0 0

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23W

f. Maximum power (engine and transmission limits) show the maximumspeeds the aircraft can structurally attain for short periods of time(less than 30 minutes). Thus the CH-47C helicopter at 2000 ft and150 C weighing 28,000 lbs has an engine that is capable of producingenough power to fly 183 kts but the transmission limits the aircraft to170 kts. Between these two columns then, the flight cannot exceed 170kts with a fuel flow rate of 3450 lbs/hr.

g. There is another limiting factor called V (velocity neverexceed). This velocity limit is determined by helopter structuralconsiderations. V 's are used in the same manner as maximum powerlimits described in naragraph f above. Since a value of 170 kts islisted for 2,000 ft., 15°C, and 28,000 lbs, this implies that theaircraft can reach its transmission limit under these conditions.

7. DETAILED FLIGHT PROFILE USING ALL PERFORMANCE DATA TABLES

The example of a Flight Profile in Chapter 2 was intentionally simplifiedto assure clarity. The description of the various tables in this hand-book, however, indicates a more complex set of considerations are normallyencountered in developing the flight profile. With the descriptionprovided in this chapter, additional information should be included inthe flight plan beyond that shown in the example and a suggested formatis provided below in Table 3-7.

TABLE 3-7

Helicopter:Alt itude:Temperature:

LEG DISTA/ICE AS CHECK TIME GC (LBS) DRAG FUELVEL.OCITYLIMIT

Needed for each take off:Weight at take off:Type of take off:Check transmission limits:Check engine limits:Check structural gross weight limit:

24

CHAPTER 4

CHINOOK (CH-47C) PERFORMANCE DATA TABLES (235 RPM)

GENERAL

The following tables are the major information presented in this hand-

bcok. If the procedure for using them is understood, a flight profile

for the CHINOOK (CH-47C) helicopter can be prepared in a matter of a few hours.

The performance data contained have been reviewed for accuracy and are

corrected to the best of our knowledge. The tables are organized in

the following manner:

Tables 4-1 to 4-24 Basic Fuel Flow Data

Tables 4-25 to 4-48 Delta Fuel Flow for Drag Data

Table 4-49 Ground Idle FuE! Flow Data

Tables 4-50 to 4-55 Gross Weight Limits Data

Tables 4-56 to 4-79 Velocity Limits Data

25

BASIC FUEL FLOW DATA

TABLES

(235 RPM)

27

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DELTA FUEL FLOW FOR DRAG DATA

TABLES

(235 RPM)

Ii

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I ix

D2 Ln A N

-w o. (L 0

CHAPTER 5

CHINOOK (CH-47C) PERFORMANCE DATA TABLES (245 RPM)

GLNFRAL

These tables are the additional ones needed when the CH-47C is operatedat a gross weight in excess of 40,000 lbs. These are for 245 RPM engineusage and are supplemental to the tables in Chapter 4. The tables areorganized in the following manner:

Tables 5-1 to 5-24 Basic Fuel Flow Data

Tables 5-25 to 5-48 Delta Fuel Flow for Drag Data

Taole 5-49 Ground Idle Fuel Flow Data

Tables 5-50 to 5-55 Gross Weight Limits Data

Tables 5-56 to 5-79 Velocity Limits Data

1

1 1

[ 117

BASIC FIEL FLOW DATA

TABLES

(245 RPM)

pREOEDUNG PAGE IL.AN

119

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DELTA FUEL FLOW FOR DRAG DATA

TABLES

(245 RPM)

•145

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m l = -- • . . . . . . .. ... . . .. ... . . ... ..... ... . . .. . . . . .. . . .. ... . . . . . . . . ... .. .. .. . . . . . .. .. . . .. . . .. . . . . . . . . .. . . . . .. . . .. . . . . . . . .. .. . .

GROUND IDLE FUEL FLOW DATA

TABLE

171

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GROSS WEIGHT LIMITS DATA

TABLES

(245 RPM)

1753_

4)0. N ly.

)~4 v) 0'

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VELOCITY LIMITS DATA

TABLES

(245 RPM)

1

t 183

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APPENDIX A

FUNCTIONS FOR CALCULATING BASIC FUEL FLOW

209

1. CH-47C Operating at 235 RPM

There are four functions that can be used to calculate the basic fuelflow for the CH-47B helicopter operating at 235 RPM. In order to usethe functions the following data is needed:

1. Flight Mode

2. Temperature

3. Pressure (altitude)

4. Gross weight

Which of the four functions will be used depends on the flight mode.The first function is for HIGE (Hover In Ground Effect).

FF (HIGE) = f (TEMP, ALT, GW)

The second function is for HOGE (Hover Out of Ground Effect).

FF (HOGE) = f (TEMP, ALT, GW)

The third function is for NOE (Nap of the Earth).

FF (NOE) = f (TEMP, ALT, GW)

The fourth function is for Forward Flight.

FF (Forward Flight) = f (AS, TEMP, ALT, GW)

The equation for FF (HIGE) is:

FF (HIGE) = A (ALT) + B (TEMP) + C (GW) + D (ALT)(TEMP)+ E (ALT) (GW) + F (TEMP) (GW)+ G (ALT) (TEMP) (GW) + K

Where ALT is the altitude, TEMP is the temperature and GW is thegross weight and the constants have the following values:

A = -4.18033488 X lO-2 E = 1.11825291 X 10-6

B = 1.08751586 F = 8.07031975 X 10-5

C = 5.33395773 X 10-2 G = 1.36606525 X 10-8

D = -3.51402949 X 10-4 K = 4.84275543 X 102

210

The equation for FF (HOGE) is exactly the same form as FF (HIGE).A new set of values for the constants is used. These values are:

A -5.88115812 X 102 E = 1.93141847 X 1O06

B -8.84728134 X 10"2 F = 1 .38394884 X l0-4

C 6.1890916 X 10-2 G = 1.60426457 X 10-8

D -3.78898469 X 10-4 K = 4.17171783 X 102

The equation for FF (NOE) is once again the same as FF (HIGE). Thenew values for the constants are:

A = -6.06951821 X 102 E = 1.82329043 X 10-6

B = -7.06558749 X 10-2 F = 1.16511314 X 10-4

C = 5.2232069 X 10-2 G = 1.72442876 X 1O"8

D = -4.1362632 X 10-4 K = 5.56116821 X 102

For the Forward Flight modes the form of the equation is:

FF A(AS) + B(AS2) + C(AS3 ) + D(TEMP) + E(GW) + F(ALT) + G(AS3 )(TEMP)

+ H(AS 2 )(TFMP) + I(AS)(TEMP) + J(AS3 )(GW) + K(AS 2 )(GW)

+ L(AS)(GW) + M(AS3 )(ALT) + N(AS 2 )(ALT) + 1n(AS)(ALT) + P(TEMP)(GW)

+ Q(TEMP)(ALT)I + R(GW)(ALT) + S(TEMP)(GW)(ALT) + T

Where AS is the air speed in kts ard the values of the constants are:

A = -3.77404814 X 10 K =-1.8393742 X 10-

B = 4.24010076 X 10-1 L = 1.12421578 X 10-3

C = -1 .01004529 X 10-3 M -1 .37884145 X 10-7

D = 1.03554213 X 10 N = 3.06267834 X 10-5

E = 1.26485201 X 102 0 =-2.34560855 X 10-

F.=F-2.29212269 X 102 P = 3.11250682 X 10-4

G = -3.12955658 X 10-5 Q = 4.4806207 X 10-4

H = 716897036 X 10- R = 2,23874594 X 106I = -5.56602478 X I0-1 S =-4.34590968 X 10-8

V J = 7.77223859 X I0-8 T = 1.8023027 X 103

211

1 - m

These functions allow anyone with a simple calculator to figurethe fuel flow of the aircraft and bypass both looking up the valuesand interpolating for points in between the data points in the tables.

The above equations calculate the basic fuel flow for the CH-47Chelicopter operating at 235 RPM with the following accuracies:

FF (HIGE) - 99.64%

FF (iHOGE) - 99.33%

FF (NOE) - 98.19%

FF (Forward Flight) - 93.86%

212

I

2. CII-47C Operating at 245 RPM

There are four functions that can be used to calculate the basic fuelflow for the CH-47B helicopter operating at 245 RPM. In order to usethe functions the following data is needed:

1. Flight Mode

2. Temperature

3. Pressure (altitude)

4. Gross weight

Which of the four functions will be used depends on the flight mode.The first function is for HIGE (Hover In Ground Effect).

FF (HIGE) = f (TEMP, ALT, GW)

The second function is for HOGE (Hover Out of Ground Effect).

FF (HOGE) = f (TEMP. ALT, GW)

The third function is for NOE (Nap of the Earth).

FF (NOE) = f (TEMP, ALT, GW)

The fourth function is for Forward Flight.

FF (Forward Flight) = f (AS, TEMP, ALT, GW)

The equation for FF (HIGE) is:

FF (HIGE) = A (ALT) + B (TEMP) + C (GW) + D (ALT)(TEMP)+ E (ALT) (GW) + F (TEMP) (GW)+ G (ALT) (TEMP) (GW) + K

Where ALT is tha altitude, TEMP is the temperature and GW is thegross weight and the cunsta..ts have tne following values:

A = -. .49443232 X lu' I1 . = 3.76395468 X 10-6

B = -4.85268086 F = 2.2449065 X 10-4

C = 5.69947064 X i0-2 G = 1.40213223 X 1O-8

D = -3.51416853 X 1-4 K 3.89716461 X 1O0

213

The equation for FF (HOGE) is exactly the same fonn as FF (HIGE).

A new set of values for the constants is used. These values are:

A = -2.29995189 X 10-1 E - 6.22147991 X 10-6

B =-5.28438944 F -2.65626702 X 10'2 8c 7.02132583 X 10-2 G -2.42121925 X 10-8

0 = -7.11802662 X 10- K = 1.28996002 X I12

The equation for FF (NOE) is once again the same as FF (HIGE). Thenew values for the constants are:

"A - -2.04558648 X 10-1 E = 5.37683025 X 10-6

B - -2.73703614 F - 1.7907843 X 10-4

C = 6.0988307 X 102 G - 4.52854692 X 10-8

D = -1.5812756 X 1O"0 K - 2.78312539 X 102

For the Forward Flight modes the form of the equation is:FF :A(AS) + B(AS2 ) + C(AS3 + D(TEMP) + E(GW) + F(ALT) + G(AS 3 )(TIMP)

+ H(AS 2)(TEMP) + I(AS)(TEMP) + J(AS 3)(GW) + K(AS 2)(GW)

+ L(AS)(GW) + M(AS 3 )(ALT) + N(AS 2 )(ALT) + O(AS)(ALT) + P(TEMP)(GW)

+ Q(TEMP)(ALT) + R(GW)(ALT) + S(TEMP)(GW)(ALT) + T

Where AS is the air speed in kts and the values of the constants are:

A = 1.4591177 X 10 K 3.95673749 X 10-6

B -1.27493959 X 10-1 L -8.94904137 X 10-

c 6.56571239 X 10- M -8.69541026 X 10-8

D 3.87899423 N 2.29792088 X 10-5

E 8.85471553 X 10-2 0 -2.08567723 X 10-3

F 6.71030849 X 10-3 p 4.72741376 X 10-5

G =-9.21962567 X 106 Q 1.11335551 X 10-

H = 4.26615639 X 10-4 R - 1.22879437 X i0-6

I =-1.24208927 X 10-2 S =-3.21402496 X 10"8

J - 4.83834617 X 10-9 T =-3.04298401 X 102

214

These functions allow anyone with a simple calculator to figurethe fuel flow of the aircraft and bypass both looking up the valuesand interpolating for points in between the data points in the tables.

The above equations calculate the basic fuel flow for the CH-47Chelicopter operating at 245 RPM with the foltowing accuracies:

FF (HIGE) - 98.26%

FF (HOGE) - 97.70%

FF (NOE) - 97.39%

Fr (Forward Flight) - 98.17%

215

APPENDIX B

FUNCTIONS FOR CALCULATING DELTA FUEL FLOW FOR DRAG

-i.

217


The function below will calculate the delta fuel flow for drag forthe CH-47C helicopter operating at 235 RPM. Recall from the discussionin chapter three that this value is added to the basic fuel flow valuewhenever drag is increasing the rate of fuel flow.,

In order to use the function the followirg data is needed:

1. Air Speed (AS)

2. Equivalent Square Footage of Drag (SQ)

3. Temperature (TEMP) in degrees centigrade

4. Altitude (ALT) in feet above sea level

That is:

FF (Drag) = f(AS, SQ, TEMP, ALT)

The equation for FF (Drag) is:

FF (Drag) = A(AS) + B(AS2 ) + C(AS 3) + D(TEMP) + E(SQ) + F(ALT)

+ G(AS3 )(TEMP) + H(AS 2 )(TEMP) + I(AS)(TEMP) + J(AS3 )(SQ) + K(AS2)(SQ)

+ L(AS)(SQ) + M(AS 3)(ALT) + N(AS 2 )(ALT) + O(AS)(ALT) + P(TEMP)(SQ)

+ Q(TEMP)(ALT) + R(SQ)(ALT) + S(SQ)(ALT)(TEMP) + T

Where the constants have the following values:

A = 1 .92351666 K =-1.84985049 X l103

B = -1.58761502 X 1O"2 L = 1.34020805 X 1l01

C = 1.22072934 X 1O-4 M = -3.96785356 X 10-8

D = 6.74994808 N = 5.21734358 X 10-6

E = -1 .57020617 0 = -4.03765589 X l0-4

F = 4.01374176 X 10- 2 P = -3.25795538 X 10- 2

G = -1.169635 X 10-5 Q = -1.83679713 X 10-5

H = 2.24108415 X l0-3 R = -2.4964305 X l0-4

I = -1.54114246 X 10- S = 1.04480392 X 10-6

J = 1.2006097 X 10-5 T =-2.27549515 X 102

*.There is no delta fuel fiow for drag for HIGE, HOGE or NOE flight

218

This equation calculates the delta fuel flow for drag value withan accuracy of 99.56%. It should be noted that in sone Instances thecomputed value will be negative. If this 'Jccurs, zero (I) should beused as the value for delta fuel flow.

521L

S219


The function below will calculate tiq delta fuel Flow for drag forthe CH-47C helicopter operating at 245 RPM. Recall from the discussionin chapter three that this value is added to the basic fuel flow valuewhenever drag is increasing the rate of fuel flow.*

In order to use the function the following data is needed:

1. Air Speed (AS)

9. Equivalent Square Footage of Drag (SQ)



That is:

FF (Drag) = f(AS, SQ, TEMP, ALT)

The equation for FF (Drag) is:

FF (Drag) = A(AS) + B(AS 2) + C(AS 3) + D(TEMP) + E(SQ) + F(ALT)

+ G(AS 3 )(TEMP) + H(AS 2 )(TEMP) + I(AS)(TEMP) + J(AS3 )(SQ) + K(AS 2 )(SQ)

+ L(AS)(SQ) + M(AS 3)(ALT) + N(AS 2 )(ALT) + O(AS)(ALT) + P(TEMP)(SQ)

+ Q(TEMP)(ALT) + R(SQ)(ALT) + S(SQ)(ALT)(TEMP) + T


A = 1.43925276 X 10 K = -3.73952542 X l0-4

B =-1.76786033 X 10l L = 9.51766968 X l0-

C = 7.34766239 X lO"4 M = -1.05205952 X 10-7

D =-3.62632334 N = 2.35495875 X 10-

E = 1.40438998 0 = -1.89449638 X 10-3

F = 7.3855726 X 10-2 P = -3.31104305 X 1O-2

G = 7.62190552 X 10- 6 Q = 3.00417855 X 10-5

H =-3.03890294 X 10-3 R = -2.39477551 X 10-4

I = 2.72080421 X 10- S = 1.2046344 X lO 6

J = 6.97609534 X 1O- 6 T = -5.04253357 X 1O2

*There is no delta fuel flow for dra. for HIGE, HOGE or NOE flight.

220

This equation calculates the delta fuel flow for drag value withan accuracy of 99.59%. It should be noted that in some inttances thecomputed value will be negative. If this occurs, zero (0) should beused as the value for delta fuel flow.

_ 221

APPENDIX C

FUNCTION FOR CALCULATING GROUND IDLE FUEL FLOW

223

The function below will calculate the ground idle fuel flow rate forthe CH-4% h.?licopte-. In order to use the function the following datais neeied:

1. Temperature (TEMP) in degrees centigrade.

2. Altitude (ALT) in feet above sea level.

That is:

FF (Idle) = f (TEMP, ALT)

The equation, for FF (Idle) is:

FF (Idle) = A(TEMP) + B(ALT) + C(TEMP)(ALT) + D(TEMP ) + E(ALT ) + F


A = -6.6749985 X 101- D = -1.24999922 X I0- 3

9 = -5.5428531 X 10-2 E = 9.99996317 X 10- 7

C = -3.00133252 X 0-I11 F = 1.47358652 X 103

This equation calculates the ground idle fuel flow rate with anaccuracy of 99.67%.

.2

'• ~224

APPENDIX D

FUNCTIONS FOR CALCULATING GROSS WEIGHT LIMITS FOR TAKEOFF

f,,5

S. i 225


The functions given below will calculate the gross weight limitsfor take off for the CH-47C helicopter operating at 235 RPM. Each ofthe functions is of the same basic form with the values of the constantschanging depending on which take off criteria is being used. In allcases the Structural Gross Weight Limit of the CH-47C helicopter is46,00G lbs.

In order to use the functions the following data is needed:



That is:

GW (Limit) = f (TEMP, ALT)

The basic equation for GW (Limit) is:

GW (Limit) = A(TEMP) + B(ALT) + C(TEMP)(ALT) + D

For take off criteria #1 the equation must be used twice, once usingthe engine limit constants and once using the transmission limit constants.For take off criteria #1 the constants for engine limits are:

A = -2.04521187 X 1O2 C = 6.45157177 X lO3

B = -1.73651493 D = 5.48574741 X lO4

For take off criteria #1 the constants for transmission limits are:

A = -5.41285706 X 10 C =-2.38285902 X l0-4

B = -5.78837119 X 10-1 D = 4.65754517 X lO4

For take off criteria #2 two checks must also be made. The constantsfor engine limits, take off criteria #2 are:

A = -1.91924759 X 102 C = 6.05328596 X 103

B = -1.62188777 D = 5.1224647 X 104


A = -4.7751194 X 10 C =-4.11927958 X 104

B = -5.19843929 X 101- D = 4.46686484 X l04

226

Also for take off criteria #3 two checks must be made. The constantsfor engine limits, take off criteria #3 are:

A -2.30310486 X 102 C = 7.26843113 X 103

8 = -1.94757777 D = 6.15208135 X l04


A = -6.06521425 X 101 C = -3.0357156 X l104

B = -6.45660669 X 10"l D = 5.21943193 X 104

This equation with the various sets of constants gives results thatare 99.89% accurate or better.

227

2. CH-47C Operating at 245 RPMThe functions given below will calculate the gross weight limits

for take off for the CH-47C helicopter operating at 245 RPM. Each ofthe functions is of the same basic form with the values :f the constantschanging depending on which take off criteria is being used. In allcases the Structural Gross Weight Limit of +he CH-47C helicopter is46,000 lbs.

In order to use the functions the following data is needed:



That is:

GW (Limit) = f (TEMP, ALT)

The basic equation for GW (Limit) is:

GW (Limit) = A(TEMP) + B(ALT) + C(TEMP)(ALT) + D

For take off criteria #1 the equation must be used twice, once usingthe engine limit constants and once using the transmission limit constants.For take off criteria #1 the constants for engine limits are:

A = -2.00498346 X 102 C = 5.51400252 X 10-3

B = -1.72648424 D = 5.52583125 X 104


A = -4.92673783 X 10 C = -5.14857456 X 10-4

B = -5.39397113 X 10-1 D = 4.64107769 X 10l

For take off criteria #2 two checks must also be made. The constantsfor engine limits, take off criteria #2 are:

A - -1.87090488 X 102 C = 5.07593085 X l0-3

B = -1.60321172 D = 5.13487373 X l04


A = -4.23669033 X 10 C = -7.50785934 X 10- 4

B = -4.76228192 X 10"I D = 4.43537026 X 104

228

Also for take off criteria #3 two checks must h, rn-de The constantsfor engine limits, take off critEria #3 are:

A = -2.25033333 X 102 C = 6.16950123 X 10"3

B = -1.93471529 D = 6.19340469 X i0


A - -5.60299997 X 10 C = -4.62999953 X 10-4

B z -6.05156399 X 10-1 D = 5.20308398 X 104

This equation with the various sets of constants gives results thatare 99.88% accurate or better.

2

• i 229

S.

APPENDIX E

SHORT DESCRIPTION OF CHINOOK (CH-47C) DATA SOURCE

L•!

231

DRDAV-EQA (A)SUBJECT: Short Description of CH-47C Performance Data Provided to TRADOC

Systems Analysis Activity (TRASANA)

MFR:

1. References:

a. United Kingdom CH-47C, Hover-out-Grovtd Effect (HOGE), Power Requiued(Boeing Vertol IOM 8-7442-1-439).

b. Determination of the Effects of Rotor Blade Compressibility on theperformance of the UH-IF; FTC-TR-65-17.

c. Airworthiness and Flight Characteristics Test, CH-47C Helicopter (Chinook)USAASTA Project No. 66-29.

d. Operator's Manual, Army Model CH-47B and CH-47C Helicopters, TM55-1520-227-10.

2. The performance data presented to TRASANA is the result of comikiningthe helicopter power required, engine power available and engine fuel flowcharacteristics. The CH-47C power required was calculated from a non-dimen-sional representation of engine power required (coefficient of power) v,s,gross weight (coefficient of thrust) and true airspeed (advance ratio).The non-dimensional power required was obtained from reference la and Ic. Allperformance in ground effect represents a 10 foot skid height. A temperaturedependent correction, based on the method outlined in reference lb, wasmade to the power required to account for compressibility which could not beaccounted for in the non-dimensional representation.

3. The T55-L-II engine power available to the CH-47C (which was used incombination with the power required to fijd helicopter take-off and speedlimits) was used as a function of altitude and temperature, from reference ic,

4. The engine fuel flow at a particular altitude and temperaturecombination was derived from a representative referred fuel flow as afunction of referred engine power. The referred fuel flow curve for theT55-L-ll engine was taken from reference ic. The calculated fuel flowsreflect 5% conservatism. A referred parameter is one which is divided bytemperature and pressure ratios in order to represent all atmospheric conditionsby one function.

5. The never exceed speeds (Vn.e.) were calculated from those shown graphicallyin reference ld.

6. The Structu.al Gross Weight limit of the CH-47C is 46000 lbs,

AMES A. O'HRALLEY hiitruc & Aeromech Br

233-- EIN -PAU i"M-A|N

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