NAVAL POSTGRADUATE SCHOOL Monterey, California5. BELL/BOEING VERTOL 25-PASSENGER TILT ROTOR -----61...
218
NAVAL POSTGRADUATE SCHOOL Monterey, California DTIC ELECTEf MAY2 1086 2 * THESIS a- A TRADE-OFF STUDY OF TILT ROTOR AIRCRAFT VERSUS HELICOPTERS USING VASCOMP II ANP RESCOMP by __ Thomas P. Walsh March 1986 Thesis Advisor: Max F. Platzer %0 Approved for public release; distributilon is mnlirvited 86 521 048
Transcript of NAVAL POSTGRADUATE SCHOOL Monterey, California5. BELL/BOEING VERTOL 25-PASSENGER TILT ROTOR -----61...
NAVAL POSTGRADUATE SCHOOLMonterey, California
DTICELECTEfMAY2 1086 2 *
THESISa- A TRADE-OFF STUDY OF
TILT ROTOR AIRCRAFT VERSUS HELICOPTERSUSING VASCOMP II ANP RESCOMP
by
__ Thomas P. Walsh
March 1986
Thesis Advisor: Max F. Platzer %0
Approved for public release; distributilon is mnlirvited
86 521 048
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REPORT DOCUMENTATION PAGEIa. REPORT SECURITY CLASSIFICATION lb. RESTRICTIVE MARKINGS
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6a. NAME OF PERFORMING ORGANIZATION 6b OFFICE SYMBOL 7a. NAME OF MONITORING ORGANIZATIONI (if appliable.)Naval Postgraduate School Code 67 Naval Postgraduate School
6C. ADDRESS (City, State, aNd ZIPCode) 7b. ADDRESS (C/y, State, and ZIP Code)
Monterey, California 93943-5100 Monterey, California 93943-5100
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11 TITLE (Include Security CaSufiCation)
A TRADE-OFF .3TUDY OF TILT ROTOR AIRCRAFT VERSUS HELICOPTERS USING VASCOMP II AND HESCOMP
12 PERSONAL AUTHOR(S)Walsh, Thomas P.
13a TYPE OF REPORT I13b TIME COVERED 14 DATE OF REPORT (Year, Month, Day) 5S PAGE COUNTMaster's Thesis FROM TO______ . 1986 March ,-218
16 SUPPLEMENTARY NOTATION
17 COSATI CODES 18 SUBJECT TERMS (Continue on reve"se if necesary and identify by block number)
FIELD GROUP SUBGROUP VASCOMP
HESCOMP. TILT ROTOR
V9ABSTRACT (Continue on revere iaf ne•ary and uaenttfy by block number)Trade-off studies were conducted wherein two versions of tilt rotor aircraft were examinedto determine optimum mission distances where the tilt rotor designs were superior to acomparable contemporary (pure) helicopter. Two FORTRAN computer programs (VASCOMP II andHESCOMP)-developed under contract for NASA Ames Research Center by the Boeing VERTOL Companywere used to predict aircraft performance. Program results were validated using data fromindependent sources. A simplified user's manual is included (with sample data and programoutput) for VASCOMP II.use at the Naval Postgraduate School, Monterey, California.
20 DISTRISUTION/AVAILABILITY OF ABSTRACT 21 ABSTRACT SECURITY CLASSIFICAT'ON['•UNCLASSIFIEDAJNLIMITED 0 SAME AS RPT 0DTIC USERS unclassified
22a '4AME OF RESPONSIBLE INDIVIDUAL 22b TELEPHONE (include Areia Coe) 22c OFFICE SYMBOLProf. Max F. Platzer (408) 646-2941 Code 67
0D FORM 1473,84 MAR 83 APR edition may be used until exhausted SECURITY CLASSIFICATION OF THIS PAGEAll other editir ns are obsolete
1
a,
Approved for public release: distribution is unlimited.
A Trade-Off Study of Tilt Rotor Aircraft versus HelicoptersUsing VASCOMP II and HESCOMP
by
Thomas P. WalshCaptain, United States Army
B.S., Norwich University, 1976
Submitted in partial fulfillment of thereauirements for the decree of
MASTER OF SCIENCE IN AERONAUTICAL ENGINEERING
from the
NAVAL POSTGRADUATE SCHOOLMarch, 1986
Author: ILThomas P. Walsh
Approved by: ./iA'-4 "Max F. Platzer, Thesis Advisor
Max F. Platzer, Chairman,Department of Aeronautics
John N. Dveý, DeAn of Science and Encineerina
2
ABSTRACT
Trade-off studies were conducted wherein two versions of
tilt rotor aircraft were examined to determine optimum
mission distances where the tilt rotor designs were superior
to a comparable contemporary (pure) helicopter. Two FORTRAN
computer programs (VASCOMP II and HESCOMP) developed under
contract for NASA Ames Research Center by the Boeing VERTOL
Company were used to predict aircraft performance. Program
results were validated using data from independent sources.
A simplified user's manual is included (with sample data and
program output) for VASCOMP I: use at the Naval Postgraduate
School, Monterey, California.
Accession ForNTIS GRA&IDTIC TARUnanuoo-nced
Just Ircation
By_.Di3tribution/
AviIlzbIlity Cedes
:Avail and/or
Dist Spocjlel3 J
TABLE OF CONTENTS
I. INTRODUCTION 8
II. BACKGROUND 0--------------------------------------10
III. PROBLEM DEFINITION ------------------------------- 17
IV. DESCRIPTION OF AIRCRAFT ------------------------- 19
A. GENERA -- ------------------------------------- 19
B. BOEING VERTOL MODEL 107 --------------------- 19
C. 44-PASSENGER TILT ROTOR --------------------- 42
r. 25-PASSENGER TILT ROTOR --------------------- 61
V. DESCRIPTION OF EXPERIMENTS ---------------------- 82
VI. RESULTS ----------------------------------------- 85
VII. DISCUSION AND CONCLUSIONS --- ;---------------------98
APPENDIX VASCOMP II USER'S MANUAL --------------------- 101
LIST OF REFERENCES ------------------------------------ 214
BIBLIOGRAPHY ------------------------------------------ 215
INITIAL DISTRIBUTION LIST ----------------------------- 217
4
LIST OF TABLES
1. BOEING VERTOL CH-46F DESCRIPTION --------------------- 20
2. COMPARISON OF ACTUAL CH-46F AND HESCOM.P CH-46F --------- 22
3. BELL HELICOPTER TEXTRON 44-PASSENGER TILT ROTOR ------- 42
4. COMPARISON OF BHT AND VASCOMP II 44-PAX TILT ROTOR --- 43
5. BELL/BOEING VERTOL 25-PASSENGER TILT ROTOR ------------ 61
6. COMPARISON OF BELL/BOEING & VASCOMP TILT ROTOR -------- 62
5
LIST OF FIGURES
1. Fuel Required vs. Range ------------------------------- 86
2. Time Required vs. Range ------------------------------- 87
3. Pax-Mi/Lb-Fuel vs. Range ------------------------------- 88
4. Loiter Time vs. Mission Radius ------------------------ 90
5. Fuel Required vs. Loiter Time ------------------------- 91
6. Hover Time vs. Mission Radius ------------------------ 92
7. Fuel Required vs. Hover Time -------------------------- 93
8. Fuel Required vs. Range ------------------------------- 95
9. Time Required vs. Range -------------------------------- 96
10. Pax-Mi/Lb-Fuel vs. Range ----------------------------- 97
6
ACKNOWLEDGEMENT
I wish to acknowledge the extensive assistance provided
by Mr. Dennis Mar of the W.R. Church Computer Center. His
technical expertise and his competernce in sharing his
knowledge in the areas of Multiple Virtual System procedures
and magnetic tape usage at NPS proved invaluable during the
conversion of VASCOMP II and HESCOMP from CDC FORTRAN to
VS FORTRAN.
Sincere thanks to Professor M. F. Platzer for his support
and guidance during the completion of the thesis research.
My appreciation is also extended to the many individuals
who were excescively cooperative in providing information and
material used for this research including: Mr. Tom Galloway,
Mr. Cliff McKeithan, and Ms. Donna Diebert, all uf NASA Ames
Research Center; Mr. Hal Rosenstein of Boeing VERTOL Company;
and Mr. Ron Reber of Bell Helicopter TEXTRON.
7
I. NTP0DUCTION
New horizons have been opened for V/STOL aircraft as a
result of the NASA/Army XV-15 Tilt Rotor Research Program and
its successful demonstration of the tilt rotor concept.
Application of tilt rotor technology lends itself aptly
to a myriad of military missions and also has significant
potential for future performance in civil roles. It is
plausible, for example, that tilt rotor aircraft could
replace conventional fixed-wing turboprops ani helicopters
for both military and civilian missions within optimal range
parameters. The limits of these "optimal" range missions can
be approximated with the aid of aircraft sizing and
"performance computer programs.
Past studies [Ref. 1, 2] used computer generated data
to make performance comparisons between tilt rotor aircraft
designs and comparable contemporary aircraft. The results
were then used to assist in evaluating the suitability of the
roles selected for the potential tilt rotor designs. This
"thesis research was conducted using a similar approach. Data
was generated for two potential tilt rotor designs using
the V/STOL Aircraft Sizing and Performance Computer Program
(VASCOMP II). Data was also generated for a contemporary
rotary-wing aircraft using the Helicopter Sizing and
Performance Computer Program (HESCOMP). Both programs were
8
developed under contract by the Boeing VERTOL Company for the
National Aeronautics and Space Administration, Ames Research
Center, Moffet Field, California.
VASCOMP II and HESCOMP are intentionally similar in
program structure, program data- requirements, and program
output. This allows comparisons between the results of the
two programs.
Application of VASCOMP II is appropriate for predicting
sizing and performance data for aircraft that employ fixed
wing surfaces to obtain lift in primary cruise flight. Use
of HESCOMP is applicable for aircraft that use rotary wing
surfaces to obtain lift in forward flijht.
9
II. BACKGROUND
A. NASA/ARMY PROGRAM
In 1972, NASA and the United States Army jointly
sponsored the XV-15 Tilt Rotor Research Aircraft Program.
The primary objective of the program was to demonstrate that
dynamic stability problems, which plagued earlier tilt rotor
designs, had been resolved. In 1973, Bell Helicopter TEXTRON
won a minimum cost contract to build two flightworthy te;tbed
tilt rotor research aircraft to be used during the flight
test portion of this proof-of-concept program. On October
22, 1976, the first cf the two prototypes (tail number N7d2)
was rolled out of Bell's Arlington, Texas facility. This
aircraft made its first hovering flight on May 3, 1977. It
was then used for extensive wind tunnel studies which had to
be completed, in accordance with the NASA/Army contract,
prior to the release of the second prototype (N703) for full
flight testing. The first full in-flight conversion fron the
helicopter mode to the airplane mode was performed in
aircraft N703 on July 24, 1979.
B. XV-15 DESIGN CHARACTERISTICS
The XV-15 is 42 feet long, has a H-tail, and utilizes a
slightly forward swept, hb4 h-winq that is 32 feet wide. The
design incorporates two three-bladed proprotors mounted on
10
wingtip nacelles, each rotor having a diameter of twenty-five
(25) feet. Maximum takeoff gross weight for the aircraft is
13,000 lbs in the VTOL mode and 15,000 lbs in the STOL mode.
Each wingtip nacelle house3 a transmission and a 1550-shp
Lycoming T-53 turboshaft engine that is modified (primarily
the lubrication systems) for both horizontal and vertical
operation. The nacelles, which can rotate through 95 degrees
during conversion between helicopter and airplane modes, are
positioned vertically for VTOL segments and rotated to the
horizontal position, after takeoff, for airplane operations.
The nacelles can be rotated five (5) degrees aft of vertical
for rapid decelerations in the air, aft translations at a
hover, and for providing a responsive means of acceleratino
and decelerating the aircraft during ground operations.
Conversion between the helicopter and airplane mode takes
roughly 12-15 seconds and has proven, without exception, to
be an uncomplicated, safe, and completely reliable procedure.
Rotor RPM is reduced when in the airplane mode to provide
greater propulsive efficiency. Maximum airspeed of the
aircraft is 301 knots in forward flight, 35 knots in sideward
flight, and 25 knots in rearward flight. As is typical with
conventional tandem rotor helicopters, the XV-15 is not
greatly affected by wind direction during hovering flight.
11
C. JOINT SERVICES OPERATIONAL TESTING
Military operational testing of the XV-15 began in 1982
to determine the suitability cf tilt rotor aircraft to
perform select military missions. Testing was conducted over
an 18-month period and included US Army evaluations at Fort
Huachuca, AZ to study the vulnerability of the aircraft in a
ground threat environment; US Navy shipboard and downwash
evaluations aboard the USS Tripoli and at Dallas NAS, TX
respectively; and JSMC mission-oriented flight tests at Yuma
Proving Grounds, AZ.
D. JVX
The versa'-ility of the XV-15 auickly convinced a study
group thiat a tilt rotor aircraft could meet future mission
recuirements of the Army, Navy, Marine Corps, t•nd Air Force.
By the end of 1982, the Joint Services Advancrd Vertical Lift
Aircreft revelopment Program (JVX) had been formed. Throuah
JVX, the services are attemptinc to procure approximately
1000 production model tilt rotor aircraft. In April IP?, a
contract was awarded to tha manufacturing team of Bell
Helicopter TEXTRON and Boeing VERTOL to begin preliminary
design work on the multimission, multiservice tilt rotor
aircraft.
E. BELL-BOEING DESIGN
As a baseline aircraft, preliminary desion studies are
using the Bel]-Boeina Model 0O1-X. The militarv Oesiqnation
12
for the world's first production model tilt rotor V/STOL
aircraft is the V-22 OSPREY. It will have a VTOL mission
weight of 43,800 pounds, a STOL mission weiaht of 55,000
pounds, a cruise speed at maximum gross weight of 260 knots,
and a service ceiling of over 30,000 feet. When loaded with
24 combat equipped troops, it will have a mission radius of
200 nautical miles and will be able to self-deploy worldwide.
Flight controls will incorporate triple-channel fly-by-wire
technoloQy and the propulsion system will utilize a 6000 shp
engine to be built by a yet-to-be-selected manufacturer.
Bell is responsible for the lift/propulpion system to include
wing, rotor, nacelle, and transmission. Boeing VERTOL is
responsible for everythinq below the winc to include the
all-composite fuselage and tail, the landing cear, and ail
subsystems. Boeing is also responsible for the aircraft's
aerodynamics, performanze, and hanalino cualities. Eioht
flying prototypes will be built with a "first fliaht" date
scheduled for wid-19S7. The first delivery, which will be to
the Mari:ne Corps, is currently , `ier~ued for mid!-]9l.
F. TECHNOLOGY TRANSFER
For nuoerous reasors, all substartiated durinq the flight
testing of the XV-15, tilt rotor aircraft could be rapidly
integrated into the civilian aviation community. Examples of
highly desirable features inherent to tilt rotor Oesigns
include:
13
1. High-speed cruise capability2. Fuel efficiency in the airplane mode3. Vertical taketff and landing capability"4. Low noise for passenger comfort/community acceptance5. Low vibration for passenger comfort/less maintenance
A tilt rotor design for use in the civilian sector can
have both national and international impact. It could be
particularly beneficial in applications where construction of
large airport facilities is either impractical or impossible.
For example, there is an abundant number of cases where small
communities are dispersed over vast areas supported by a poor
ground infrastructure. Alaska, Brazil, Indonesia, Canada,
Japan, and the Carribean Basin are prime example3 of areas
which desperately need a resource with the high productivity
potential of a tilt rotor aircraft.
Alask• dependence on aviation is significantlv 'hiqher
"than any other state in the nation. Statistics rRef. 51
show that Alaska has 16 times more aircraft and P times more
pilots logging 15 times more flight hours per capita as
comnared to the rest of the United States. Accessibility to
Alaska's natural resources (offshore oil, minerals, timber,
fish, etc.) could be greatly enhanced using V/SrOL aircraft
with the versatility found in tilt rotor designs. As this
state's economy grows, the need for additional conventional
airports could be greatly reduced or eliminated throuoh use
of V/STOL aircraft like the tilt rotor resultinc in enormous
savinqs in construction costs. Also, expenses associated
with the removal of snow and ice from Iona runways (75% of
14
[t •' ', ," "L " . " . • - " ' "•' L• " " . . . . . . . . . .
expenditures at community airports is allocated to this)
could be substantially reduced. Canada and Japan, in many
ways, have conditions similar to those found in Alaska.
Brazil, Indonesia, and the Carribean are examples of
developing nations that could greatly enhance their efforts
towards economic advancement through use of V/STOL aircraft
which would permit industrial and agricultural growth without
the necessity of having to build railroads, harbors, arid/or
airports to support expansion operations.
G. MISSION POTENTIAL
The VTOL capability of a tilt rotor aircraft coupled with
its capacity for high cruise speeds makes it a fierce
competitor for missions currently performed by conventional
helicopters. Some of the military and civilian applications
include:
1. Troop transport2. Search and rescue3. Reconnaissance/surveillance4. Law enforcement5. Medical evacuation6. Public transportation7. Corporate/executive transport8. Offshore oil exploration and production
H. EXCESSIVE INITIAL COST
Although tilt rotor aircraft have advantaaes which may
never be matched bv conventional helicoptc-rs, it is not
expected that rotary-wino aircraft will hecomp totally
obsolete. There are various parareters which, eependina on
15
their value, could make use of helicopters more feasible than
use of tilt rotor aircraft. One such parameter is cost. Any
new aircraft design necessitates a development proaram which
translates to high monetary expenditures. This fact could
make acquisition of tilt rotor aircraft "cost prohibitive" to
users requiring only a small auantity of aircraft, say, five
or less. The percentage of businesses in the civilian market
that would fit into this category is large enough to prevent
a civilian development program until after the completion of
a comparable military development proqram. The dancer of
postponing a civilian program is that, in doing so, the
United States may very well lose its competitive edge and
allow foreign competitors to seize the initiative and capture
the international market that is beginnina to form for tilt
rotor designs. The Soviet Union and France are currently
working on their own tilt rotor aircraft. F3r an example of
lost opportunities one has only to look as far as the Ouiet
Short-Haul Research Aircraft (OSRA) which underwent extensive
study and development at NASA Ames Research Center, Moffett
Field, California. This design was noted to have siqnificant
potential as a STOL transport hut the lack of a military
development program stalled the desian at the research
prototype stage. Recently, Japen announced the successful
maiden fliqht of a commercial STOL aircraft, soon to he made
available on the international market, which has an uncanny
resemblance to the NASA Ames OSRA.
16
III. PROBLEM DEFINITION
A. GENERAL
Perhaps the most difficult decision that will face a
potential user of tilt rotor technology is whether or not the
vastly increased productivity available through tilt rotor
aircraft de;Aqns justifies the substantially higher costs of
initial acquisit 4 on. Clearly, features that would qualify a
transport aircraft as "successful" include:
1. Low noise2. Long range3. Low vibration4. High performance5. Low operating costs6. Low fuel consumption
Additionally, parameters such as the cost of acquiring
real estate and the escalating costs of construction (for new
airports), aggravation of air traffic congestion at existing
airports, and the importance of time tc the traveller, will
play a major role in the decision making process of civilian
communities, businesses, and individuals who miaht be
considering the utilization of tilt rotor aircraft for their
their aviation transportation requirements.
Trade-off studies provide a low risk, relatively low cost
method of analyzing available options during any selection
process. In the case of judging the suitability of a new
aircraft design, computer programs, used for predicting
17
sizing and performance parameters, have become a vital tool
used in the early stages if design work.
B. RESEARCH GOAL
The objective of this research was to make comparisons
between the performance of a Boeing VERTOL Model 107 tandem
rotor helicopter (military designation: CH-46F SEA KNIGHT)
and the performance of compafable tilt rotor aircraft.
i#A Specificaliy, it was desired to use sizino and performance
computer programs to predict values of "rance" which would
show the superiority of one aircraft type over the other.
C. RESEARCH PARAMETERS
Parameters considered during this research included:
1. Fuel required versus distance2. Time recuired versus distance3. Passenger mile per pound of fuel versus d4stance
18
IV. DESCRIPTION OF AIRCRAFT
A. GENERAL
The aircraft studied during this thesis research project
included the following:
1. Boeing VERTOL Model 1072. 44-Passenger Tilt Rotor3. 25-Passenger Tilt Rotor
All three aircraft are described in this chapter and the
results of calibration runs ar6 discussed. It should be
noted that these particular aircraft were not selected for
reasons related to their size or performance capabilities.
Selection was based solely on the fact that these were the
aircraft for which the greatest quantity of descriptive
information was obtained during the period of research.
B. BOEING VERTOL MODEL 107
1. Description
There are numerous versions of this Boeing VERTOL
product that first flew in 1958. The Model 107 1I is the
standerd commercial version equipped with two 1,250 shp
General Electric CT58 turboshaft engines. The military
version of this aircraft is the CH/UH-46 SEA KNIGHT. There
are several variations to the basic CH-46 to include the
CH-46A, CH-46D, and the CH-46F. Typical differences between
versions include uprated power plants, additional electronic
19
equipment, and cambered rotor blades. The CH-46F was used
for all experiments discussed in this thesis. Reference 6
was used to obtain some of the more noteworthy specifications
of the aircraft as shown below in Table 1.
TABLE'1
BOEING VERTOL C•!-46F DESCRIPTION
Type: Twin-engined, transport helicopterEngines: Two 1400 shp General Electric T58 shaft-turbinesRotors: Two three-bladed rotors in tandem
• Dimensions:Diameter of main rotors: 51 ft 0 in
"Length overall, blades turning: 84 ft 4 inLength of fuselage: 44 ft 10 inMain rotor blade area: 39.85 sq ftMain rotor disc area: 4,086 sq ft
Weights and Loadings:Weight empty, equipped: 13,342 lbMax taeoff and landing weiat;t: 23,000 lbMax disc loading: 5.63 psf
Performance:Max permissable speed: 144 knotsMax cruising speed: 143 knotsService ceiling: 14,000 ft
Ranges (with 10% reserve fuel):At 20,800 lb (4,275 lb payload): 206 naut miAt 23,000 lb (6,475 lb payload): 198 naut mi
Fuel Capacity:Standard configuration: 380 US gal
Accommodation:Crew; 3Passengers: 25
2. Aircraft Calibration Using HESCOMP
Data for the CH-46F was cbtained from NASA Ames
Research Center. To insure the validity of results obtained
using HESCOMP, calibration runs were made to match computer
generated results with descriptive data from Table 1 above.
There were several key parameters to be considered during the
22'
experiments including: fuel, time, power, distance, and
airspeed. As such, emphasis was placed on calibrating the
items that would affect these key parameters. For example,
to match the aircraft dercription in Table 1, the HESCOMP
results had to depict an aircraft that could carry a payload
of 4275 pounds for a distance of 206 nautical miles and, in
so doinc, use all available fuel except for a 10% reserve.
Calibration was accomplished using the flexibility built into
HESCOMP. Many data locations are established to describe a
component's weight or performance, as applicable. Some )f
the data locations represent constants and some represent
multiplicative factors. Data can be input based on actual
values or the user can use the program, in some cases, to
calculate approximate values. The values of component
weights, for example, can be input based on known data or the
program can calculate them based on weight trends of other
aircraft in the same weight class. For the CH-46F, when the
empty weight did not match data from Reference 6, the "Body
Group Weight Factor" (location 2622), which varies the
fuselage weight, was manipulated until the desired empty
weight was obtained. After consideration of the parameters
being analyzed, it was not felt that this "fudge factor"
would have any impact on the results. This flexibility
allows for convenient and very precise control of many of the
parameters in the program. As can be noted below in Table 2,
the calibrated aircraft description used for HESCOMP very
21
nearly matches the performance specifications of the actual
aircraft as described in reference 6.
TABLE 2
COMPARISON OF ACTUAL CH-46F AND HESCOMP CH-46F
ACTUAL HESCOMP % DIFFERENCE
Dimensions:Main rotor diameter: 51.000 ft 50.966 ft 0.067Length overall: 84.333 ft 83.900 ft 0.513Fuselage length: 44.833 ft 43.700 ft 2.528
Weights and Loading:Weight empty: 13,342 lb 13,342 lb 0.000Operating weight: 14,055 lb 14,055 lb 0.000Payload: 6,475 lb 6,475 lb 0.000Fuel: 2,550 lb 2,550 lb 0.000Gross weight: 23,000 lb 23,000 lb 0.000"Disc loading: 5.629 psf 5.637 psf 0.134
Ranges (normal power):At 20,800 Ib: 206 nnm 206 nm 0.000At 23,00b lb: 198 nm 198 nm 0.000
3. Program Data
The program output that was generated by HESCOMP for
the CH-46F calibration run is shown below on pages 23 - 41.
Pages 23 - 26 show the "echo" of the input data file.
Mission performance data begins on page 34. It can be seen
that two missions were flown. The first mission was flown at
the aircraft's maximum gross weight of 23,000 pounds. This
was done to insure that the aircraft was sized properly by
the program. The payload was reduced for the second mission
which was flown at a gross weight of 20,800 pounds. The
aircraft's maximum range (at normal power setting) was
calibrated for this weight.
22
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C. 44-PASSENGER TILT ROTOR
Description
The first of the two tilt rotor designs used in this
study is a civilian derivative of the V-22 OSPREY. This
aircraft design was the subject of a study [Ref. 7) conducted
by Bell Helicopter TEXTRON (BHT) wherein the 44-passenger
tilt rotor was shown to be substantially more cost-effective
(despite a higher acquisition cost) than the 44-passenger
Boeing VERTOL 234LR. This tilt rotor design features two
General Electric T64-717 engines that each produce 4855 shp.
Reference 7 was used to compile the information in Table 3.
TABLE 3
BELL HELICOPTER TEXTRON 44-PASSENGER TILT ROTOR
Type: Twin-engined, commercial transport aircraftEngines: Two 4855 shp General Electric T64 shaft-turbinesRotors: Two three-bladed rotors on wingtip nacellesDimensions:
Diameter of main rotors: 38 ft 0 inLength overall: 60 ft 11 inLength of fuselage: 60 ft 11 inWing span: 47 ft 10 inMain rotor disc area: 2,268 sa ft
Weights and Loadings:Weight empty, equipped: 26,676 lbMax takeoff and landino weight: 44,000 lbMax disc loading: 19.4 psf
PerformanceMax permissable speed: 360 knotsMax cruising speed: 300 knotsService ceiling: 34,000 ft
Ranges (with reserve fuel)At 44,000 lb (9124 lb payload): 725 naut ml
Fuel Capacity:Standard configuration: 1043 US gal
Accommodation:Crew: 4Passengers: 44
42
2. Aircraft Calibration Using VASCOMP II
Basic data for the 44-passenger tilt rotor was also
obtained from NASA Ames Research Center. The similarity
between VASCOMP II and HESCOMP permitted using an identical
calibration technique as that described in paragraph IV B 2
above. Table 4 shows the comparison3 between the aircraft as
described in Ref. 7 and as portrayed through the VASCOMP II
output results.
TABLE 4
COMPARISON OF BHT AND VASCOMP II 44-PAX TILT ROTOR
BHT VASCOMP % DIFFERENCE
Dimensions:Main rotor diameter: 38.000 ft 38.000 ft 0.000Length overall: 60.917 ft 60.900 ft 0.028
Weights and Loading:Weight empty: 26,676 lb 26,676 lb 0.000Operating weight: 27,876 lb 27,876 lb 0.000Payload: 9,124 lb 9,124 lb 0.000Fuel: 7,000 lb 7,000 lb 0.000Gross weight 44,000 lb 44,000 lb 0.000
Ranges (normal power):At 43,676 lb: 725 nm 725 nm 0.000
3. Program Data
VASCOMP II output from the calibration run is shown
on pages 44 - 60. It should be noted that the output closely
parallels the format and sequence of output generated using
HESCOMP. Due to the limited information availalle, only one
mission was programmed. A maximum takeoff qross weight o'
44,000 pounds was used for the mission to calihrate the
aircraft's maximum range.
43
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D. 25-PASSENGER TILT ROTOR
1. Description
This tilt rotor design is, in actuality, the V-22
OSPREY that was discussed in Chapter IT - BACKGROUND. It is
describec in reference 6 and reference 7. Both references
werb !sed to obtain the specifications listed in Table 5.
TABLE 5
BELL/BOEING VERTOL 25-PASSENGER TILT ROTOR
Type: Twin-engined, military transport aircraftEngines: Two 4855 shp General Electric T64 shaft-turbinesRotors: Two three-bladed rotors on wingtip nacellesDimensions:
Diameter of main rotors: 38 ft 0 inLength overall: 56 ft 10 inLength of fuselaqe: 56 ft 10 inWing span: 46 ft 6 inMain rotor disc area: 2,268 so ft
Weights and Loadings:Weight empty, equipped: 26,858 lbMax takeoff and landing weight: 43,800 lbMax disc loading: 19.300 psf
Performance:Max permissable speed: 360 knotsMax cruising speed: 300 knotsService ceiling: 34,000 ft
Ranges (with reserve fuel):At 43,800 lb (10,000 lb payload): 400 naut miAt 35,400 lb ( 1,600 lb payload): q20 naut mi
Fuel Capacity:Standard configuration: 1043 US qal
Accommodation:Crew: 4Passenaers: 25
2. Aircraft Calibration Usino VASCOMP II
Basic data for Qeneral dimensions, aerodynamics, and
engine performance is identical to that used for the
44-passencer tilt rotor aircraft. Table 6 shows the values
61
for aircraft specifications ao obtained from references 6 and
7 and as obtained using VASCOMP II.
TABLE 6
COMPARISON OF BELL/BOEING AND VASCOMP 25-PAX TILT ROTOR
BELL/BOEING VASCOP. % DIFFERENCE
Dimensions:Main rotor diameter: 38.000 ft 38.000 ft 0n0o0Lenqth overall: 56.833 ft 56.600 ft 0.410
Weights and Loading:Weight empty: 26,858 lb 26,858 lb 0.000Operating weight: NOT SHOWN 29,268 lbPayload: 10,000 lb 10,000 lb 0.000Fuel: 7,000 lb 7,000 lb 0.000Gross weight: 43,800 lb 43,800 lb 0.000
Ranges (normal power):At 43,800 lb: 400 nm 400+ nm 0.000At 34,600 Ib: 920 nm 920 nm 0.000
3. Program Data
The following pages are a reproduction of the input
data used for VASCOMP and the output of the program for the
calibration runs. Two missions were flown. The first
mission was flown at a gross weiqht of 43,800 pounds to
properly size the aircraft. This mission was based on the
US Marine Corps requirement for the V-22, in the medium
assault transport role, to have the capability to carry 2-;
troops plus equipment on a mission radius of 200 nautical
miles. The second mission was representative of the US Navy
requirement for the V-22, in the combat search and rescue
role, to be able to fly a 460 nautical mile radius to rescue
four people.
62
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V. DESCRIPTION OF EXPERIMENTS
" A. EXPERIMENT 1 - MAXIMUM RANGE
In this experiment, the 25-passenger Model 107 and the
25-passenger tilt rotor we-e used to examine aircraft
performance for a single-leg mission. The aircraft were
programmed to carry identical payloads of 5000 pounds (which
represented a load of twenty-five 200 pound passengers,
including baggage) and to fly identical mission profiles at
an altitude of 5000 feet mean sea level (MSL). The standard
flight profile used for any given "mission" followed a
se":jence which included:
* Taxi* Hover/7akeoff* Climb
* Cruise,* !escent
La* Lnding/'-'o, er* Tarl
The objective was to determine the mission raroe, w~t'-Kn
"the maxi'um ranae cardabili tv of the CH-46F, where t~'e
performance of the tilt rertor surr;,sse4 the perform4nc- '_:f
the tandem rotor he1licopter.
P. 7XPEPIMENT 2 - LOJ',ER FN.rU?ANCE
SThe second experiment involved comparisons !'Ptween the
performance capý bilites of , 25-passenaer tilt rotor
* aircraft and the 25-n3ssenaer CP-46F for missions reouirina
P2
I4
the aircraft to traverse a specified range (mission radius)
and conduce a loiter mission for the maximum time permissible
while maintaining sufficient fuel to return the same distance
as the outbound leg (plus reserves). Payloads carried by the
aircraft were 3500 pounds and the cruise segments were flown
&-. an altitude of 5000 feet MSL. Each "mission" consisted of
the following segment seqtuence:
*Taxi
* Hover/Takeoff* Climb* Cruise* Loiter* Cruise
*• * Descent* Landing/Hover*Taxi
The objective was to compare the capabilities of each
type aircraft at given tissi ,v radius values.
C. EXPERIMENT 3 - HOVER ENDURANCE
This experiment is an extension of experiment 12 in that
payload, altitude parameters, and aircraft t,¢es 7emained
unchanged. Flight profiles were similar except that the
Iciter segn.ent was replaced with a descent-hover-climb series
0of segments. From a performance standpoint, tle difference
* . lien in the fact that the hover fliaht mode As a more
demanding flight mode in terms of fuel (and powerl reauireO.
The objective in conducting this experiment was to
aanalyze the differences in aircraft canabi]it'es for civet!
values of mission radius.
SP3
D. EXPERIMENT 4 - ONE TILT ROTOR VS. TWO HELICOPTERS
In the fourth and final experiment, the capability of one
44-passenger tilt rotor was compared to the capability of two
25-passenger Boeing CH-46F helicopters.
In this mission scenario, the two tandem rotor aircraft,
with no payload onboard, simultaneously depart from a site
designated "Helipad A". Each helicopter flies a distance
equal to one-half of its maximum range capability (103 NIM),
receives a payload increase of 4400 pounds (to simulate the
boardina of 22 passeners, each weight a total of 200 pounds
with baggage), and returns to the oricinal departure sit-.
One helicopter flies to "Helipad F" while the other travels
to "Helipae C". Helipads B and C are separated4 v a distance
designated "Iea BC".
Starting from Helipad A with no payload, the ,'4-passenaer
tilt rotor departs at the same time as the two helicoDters,
flies to Helipad B, picks up a 4400 pound payload, then flies
to Helipad C and picks up an additional 4400 nounds. The
tilt rotor complets its mission by returnina to Helipad 4.
The standard fliaht profile used for this experiment was the
same as that used for experiment *I.
"The objective for this experiment was to fin' the valups
for Lea BC where the tilt rotor coule transnort t)e navloads
r further distances more efficientlv thar t-e two
ihelicopters.
P41
VI. RESULTS
A. EXPERIMENT #1
1. Fig 1, Pg 86 - Fuel Required vs Range
In addition to the obvious advantage of the extended
range capability (on a single fuel load) that the tilt rotor
maintains over the tandem rotor helicopter, the graph on
page 86 shows that for ranges of approximately 150 NM and
beyond, the tilt rotor requires less fuel than the helicopter
to complete the mission. It is noted that to traverse the
distance covered by the tilt rotor in one fuel load the
CH-46F would have to refuel three times.
2. Fig 2, Pg 87 - Time Required vs Range
For the data presented, within the single fuel load
range of the tandem rotor helicopter, it can be seen that for
any given range beyond approximately 75 NM the helicopter
requires at least tice as much time to travel thle same
distance as the tilt rotor. Refueling time woull further
compound this disadvantace for the helicopter.
3. Fig 3, Pg 88 - Passenaer-Mile per Lb of Fuel vs Rance
This graph shows that at a ranqe of just under 150 NY
the tilt rotor will perform more efficiently '-v beina a'lIe to
complete more passenger miles per pound of expended fuel Cvhan
the CH-46F.
85
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B. EXPERIMENT #2
1. Fig 4 Pg 90 - Loiter Time vs Mission Radius
The tilt rotor clearly holds a significant loiter
endurance advantage over the tandem rotor helicopver. At the
maximum mission radius for the CH-46F, which allows no loiter
time, the tilt rotor can travel the same distance and still
conduct a three hour (plus) loiter mission.
2. Fig 5, Pg 91 - Fuel Required vs Loiter Time
The graph reveals the tilt rotor's disadvantace of
requiring roughly 50' more fuel than the helicopter for the
same amount of loiter time.
C. EXPERIMENT *3
1. Fig 6, Pa 02 - Pover Time vs Mission Radius
Compared to the loiter mission, the tilt rotor does
not maintain as large d marain of superiority for the hover
mission. At the maximum mission radius for the CP-46r, (nc
hover time capability) the tilt rotor *-as suffici-rt
remainina fuel for a little over one hour cf hover tire.
2. Fia 7, Pg 93 - Fuel Pecuired vs Hover Tire
The hoverina efficiency of rotary wino aircraft
made apparert bv this arath whicl- silowq t1'at tl t: It r-trr,
for a aiven hover time reauires roua~hlv 2.7 times mere ftusl
than the tandem rotor 1eliconter.
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D. EXPERIMENT #4
1. Fig 8, Pg 95 - Fuel Required vs Range
This graph shows that one 44-passenger tilt rotor can
do the job of two 25-passenger helicopters (operating at 88%
seating capacity) usinq less fuel even when travelling to
sites that are separated by distances of up to 250 NM. It is
evident that even if the helicopters were operating at full
capacity, the tilt rotor's advantage would not decrease
substantially.
2. Fig 9, Pg 96 - Time Reauired vs Range
The 44-passenger tilt rotor, when travelling to
locations separated by distances of up to nearly 300 NM,
performs better than two CH-46F helicopters that are flvina a
103 NM radius to the same locations.
3. Fiq 10, Pq 97 - Passenger-Mile per Lb of Fuel vs Rnq
This final graph shows that using one 44-passencer
tilt rotor travellinq to two landinq sites separatee by 150
IM or more is more efficient th~an usina twco• tandem rotor
CH-46F helicopters to cover the same two landina sites.
94
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97-
VII. DISCUSSION AND CONCLUSIONS
A. EXPERIMENTS
"The experimental results substantiate that the V-22,
designated to replace the CH-46F among other aircraft, will
offer significant improvements in speed, loiter endurance,
hover endurance, performance, and efficiency. The results
lend further credence to the manufacturer's claims that
tilt rotor aircraft can transport more passencers/payload
over longer distances in less time than conventional
helicopters while retaining the important advantaae of
vertical takeoff and landing. It is acknowledged that the
tilt rotor can readily perform transport missions using less
fuel than helicopters but if a large percentage of hoverinq
flight is required, conventional rotary wina aircraft are far
more efficient from a fuel consumption standpoint. Fowever,
they do not have the staying power that the tilt rotor
demonstrates. ! -
B. TILT ROTOR ADVANTAGES TO THE MILITARY
The speed and efficiency of the V-22 will allow oreater
stand-off ranges for naval assault fleets; permit more rapid
buildup of assaLoit forces at objectives while retaininq the
ability to operate from small ships and/or to maintain an
independence from runways.
9R
C. TILT ROTOR ADVANTAGES TO THE CIVILIAN AVIATION COMMUNITY
The high productivity of the tilt rotor will provide much
needed relief in air traffic congestion at major airports by
providing a vehicle that has twice the speed of conventional
helicopters without sacrificing the ability to transport
passengers on a city-center to city-center basis. This will,
effectively, reduce the recuirement to increase the number of
available commuter jets and turboprops as transportation
demands increase.
D. COMPUTER PROGRAMS
Both VASCOMP II and HESCOMP were found to be relatively
straightforward in their apnlicatior. Each proaram has an
enviable range of versatility and flexibility. The most
difficult factor in workinq with the proqrams is, without a
doubt, obtaining the input data. Without the cooperation of
an agency. such as NASA Ames Tesearc!. Center it wou)O be
extremely difficult to compile a complete data set for any
given aircraft. Aircraft manufacturers are a potential
source for data. The aircraft for which it is easiest to
obtain data are those produced for the military. As rart of
the nrocurement process, manufacturers are recuired to suhrit
a MIL-STD Form 1374 entitled GROUP WEIGHT STATEMENT wl-c"
provides detailed weight data as well as dimensional ane
structural data. This document is cenerallv availale for
official uses.
09
Concerning the accuracy of the programs in the faithful
representation of the aircraft described bv the input eata,
it is felt that the trends routines built into the program do
an adequate job of approximating the vehicle and the
flexibility of the program more than adequately allows for
adjusting the output to obtain recults which match actual
flight test dta.
100
APPENDIX
VASCOMP USER'S MANUAL
This User's Manual was completed in conjunction with
thesis research. It was prepared using the guidelines that
the manual should:
1. Provide helpful information for those individualsinterested in using VASCOMP II at the NavalPostgraduate School (NPS), Monterey, California.
2. Provide information in such a way that a user could runVASCOMP II without relying on material elsewhere inthis thesis.
3. Simplify, to a large extent, the'material present-1 in
the Boeing VERTOL Company VASCOMP II User's M4anual.
4. Include examples of:
a. Required Job Control Language (JCL) statements forusing VASCOMP II on the IBM 3033 computer at NPS.
b. Sample data for a V/STCL aircraft.
c. Sample output for a V/STOL aircraft.
* 10l-
I. INTRODUCTION
A. PURPOSE
This manual was designed to provide the user with a
simplified version of the Boeing VERTOL Company's VASCOMP II
User's Manual. A copy of the Boeing VERTOL user's manual is
catalogued and available at the Dudley Knox Library at NPS.
Also included in this manual are examples of:
1. Required Job Control Lanugage (JCL) statements forusing VASCOMP II on the IBM 3033 computer at NPS.
2. Sample data for a V/STOL aircraft.
3. Sample output for a V/STOL aircraft.
B. APPLICABILITY
VASCOMP II, the V/STOL Aircraft Sizing and Performance
Computer Program, is a viable computer program for predicting
size and performance data for V/STOL aircraft. It can be
used for any aircraft which employs fixed wing lift during
primary cruise flight.
C. PROGRAM OPERATION AT NPS
Due to the large size of VASCOMP II, (in excess of 16,000
lines of FORTRAN code), production runs at NPS should be
accomplished using the batch operating system on tý;e IBM 3033
Network referred to as MVS (Multiple Virtual System).
Chapter II of *his manual provides an explanation and an
102
example of the Job Control Language statements required for
running VASCOMP II on the batch system. Chapter II also
describes the procedure for running the program on VM/CM3.
D. PROG3RAM DATA
Production runs of VASCOMP II require a significant
quantity of data. Chapter III provides information on the
format of the data file and a list/description of the data
locationa.
E. V/STOL AIRCRAFT EXAMPLE
Chapter IV consists of output genereted by running the
program with the data shown in Chapter III.
I1
103
IlI PROGRAM OPERATION AT NPS
A. The VASCOMP II program can be run at NPS by one of two
means. The user is highly encouraged to utilize the MVS
batch system due to the excessive size of the program. This
is primarily for the user's convenience since, in order to
run the program using VM/CMS, the program would have to be
filed on an A-disk and would require 37% of the normal eight
cylinder allocation for storage alone. Using only eight
cylinders, it is not possible to compile the program. If it
is necessary to use VM to run the program, the user can
request allocation of a B-disk or, as an alternative, it is
possible to create a temporary disk cf variable size any time
the user logs onto VM/CMS. NOTE: Files on a temporary disk
fre tr',iy temporary. FILES ARE LOST IF THE USER LOGS OFF OR
IF THE SYSTEM GOES DOWN UNEXPECTANTLY. A user can create a
temporary disk using the following steps:
1. Create an EXEC File
In CMS type: XEDIT TDISK EXEC
2. Temporary Disk EXEC File Commands
Type the following lines into the file:
& TRACECP DEFINE T3350 200 CYL 24& STACK YES& STACK TDISKFORMAT 200 CACCESS 191 B
104
3. Create a Second EXEC File
In CMS type: XEDIT MODES EXEC
4. Modes EXEC File Commands
Type the following lines into the file:
ACCESS 200 A
5. Activate the EXEC Files
In CMS type: T"DISK. Wait for the "RT" response thentype: MODES.
6. File Alignrunt
The user now has a 24 cylinder "temporary" A-disk and
an 8 cylinder B-disk. All default files will go to the
temporary A-disk. Move files for permanent storage on the
191 disk by typing the following command next to the file in
FLIST: COPY / = = B
B. PROGRAM OPERATION USING VM/CMS
Use VS FORTRAN if the program is to be run on VM.
C. PROGRAM O:?ERATION USING VMS BATCH
The preferred mode of operation is to the MVS batch. To
run VASCOMP II on MVS batch at the Naval Postgraduate School,
the user must access the "AERO DISK" and locate the file
entitled: VASCOMP FORTRAN Al. Browse the file for furt'ier
information 3n program operation.
D. JOB CONT.1ROL LANGUAGE (JCL) STATEMENTS FOR DATA FILE
When the user has compile the necessary data to run the
program, create a FORTRAN file by following the steps below.
105
1. Create the File
In CMS type: XEDIT (filename) FORTRAN
2. Job, Main, Procedure, and DD Statements
Place the following Job Control Language (JCL)
statements at the top of the file:
//jobname JOB (nnnn,9999),'ident°,CLASS=C//*MAIN ORG=NPGVMI.nnnnP
?/cO EXEC PGM=VASII//STEPLIB DD DSN=MSS.Fxxxx.VASCOMP,DISP=SHR// DD DSN=SYSl.pp.VFORTLIB,DISP=SHR//FT06F001 DD SYSOUT=A//FT07F001 DD SYSOUT=A, DCB=(RECFM=FBA,LRECL=133,// BLKSIZE=1300)//FT05F001 DD *
where,
• The jobname may contain eight alphanumeric characters.The first character must be alphabetic.
* nnnn is the user number assigned by the computer center.
• ident may contain twenty characters (including blanks).
* Correct spelling and spacing on the JOB and MAINstatement (first and second line, respectively)is critical.
* To send the program output to the remote printer inHalligan Hall, modify the MAIN statement by replacingNPGVMI.nnnnP with NPGVM!.HAL
3. Program Data
The data for the program is placed immediate!y
after the last JCL statement (see example below).
4. Delimiter and Null Statements
Place the following JCL statements immediately
after the last line o.' the program data:
/10*//
106
4.
The delimiter (/*) statement is an end of fiue statement for
marking the end of a data set. The null (//) statement is
used to mark the end of the job. Without the null statement
the user may find that the prograr will not run with any
degree of consistency.
B. DATA FILE EXAMPLE
A complete data file, including proper JCL statements,
for a program calibration run using an 8-passenger tilt rotor
aircraft is included in Chapter III of this manual. The
following is an example of the proper placement of the JCL
statements.
//VASWAL01 JOB (1053,9999),'TOM WALSH SMC 2986',CLASS=C//*MAIN ORG=NPGVM1.1053P//GO EXEC PGM=VASII//STEPLIB DD DSN=MSS.Fxxxx.VASCOMP,DISP=SHR// DD DSN=SYS1.PP.VFORTLIB,DISP=SHR//FT06F001 DD SYSOUT=A//FT07F001 DD SYSOUT=A, DCB=(RECFM=FBA, LRECL=133,// BLKSISE=I300)//FT05F001 DD *
(data)
/.
107
I-. - •. .'.r . W
III. DATA LOCATIONS AND DESCRIPTIONS
A. INFORMATION SOURCE ACKNOWLEDGEMENT
It should be noted that a significant portion of the
information in this chapter is either paraphrased or taken
directly from the Boeing VERTOL Company VASCOMP II Users's
Manual. In particula', most of the variable descriptive
information is taken from Chapter 5 of the Boeina Manual.
B. ORGANIZATION OF DATA
For efficiency and program logic, each data value is
identified by a four-digit integer. For convenience, the
data is read in uslng rows of t; tu five (5) data values.
The format of the FORTRAN "READ" statement is:
FORMAT( 14, lx, 11, 5'E14.7) )
When preparing the data file, spacing must matcn the above
format precisely. Failure to accomplish this will result in
erroneous data vAlues. The correct seouencinq of values for
any given line of data is:
COLUMNS VARIABLE TYPE VATA DESCRIPTION
01 - 04 INTEGER First value ide,.tification5 N/A Blank6 INTEGER Number of vaiues it: row7 N/A Blank
08 - 21 R E A L Value identified in COL 4-22 - 35 R E A L ' Talue in COL I - 4 clus e36 - 49 R E A L Value in COL I - 4 plus I50 - 63 R E A L Value in COL I - 4 plus t.1ret64 - 77 R E A L Value in CO. 1 - 4 plus four77 - 80 N/A Blank
I08i
1234567890123456789012345678901234567890123456789012345678901
EXAMPLE:
L. 0227 4 0.123 72.92 3.0 0.25
In this example, the data line contains four data values (the
numbers at the top of the page are for the convenience of the
*••- manual user to verify the column locations of the dataW. The
first data value, number 0227, is 0.123. The second data
value, number 0228, is 72.92. The third value, number 0229,
is 3.0. The fourth data value, number 0230, is 0.25.
C. COMMENT, END-OF-CASE-STUDY, AND EXIT LINES
1. Two types of comment lines e-dn be included in the
program data.
a. The first type of comment is one which the user
desires to have repeated at the top of each page of proaram
output. To accomplish this, place a "7" in column 1 thru 8
on the first line immediately following the JCL statements,
On the next line, columns 1 thru 6 must be left blank.
Columns 7 thru 80 are then used for any comment (such as type
of aircraft being studied, etc.) .-o be repeated at the top of
each and every page of proctram output.
b. The second cype of comment is one which the user
desir.s for separating qroups of e',ta. This tvpe of comment
card is optional. It is accomplished in the same manrer as
the first type (by placing a "7" in columns 1 thru 8, etc.)
but the comment information will be printed in the output
109
only once. Examples of both types of comment lines are
included in the sample data file at the end of this chapter.
5. If the user desires to analyze more than one type
of aircraft using only one data file, the data groups can be
3eparated by placing an "8" in columns 1 thru 8. A "7" card
and a repeating comment line should immediately follow the
"8" card.
d. To properly exit the proqram, the last data line
must contain a "9" in columns 1 thru 8.
D. DATA CATEGORIES
"The Roeing VERTOL Company VASCOMP II User's Manual
includem a total of twenty-four different data input sheets.
Data can be loosely grouped into six categories:
1. Aircraft General Information2. Aircraft Dimensional Information3. Mission Profile Information4. Engine Cycle Information5. Propeller Data6. Supplementary Information
The following paragraphs describe the purpose of each data
loration
E. AIRCRAFT GENERAL INFORMATION
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0001 Option Indicator (OPTIND)
0 = Program calculates aircraft arcss weight,dimensions, and power required.
1 = Program calculates aircraft cross weiaht,dimensions, and power required to completea user specified mission flight profile.
110
2 = Input gross weight and mission profile.Aircraft size remains fixed and programcalculates time history performance dataand fuel required to complete the mission.
3 = Input operating-weight-empty and missionprofile. Aircraft size remains fixed an,program calculates takeoff gross weightand fuel required to complete the mission.
For a combination of options 1 and 2, input data to describethe aircraft and mission flight profile at the top of the
data file. This data should then be immediately followed by
the additional missions for off-design-point performance
calculations.
0002 Print Indicator (TNIRPK)
0 = Mission performance data output for thefollowing flight segments will consist of:
All - time, range, fuel used, aircraftweight, press. alt., true airspeed, eng.turbine temp., engi code specifying eng.performance, and the primary eng. thrustor horsepower fraction (PETF or PEHF).
Takeoff, Hover, and Landinq - Thrust toweight ratio, propeller power coeff.,propeller thrust coeff., prop tip speed.
Taxi/Takeoff, Hover/Landing - Lift eng.thrust fraction (LETF).
Climb/Cruise/Descent/Loiter - Mac. no.,equivalent airspeed.
Climb, Cruise, and Descent - Mach numberfor drag diverqence.
Climb and Descent - Flight path angle,and fuselage attitude angle.
Climb - Rate of climb.
"Cruise - Specific ranqe.
11
Descent - Rate of sink.
Loiter - Time rate of fuel consumption.
Takeoff, Hover, Landing, Cruise, andLoiter - Propeller efficiency.
1 = All data output for TNIRPK = 0 is printed.The following information is also printed:
Climb/Cruise/Descent/Loiter - Fuel flowrate, lift, drag, horsepower, thrust,lift coeff., drag coeff., propellerpower coefficient, prop advance ratio,propeller thrust coefficient, propellertip speed, and propeller efficiency.
2 = Delete mission performance data output.
0003 Drag Indicator (DRGIND)
0 = Program calculates draa rise due tocompressibility effects.
1 = If drag rise characteristics of the acftare known, a 3-D table of compressibilitydrag coeff. can be input as a function ofMach number and lift coefficients.
2 = Used for supercritical drag liveraence.
0004 Oswald's Efficiency Factor Indicator (OSWIND)
0 = The user inputs a fixed value for Oswald's(spanwise loadina) efficiency factor.
1 = Proqram calculates Oswald's efficiencvfactor as a function of wina aspect ratio.
0005 Propeller Dimension Indicator (PDMIND)
- Input dia. and activity factor per blade.
2 = Input disc loading and activity factor perblade.
3 = input diameter and thrust coefficient tosolidity ratio.
4 = Input propeller disc loadina and thrustcoefficient to solidity ratio.
112
0006 Fuselage Dimension Indicator (FDMIND)
0 = Input fuselage length and wetted area.
1 = Input cabin length (constant diameter),nose and tail fineness ratios. Proaramcalculates acft length and wetted area.
2 = Input desired capacity, seat width andpitch, number and width of aisles, numberof seats abreast for tourist and firstclass, galley and lavatory size. Programcalculates fuselage size.
0007 Wing Dimension Indicator (WDMIND)
0 - Input wing loading and aspect ratio
1 = Input chord to diameter ratio and eiscloading. The size trends subroutine thencalculates the wing loadina.
2 = Input wing loading and disc loading. Thesize trends subroutine then calculates thechord to diameter ratio and aspect ratio.
3 = X-Wing configuration
** * D1 NOT SET WDMIND = 1 or 2 IF ENGIND = I 1
0008 Horizontal Tail Indicator (HTIND)
0 = Program computes H-tail volume coefficient
1 = Input horiz. tail volume coeff. and momentarm. Horiz. tail sfc is then calculatedby the size trends subroutine.
2 = Input the horizontal tail area as a fixedsize surface.
0009 Vertical Tail Indicator (VTINDN
0 = Program computes vert. tail volume coeff.
= Input verf. tail volume coeff. and momentarm. The size trends subroutine thencalculates the vertical tail surface area.
2 = Input the vertical tail area as z fixedsize surface.
113
0010 Engine Size Indicator (FIXIND)
0 = Input level of maximum power or thrust.Engine size is fixed.
1 = The engine size is "rubberized" and theengine sizina subroutine calculates thelevel of maximum power or thrust.
0011 Engine Indicator (ENGIND)
0 = Turboshaft engine cycle.
1 = Turbofan or turbojet enaine cycle.
2 = Turbofan engine cycle. Program simulatesoperation of a convertible encine cycle.
0012 Engine Sizing Indicator (ESZIND)
0 = Engines sized for takeoff conditions only.
1 = Engines sized for takeoff or cruiseconditions whichever requ'Les more power.
(No input if LFTIND = 0 or 1)
0013 Lift Enoine Indicator (LFTIND)
0 = No separate lift propulsion encine
1 = Propulsion svs. includes a primary enqinecycle (cruise) and a lift enaine cycle.
0014 Gross Weight Initial Condition (WGOO)
rLBS1
0015 Pressure Altitude Initial Condition (HOO)
rFEET?; (normally zero except for partialmission analvsis).
0016 Range Initial Condition (ROO)
rNAUTICAL MILES-; (normally zero exceptfor partial mission analysis).
0017 Time Initial Condition (STOO)
rJOURS?; (normally zero'.
114
0018 Optimum Altitude Indicator (HOPTIN)
0 = Input max alt. for each cruise segment.
1 = Input max alt. Program calculates optimumcruise altitude for segments preceded by aclimb or transfer altitude.
0019 Flight Speed Limit Indicator (VLMIND)
0 = No constraints on eaaivalent airspeed
1 = Max of 250 knots for flight altitudes ator below 10,000 feet as per FAA reas.
0020 Maximum Operating Mach %umber (FMMO)
0021 Maximum Operating Ecuivalent Airspeed (VMO)
0022 Design Dive Speed (VDIV)
0023 Maneuver Load Factor (E.MLF)
rG'S]; As prescribed by Fed Avn Rea 31.
0024 Fuel Req'd Multiplicative Reserve Factor (CKI)
Any fraction greater than 1.0 representsthe percent of reserve fuel (e.a. 1.1reuresents a 10% fuel reserve).
0025 Reserve Fuel Factor (DELWF)
0026 Fuel. Flow Multiplicative Drao Factor (CKFF)
0027 Mission Profile Informatic. (SGTIND)thru0076 0 = End of mission
1 = Taxi2 = Takeoff, hover, and landina3 = Climb4 = Cruise5 = Descent6 = Loiter7 = Chanqe of fuel weight8 = Change of payload weiaht9 = Transfer altitude
10 = X-Y plotter output11 = General performance
100 = End of case
115
Mission profiles are programmed using combinations cf
the following elements:
1. Segment - A unique portion of the mission such ascruise or climb. A segment starts with a set ofinitial conditions and ends when a terminalcondition has been satisfied.
2. Hop - A set of segments ending at some logicalterminal location (such as ground level at thedesired range). Thus, a hop might consist offlying from point "A" to point "B" by means ofcombining the following segments: taxi, takeoff,climb, cruise, descent, landing, and taxi.
3. Leg - A set of hops ending in a refueling of theaircraft.
4, Mission - A set of legs (or hops or seqments)which satisfy specific operational recuirements.In this program, the mission is the basic elementfor which the aircraft is sized.
5. Case - A consecutive series of missions for thesame aircraft. This program permits the user toanalyze a case which consists of a mission forwhich an aircraft is sized, followed bv adiffecent mission which the now sized aircraftperforms, followed by yet additional missions.
An array of segment indicators is input to specify
the mission being studied. A typical array might be:
SGTIND =
1,2,3,4,5,2,1,1,2,3,4,3,4,5,2,7,2,3,4,5,4,5,6,2,1,0,I I -
SI I III SEGMENT REFUELINGI
1<--HOP *1-->-< ---- HOP -2 ---. M ----- HOP *3----->I I1< --------- LG 1 ----------- >< ---- LEG *2 ----- >1II
< . ------------------- MISSION *1 -------------------- >1
0077THRU NOT USED0093
1IF;S
0094 Winq Location for 3-View Drawing (SPACE1(18))
0 - Low
I = High
0095 Location of Engine on Fuselage (SPACE1(19))
Expressed as fraction of length.
0096 Horiz. Tail 1/4 Chord Sweep (SPACE1(20))
0097 Vert. Tail 1/4 Chord Sweep (SPACE1(21))
0098 Aircraft 3-Views Plot Indicator (SPACE1(22))
Greater than 0.0 qenerates 3-view plot
0099 NOT USED
0100 1.0 = NPS THESIS2 format; 0.0 = 133 characterformat
F. AIRCRAFT DIMENSIONAL INFPtRMATION
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0101 Wing Aspect Ratio (DAM2)
(No input when WDMIND = 1,21
0102 Mean Wina Chord to Prop. Diameter Ratio (DAM3)
(No input when WDMIND = 0,2)
0103 ling Incidence Angle (EYFW)
[DEG]; With respect to the fuselage.
0104 Wing Root Thickness to Chord Ratio (TCR)
0105 Wing Tip Thickness Chord Ratio (TCT)
0106 Wina Loading at Design Gross Weight (DAM4)
[LBS/SQ FT3: (No input when WDMIND = 1).
0107 Ouarter Chord Mean Sweep Angle (DLMCH)
rDEG]: (No input if DRGIND=1 & OPTIND=2).
117
0108 Taper Ratio of Wing (SLM)
Tip chord/root chord
0109 Horizontal Tail Aspect Ratio (ARHT)
0110 Horiz. Tail Position on Vertical Tail (SAH)
Fraction of vert. tail span. 1.0 = "T"tail, 0.0 = horiz. tail on or below thevert. tail root chord.
0111 Horizontal Tail Moment Arm (ELTH)
0112 Horiz. Tail Mean Thickness/Chord Ratio (TLCT)
0113 Horizontal Tail Volume Coefficient (VBARH)
(No input when HTIND = 2)
0114 Horizontal Tail Taper Ratio (SLMH)
0115 Horizontal Tail Planform Area (AAW1I)
rSOUARE FEETI; (No input when HTIND = 2)
0116 Prop Blade Attachment Distance (SR)
Measured fron centerline nf huh as afraction of the prop radius. (No inputwhen ENGIND = 1).
0117 Distance Between inboard Prop Tips (YCL)
[FEET): Measured from the inboard prop tipon one side of the fuselage to the inboardprop tip on the opposite side of thefuselage. :No input when WDMINr = 0).
0118 Prop-Over-Prop Overlap (ZETAl)
Measured as a fraction of the prop radius.(No input when WDMIND = 0).
0119 Prop-Over-Wing-Tip Overlap (ZFTA2)
Measured as a fraction of the prop radius.(No input when 1*PMI",qD = 0).
lip
0120 Increment in Wetted Area (DLSWSW)
Utilized for protrusions such as landinggear, etc. Ratio of incremental wettedarea of airplane to wing planform area.
0121 Fuselage Height (HF)
[FEETI; (No input when FDMIND = 0,2).
.0122 Fuselage Length (DAM5)
[FEETI; (No input when FDMIND = 1,2).
0123 Nose Section Fineness Ratio (ELPD)
(No input when FDMIND = 0,2)
S0124 Tail Section Fineness Ratio (ELTD)
(No input when FDMIND = 0,2)
0125 Cabin Section Length (ELC)
rFEET]; Length of constant dia. fuselace(No input when FDMIND = 0,2).
0126 Length of Ramp Well (ELRW)[FEET); May also represent lenath of
strengthened fuselage portion such as thatfor rear engine attachment. Used in thecalculation of fuselage weight penalty.
0127 Fuselage Wetted Area (DAM6)
[SO FT); (No input when FPMIND = 1,2)
0128 Fuselage Width (SWF)
rFEET]
0129 Vertical Tail Aspect Ratio (ARVT)
0130 Vertical Tail Moment Arm (ELTV)
rFEETj
0131 Vert. Tail Mean Thickness Chord Ratio (TCVTI
119
0132 Vert. Tail Volume Coefficient (VBARV)
(No input when VTIND = 2)
0133 Taper Ratio of Vertical Tail (SLMV)
0134 Area of Vertical Tail (AAW12)
(No input when VTIND = 1)
0135 Position of Main Landing Gear (YMG)
Measured outboa£• imm the side of thebody as a fraction of vinq semi-sppn.
0136 Mean Position of Primary Engines (YP)
Measured outboard from airplane centerlineas a fraction of wing semi%.-sp,.n.
0137 Mean Position of Lift Engines (YL)
Measured outboard from aircraft centerlineas a fraction of wing semi-span (No inputwhen LFTIND = 0).
0138 Lift Engine Cluster Gap Factor (EPSLON)
Set by engine type, encine size, and bythe no. of engines which may be clusteredtogether (No input when LFTIND = 0).
0139 Primar- Enq. Nacelle Dimen. Factor (AZETAI)
0140 Primary Enq. Nacelle Dimen. Factor (AZETA2)
0141 Primary Eng. Nacelle Dimen. Factor (AZETA:)
0142 Rotor t/c Ratio at 0.25R (SKIP(I})
0143THRU NOT USED0150
120
G. PASSENGER DATA REQUIRED FOR FUSELAGE SIZING (FDMIND = 2)
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0151 Galley Indicator (DNI1GN)
0 = galley area calculated by program
1 = galley area input by user
0152 Total Area of the Galley(s) (AGLLEY)
rSQUARE FEET]; (No input if DNI1GN = 0)
0153 First Class Section Passenger Capacity (ANPXI)
0154 No. of Seats Abreast in First (lass ANAMl)
0155 No. of Aisle: in First Class (ANISLi)
0156 Width of Seats in i'4rst Class (WSEATI)
[INCHES]; A typical value is ?7 inches.
0157 Seat Pitch in First Class (PSEATI)
[INCHES]; A typical value is 38 inches.
0158 Aisle Width in k'rst Class (WAISLl)
[-NCHES]; A typical value is 20 inches.
0159 Lavatory Indicator (PNIVAL)
0 = Number of lavatories calculated by program
1 = Number of lavatories input by user
0160 Number of lavatories (ANLAVS)
(q'o input when DNIVAL = 0).
0161 Tourist Section Passenoer Capacity (ANPXT)
0162 No. of Seats Abreast in Tourist (ANABT)
0163 No. of Aisles in Tourist (ANISLT)
0164 Width of Seats in Tourist (WSEATT)
[INCHES?; A typical value is 20 inches.
121
0165 Seat Pitch in Tourist (PSEATT)
[TNCHESJ; A typical value is 34 inches.
0166 Aisle Width in Tourist (WAISLT)
[INCHES]; A typical value is 16 inches.
0167 Number in Flight Deck Crew (SPACE2(l))
0168 Number of Flight Attendants (SPACE2(2))
0169vinRU NOT USED0199
H. AIRCRAFT PROPULSION INFO•iMATION REQJJRED WHEN ENGIND = 0
(TURBOSHAFT ENGINES)
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0201 Primary Engine Cycle Number (CYCPRP)
TABLE 1
II I - I IPRIMARY I ENGINE I MAX. TURBINE I COMPRESSOR ! FANCRUISE I CYCLE I INLET TEMP. I DESIGN I BYPASS IENGINES I NUMBER I (DEGREES R) I PRESS RATIO I RATIO I
I II II 1 2600 13 II 2 2600 16 II 3 2900 13 I
TURBOSHAFT I 4 2900 161ENGINE I 5 2900 19 I
CYCLES I 6 3200 13 I
I 7 3200 16 II 8 3200 19 I1 9 3200 22 I
0202 Primary Engine Max. Static Horsepower (PAM7)
Total for all encines at stnd. sea levelconditions (No input when FIXIND = 1).
0203 NOT USED
0204 Number of Primary Engines (ENP)
122
0205 NOf USED
0206 Transmission Efficiency (ETAT)
0257 Transmission Indicator (XMSND)
0 = Trans. sized at fraction of installedpower (see LOC 0258).
1 = Trans. sized at fraction of installedpower (see LOC 0258) or at cruise powerrequired, whichever is more critical.
(No input when ENGIND = 1)
0258 Fraction of Power for Trans. Sizing (XMSMRT)
Ratio of trans. SHP to prim. eng. maxstatic HP (No input when ENGIND = 1).
0259 Accessory Horsepower (DSHPAC)
0223 Number of Rotors or Propellers (ENR)
0224 Propeller Tip Speed (VT)
[FT/SEC]
0225 Disc Loadi.ng [Fan Loading if ETAIND 31 (WCA!
[LBS/SQ FT); Req'd if OPTIND = 2 or 3 (Noinput if OPTIND = 1 AND PDMIND = 1 or 3).
0226 Propeller Diameter (DI)
EFT]; NOTE: If ETAIND=-3, see note in Ref.3, pg. 5-7. (No input when PDMIND=2 or 4)
*** LOCATIONS 0207 - 0212 MUST BE INPUT IF FIXIND = 0 ****
0207 Takeoff Altitude (HES)
rFEET], Typically set eaual to zero.
0208 Thrust-to-Weight Ratio (SENE)
0209 Ambient Temp. Increment for Takeoff (TINY)
._ rDEG FARENHEIT-; Ambient temp. increment"for engine sizing at takeoff conditions(For standard atmosphere, TINY = 0.0).
123
0210 Power Turbine Speed Ratio for Takeoff (AN2TO)
Ratio of operating power turbine speed tomax power turbine speed (No input whenN2IND = 0 or 1). Required when sizingprimary engines for takeoff (see note inRef. 3, page 5-41).
0211 Number of Inoperative Primary Engines (ENPO)
(No input when FIXIND = 0).
0212 Number of Inoperative Lift Engines (ENLO)
(No input when FIXIND = 0).
"* 0260 Fraction of Power (SHPTO)
Ratio of engine SHP to primary engine maxstatic HP (LOC 0202). Required for sizingengines; nominally input as 1.0.
0261 Vertical Rate of Climb for Takeoff (VRCRC)
[FT/MINI; For engine sizing at takeoff.
0262 Takeoff Vertical Climb Power Constant (CKRC)
Climb pot multiplicative constant;Nominally 1 turboshaft engines; Lessfor high diL led aircraft and fans.
*** LOCATIONS 0213 - 0217 MUST BE INPUT IF XMSNIND = 1 I
0213 Power Indicator (POWESI)
0 = Maximum rated power
1 = Military rated power
2 = Normal rated power
(No input when FIXIND = 0 or ESZIND 0)
0214 Cruise Altitude (HC)
[FT]; (No input if FIXIND=0 or ESZIrD=o).
0215 True Airspeed at Cruise (VC)
[KTS]; (No input if FIXIND=-0 or ESZIND=0).
124
0216 Ambient Temp. Increment at Cruise (ATIMIY)
[DEG FAREN.]; (No input when FIXIND = 0 orESZIND = 0. For stnd. atmos., ATMIY - 0).
0217 Power Turbine Speed Ratio for Cruise (AN2CR)
Ratio of operating power turbine speed tomaximum power turbine speed (No input whenN2IND = 0 or 1 or FIXIND = 0 or ESZIND =0). Required if sizing prim. engines forcruise (see note in Ref. 3, page 5-40).
I. PROPELLER DATA REQUIRED WHEN "ENGIND" = 0 (TURBOSHAFT
ENGINES)
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0200 Primary Engine Efficiency Indicicator (ETAIND)
0 = ezopulsive efficiencies input by user
1 = Propulsive table input by user
2 = Propulsive performance calculated by prog.
3 = Fan table input by user
* ** PROPELLER DATA WHEN "ETAIND" = 0 *
0227 Thrust Coeff./Prop. Solidity Ratio (CTSIG)
Ratio of thrust coefficient to propellersolidity. If acft is in the helo mode:Ct = thrust/[(density)*(area)*(tip speeO)l(No input when PDMIND = 1 or 2)
0228 Activity Factor of Propeller (AF)
[PER BLADE]; (No input if PDMIND = 3 or 4)
0229 Number of Blades on Propeller or Rotor (BLDN)
(No input when PDMIND = 3 or 4)
0232 Static Propeller Efficiency (ETAP2)
"Figuze of Merit" for calculations duringtakeoff/hcver/landing (when SGTIND = 2).
125
0233 i:limb Propeller Efficiency (ETAP3)
For climb calculations (SGTIND = 3).
0234 Descent Propeller Efficiency (ETAP5)
For descent calculations (SGTIND = 5).
0235 Mach Number Table (TBEMS)THRU0244 Mach no. values to be paired with primary
engine propulsive efficiencies for usewhen SGTIND = 4 (cruise) or 6 (loiter).
0245 Number of Mach No./Efficiency Pairs (ETAP4N)
0246 Propulsive Efficiency Table (TB8AP4)THRU0255 Primary eng. propulsive efficiency values
to be paired with Mach no. values in LOC0235 thru 0244. Permits rapid evaluationof sensitivity of aircraft perfcrmance andsize to changes in propeller performance.
* ** PROPELLER DATA WHEN "ETAIND" = 1 *
0227 Thrust Coeff./Prop. Solidity Ratio (CTSIG)
Ratio of thrust coefficient to propellersolidity. If acft is in the helo mode:Ct = thrust/[(density)*(area)*(tip speed))(No input when PDMIND = 1 or 2).
0228 Activity Factor of Propeller (AF)
"[PER BLADE]; (No input if PDMIND = 3 or 4)
0229 Number of Blades on Propeller or Rotor (BLDN)
(No input when PDMIND = 3 or 4)
0234 Descent Propeller Efficiency (ETAP5)
For descent calculations (SGTIND = 5).
0256 Propeller Table Number (CYPROP)
CYPROP values for live general avn propsare given in the table below.
126
TABLE 2
PROPELLER CHARACTERISTIC SUMMARYALL PROPELLERS ARE 3-BALDED, CONSTANT SPEED
I DESIGNATION1 INTEGRATED I ACTIVITY IIMANUFACTURERI I DESIGN LIFT! FACTOR I APPLIC. I
I [TABLE NO.11 COEFFICIENT! PER BLADE!
IHARTZELL T10282H 114 TWIN1PROPELLERS, 1 1 0.555 1 OTTER IIINC. [10282.3) 1 (118) 1 SKYVAN
IHARTZELL T10173-8 0.620 104 BEECHIPROPELLERS, I I KING AIR!IINC. 1 [101738.3) 1 (0.700) 1 (107) 1 99
1HAMILTON 13TLF 1033A-01 HAWKISTANDARD 1 1 0.424 1 127 1 CMDRDIV., UAC 1 [1033..3) 1
IHAMILTON 133LF 1027A-01ISTANDARD 1 1 0.500 I 110 1IDIV., UAC 1 [1027.3] 1
IHAMILTON 133LF 1013A-01ISTANDARD 1 1 0.483 I 97 IIDIV., UAC 1 [1013.3) 1
NOTE: Values :n parenthesis are quoted by Hartzell Props,Inc. Values not in parentheses are consistent withthe blade geometric Oata supplied by Hartzell.
1700 User's Propeller Table Number (PROPCY)
A table of actual prop performance dataobtained from test data can be input.Refer to Ref. 3, page 4-63 for a completedescription of the table requirements.
1701 Number of Advance Ratio Values (XPXNO)
To be assigned to loc 1702-1721 NOTE: Ifused, XPXNO must at least equal 3.
127
1/02 Propeller Advance Ratio Table (XPJ)THRU1721 Table of propeller advance ratio values to
be input by user (minimum of 3 values).
1722 Number of Prop Power Coefficients (CPPNO)
No. of prop pwr coeff. to be input by theuser in loc 1723-1742. If useda table of at leasts used, a value ofat least "3" must be assigned to CPPNO.
1723 Propeller Power Coefficients (CPPROP)THRU1742 Table of propeller power coeff. values to
be input by user (minimum of 3 values).
1743 Values of Prop Thrust Coefficients (CTPROP)THRU2142 Table of propeller thrust coeff. values.
** * PROPELLER DATA WHEN "ETAIND" = 2 *
0227 Thru3t Coeff./Prop. Solidity Ratio (CTSIG)
Ratio of thrust coefficient to propellersolidity. If acft is in the helo mode:Ct = thrust/[(density)*(area)*(tip speed)l(No input when PDMIND = 1 or 2)
0228 Activity Factor of Propeller (AF)
[PER BLADE]; (No input if OPTIND = 1 andPDMIND-3,4; ALWAYS req'd if OPTIND 2, 3)
0229 Number of Blades on Propeller or Rotor (BLDN)
0230 Prop Integrated Desigr Lift Coeff. (CLEYE)
0234 Descent Propeller Efficiency (ETAP5)
For descent calculations (SGTIND = 5).
*****i****** PROPELLER DATA WHEN "ETAIND" = 3 ************
0227 Thrust Coeff./Prop. Solidity Ratio (CTSIG)
Ratio of thrust coefficient to propellersolidity. if acft is in the helo mode:Ct = thrust/[Pdensity)*(area)*(tip speed)l(No input when PDMINP = I or 2)
12P
0228 Activity Factor of Prcpeller (AF)
EPER BLADE]; (No input if PDMIND = 3 or 4)
0229 Number of Blades on Propeller or Rotor (BLDN)
(No input when PDMIND = 3 or 4)
0230THRU NOT USED0233
0234 Descent Propeller Efficiency (ETAPS)
For descent calculations (SGTIND = 5).
0256 Fan Table Number (CYPROP)
User selected no. to identify Fan Table.
0408 Tiltina Mechanism Constant (SKTM)
Tilt wing cr tilt rotor tilt mechanismweight factor. This constant calculates atilt mechanism weight proportional to theacft gross weight.
0457 Rotor/Prop Weight Adjustment Factor (SKRP)
To avoid using the prop weights eauations,input this variable value as zero.
1700 User's Fan Table Number (PROPCY)
As assigned by user in LOC 0256.
1701 Number of Mach Number Values (XPXNO)
To be assigned to locations 1702 - 1721NOTE: If this table is used, a value ofleast "3" must be assigned to XPXNO.
1702 Mach Number Values (XPJ)THRU1721 Table of Mach Number values to be input by
the user (minimum of 3 values).
1722 Number of Referred Power Coefficients (CPPNO)
No. of referred power coefficients to beinput by the user to LOC 1723 - 1742.
129
NOTE: If this table is used, a value ofat least "3" must be assigned to CPPNO,
1723 Referred Power Coefficient Values (CPPROP)THRU1742 &able of referred power values to be
input by the user (minimum of 3 values).
1743 Referred Thrust Coefficient Values (C?.PROP)THRU2142 Table of referred thrust coeff. values.
J. AIRCRAFT PROPULSION INFORMATION REQUIRED WHEN ENGIND = 1
(TURBOJET OR TURBOFAN ENGINES)
LOCJTION DATA DESCRIPTION (FORTRAN NAME)
0201 Primary Engine Cycle Number (CYCPRP)
To be selected from the TABLE 3 (below).
TABLE 3
II I i I iPRIMARY I ENGINE I MAX. TURBINE I COMPRESSOP I FAN ICRUISE I CYCLE I INLET TEMP. I DESIGN I BYPASSENGINES 1 NUMBER I (DEGREES R) I PRESS RATIO I RATIO I
I II Ii 10 2600 I 13
11 I 2600 I 1.6I 12 2900 i 13 I
"ITURBCJZT 13 2900 i 16 IENGINE 14 2900 1 19 ICYCLES 15 3200 13 I
16 3200 16 I17 3200 I 19 Ii18 3200 22 I
I -I I I119,20,211 2600 16 2,4,6 1
1 122,23,241 2600 20 I 2,4,6 1125,26,271 2q00 16 2,4,6 I
TURBOFAN 12S,29,301 2900 20 I 2,4,6 1I ENGINE 131,32,331 2900 24 I 2,4,6 11 CYCLES 134,35,361 3200 16 I 2,4,6
137,38,391 3200 20 i 2,4,6 (140,41,421 3200 24 I 2,4,6143,44,451 3200 1 28 I 2,4,6
130
0203 Primary Engine Maximum Static Thrust (DAM8)
Total thrust for all engines at stnd. sealevel cLonditions (No input if FIXIND = 1).
0204 Number of Primary Engines (ENP)
0218 Lift Engine Cycle Number (CYCLFP)
See table below (No input if LFTIND = 0).
TABLE 4
IENGINE MAX. TURBINE COMPRESSOR FANLIFT I CYCLE INLET TEMP. DESIGN BYPASSENGINES I NUMBER (DEGREES R) PRESS RATIO RATIO I
I INDEPENDENTI46,47,481 2400 1 7 1 2,4,6 I1 LIFT 149,50,511 2700 1 7 1 2,4,6 11 ENGINES 152,53,541 3000 1 7 1 2,4,6 11I
155,56,571 2600 I 13 18,11,14 I158,59,601 2600 1 16 12,11,14 I161,62,631 2900 13 18,11,14 1
1 GAS COUPLEDI64,65,66I 2900 16 18,11,14 I167,68,691 2900 19 18,11,14 I
LIFT FANSI70,71,721 3200 13 18,11,14 I!73,74,751 3200 1 16 18,11,14 1176,77,781 3200 1 19 18,11,14 1179,80,61I 3200 1 22 18,11,14 1
0219 Lift Engine Maximum Static Thrust (DAM9)
Total for all eng; stnd. sea level;(No input if FIXIND = 1 or LFTIND = 0).
0220 Number of Lift Engines (ENL)
(No input when LFTIND = 0)
0221 Number of Clusters of Lift Engines (ENC,
(No input when LFTIND = 0)
0231 Lift Engine Efficiency (ETAC)
(No input when LFTIND = 0).
131
0232 Primary Engine Propulsive Efficiency (ETAP2)
To be used for Takeoff, Hover, and Landingconditions (SGTIND = 2).
0207 Takeoff Altitude (HES)
[FEET]; Required for sizing engines and istypically set equal to zero.
0208 Thrust-to-Weight Ratio (SENE)
0209 Ambient Temp. Increment for Takeoff (TINY)
rDEG FARENHEITI; Ambient temp. incrementfor engine sizinqr at takeoff conditions(For standard atmosphere, TINY = 0.0).
0210 Power Turbine Speed Ratio for Takeoff (AN2TO)
Ratio of operating power turbine speed tomaximum power turbine speed (No input whenN2IND = 0 or 1 or FIXIND = 0). Reauiredwhen sizing primary engines for takeoff(see note in Ref. 3 page 5-41).
0211 Number of Inoperative Primary Engines (ENPO)
Required for engine sizing (No input whenFIXIND = 0).
0212 Number of Inoperative Lift Engines (ENLO)
Required for engine sizing (No input whenFIXIND = 0).
* LOCATIONS 0213 - 0217 MUST BE INPUT IF LFTIND = 1 ** ***************************** **** *** ** ** *** *** *** *** ****** **
0213 Power Indicator (POWESI)
0 = Maximum rated power
1 = Military rated power
2 = Normal rated power
(No input when FIXIND = 0 or ESZIND = 0)
132
0214 Cruise Altitude (HC)
[FEET]; This data is required for sizingthe engines (No input when FIXIND = 0 orESZIND = 0).
0215 True Airspeed at Cruise (VC)
[KNOTS]; This data is required for sizingthe engines (No input when FIXIND = 0 orESZIND = 0).
0216 Ambient Temperature Increment at Cruise(ATMIY)
[DEGREES FARENHEITJ; This data is usedfor sizing the engines (No input whenFIXIND = 0 or ESZIND = 0. For standardatmosphere, ATMIY = 0.0).
0217 Power Turbine Speed Ratio for Cruise (AN2CR)
Ratio of operating power turbine speed tomaximum power turbine speed (No input whenN2IND = 0 or 1 or when FIXIND = 0 orESZIND = 0). Required when sizing primaryengines for cruise (see note in Ref. 3,page 5-40).
0200 Primary Engine Propulsive Efficiency Indicator(ETAIND)
MUST be input as "3". The user isrequired to input a fan table (Locations1700 to 2142). Also, the rotor/propellerweight adjustment factor must be input aszero (Location 0457).
K. AIRCRAFT PROPULSION INFCRMATION REOUIRED WHEN ENGIND = 2
(CONVERTIBLE ENGINES)
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0201 Primary Engine Cycle Number (CYCPRP)
To be selected from Table 5 on page 134below. NOTE: The number entered at thislocation must match the number entereO atlocation 1301 to avoid an error message.
133
TABLE 5
II I I IPRIMARY I ENGINE AXL.. TURBINE I COMPRESSOR I FANCRUISE I CYCLE INLET TEMP. I DESIGN I BYPASSENGINES I NUMBER (DEGREES R) I PRESS PATIO I RATIO
1 2600 132 2600 163 2900 13
TURBOSHAFT 4 2900 16ENGINE 5 2900 19CYCLES 6 3200 13
7 3200 168 3200 199 3200 22
1Ii0 2600 13I II 2600 16I 12 2900 13
TURBOJET I 13 2900 16ENGINE 1 14 2900 19CYCLES I 15 3200 13
I 16 3200 16I17 3200 19
i 18 3200 22
119,20,211 2600 16 1 2,4,6r122,23,241 2600 20 I 2,4,C,125,26,271 2900 16 I 2,4,6
I TURBOFAN 12S,29,301 2900 20 I 2,4,6I ENGINE 131,32,331 2900 24 I 2,4,6I CYCLES 134,35,36! 3200 16 1 2,4,6
137,38,391 3200 20 I 2,4,6140,41,421 3200 24 I 2,4,6143,44,451 3200 28 I 2,4,6
134
0203 Primary Engine Maximum Static Tnrust (DAM8)
[LBS-FORCE]; Total for all engines atstnd. sea level (No input if FIXIND 1).
0204 Number of Primary Engines (ENP)
0205 Convertible Engine Conversion Ratio (BETA)
rLBS-FORCE PER HORSEPOWER]
0206 Transmission Efficiency (ETAT)
0257 Transmission Indicator (XMSND)
0 = Trans. sized at fraction of installedpower (fraction at LOC 0258).
1 = Trans. sized it fraction of installedpower (fraction at location 0258) or atcruise power required, whichever is morecritical.
(No input when ENGIND = 1)
0258 Fraction of Power for Trans. Sizinq (XMSMRT)
Ratio of transmission shaft horsepower toprimary engine maximum static horsepower(location 0202)
(No input when ENGIND = 1)
0259 Accessory Horsepower (DSHPAC)
[HORSEPOWER]
0223 Number of Rotors or Propellers (ZNR)
0224 Propeller Tip Speed (VT)
FFEET PEP SECOND]
0225 Disc Loaiinq fFan Loadinc if FTAIND = 31 (WGA)
rLBS/SO FT3; This is always required ifOPTIND = 2 or 3 (No input whien OPTIND = IAND PDMIND = 1 or 3).
135
0226 Propeller Diameter (DI)
[FEET]; If ETAIND = 3, see note in Ref. 1page 5-7 (No input when PDMIND = 2 or 4).
0227 Thrust Coeff./Prop. Solidity Ratio (CTSIG)
Ratio of thrust coefficient to propellersolidity. If acft is in the helo mode:Ct = thrust/[(density)*(area)*(tip speed)l(No input when PDMIND = 1 or 2)
0228 Activity Factor of Propeller (AF)
[PER BLADE]; (No input if PDMIND = 3 or 4)
0229 Number of Blades on Propeller or Rotor (BLDN)
(No input when PDMIND = 3 or 4)
0232 Static Propeller Effic.ency (ETAP2)
"Figure of Merit" for calculations duringtakeoff, hover, and landing (SGTIND = 2).
* LOCATIONS 0207 - 0212 MUST BE INPUT IF FIXIND = 0 *
0207 Takeoff Altitude (HES)
[FT?; Typically set equal to zero.
0208 Thrust-to-Weight Ratio 1SENE)
0209 Ambient Temp. Increment for Takeoff (TINY)
FDEG FARENHEIT]: Ambient temp. incrementfor engine sizing at takeoff conditions(For standard atmosphere, TINY = 0.0).
0210 Power Turbine Sneed Ratio for Takeoff (AN2TO)
Ratio of operating power turbine speed tomaximum power turbine speed (No input whenN2IND = 0 or 1 or FIXIND = 0). (see notein Ref. 3, page 5-41).
021i Number of Inoperative Primary Engines (ENPO)
(No input when FIXIND = 0).
136
F!
0212 Number of Inoperative Lift Engines (ENLO)
(No input when FIXIND = 0).
0260 Fraction of Power (SHPTO)
Ratio of engine SHP to primary engine maxstatic HP (LOC 0202). Nominally set equalto 1.0.
0261 Vertical Rate of Climb for Takeoff (VRCRC)
[FT/MINs: Req'd for sizing eng. at takeoff
0262 Takeoff Vertical Climb Power Constant (CKRC)
Climb power multiplicative constant;nominally 2.0 for turboshaft engines; Lessfor high disc loaded aircraft and fans.
0213 Power Indicator (POWESI)
0 = Maximum rated power
1= Military rated power
2 = Normal rated power
(No input when FIXIND = 0 or ESZIND = 0)
0214 Cruise. Altitude (HC)
[FT]: (No input if FIXIND=O or ESZIND=O).
0215 True Airspeed at Cruise (VC)
[KTS]; (No input if FIXIND=-0 or ESZIND0O).
0216 Ambient Temp. Increment at Cruise (ATMIY)
[DEG FARENHEITI; (No input if FIXIND = 0or ESZIND as zero).
0217 Power Turbine Speed Ratio for Cruise (AN2CR)
Ratio of operating power turbine speed tomaximum power turbine speed (No input whenN2IND = 0 or 1 or FIXIND = 0 or ESZIND =0). Required when sizina orimary enainesfor cruise (see note in Ref. 3, paqe5-40).
.N 137N9
L. "TRCRAFT AERODYNAMICS INFORMATION
_•TION DATA DESCRIPTION (FORTRAN NAME)
0301 Vertical Tail Profile Drac Coefficient (CDVTI)
Based on vertical tail planform area and aReynold's number of 1.OE+07.
0302 Lift Eng. Nacelle Profile Drag Coeff. (CDHTI)
Based on total wetted area of nacellecluster and a Reynold's number of 1.0E+07.
0303 Primary Eng Nacelle Profile Drag Coeff. (CDNI)
Based on wetted area of all nacelles and aReynold's number of 1.OE+07.
0304 Horizontal Tail Profile Drag Coeff. (CDLNI)
Based on horizontal tail planfom area anOa Revnold's num'ser of l.OE+07 (No inputwhen LFTIND = 0).
0305 Profile Drag Increment (DELCD)
Based on winq planform area
0306 Oswald's Efficiency Factor (DAM]0)
Spanwise efficiency factor (No input whenOSWIND = 1).
0307 Ecuivalent Flat Plate Area Increment (DELFFI
rSO FTI; Based on fuselaae parasite ararT.
0308 Number of Pairs in Table (TLLN)
Number of pairs of values in the table ofwino profile erac coeff. (LOC 0335 - 03421versus lift coeff. (LOC 0317 - 032A).
0309 Number of Mach Numbers (TENN)
Nmber of Mach no. values in LOC 0325-0320for use in table of comoressibilitv e:aaas a function of Macli number and liftcoefficient (No input when PRGIND =
13R
0310 Number of Lift Coefficients (TCLZN)
Number of lift coefficient values in LOC0343 - 0349 for use in the table ofcompressibility drag as a function of Mach.no. and lift coeff. (No input if DPGIND=-0)
0311 Lift Nacelle Multiplicative Drag Factor (CKLN)
0312 Wing Multiplicative Drag Factor (CKW)
0313 Prim. Nacelle Multiplicative Drac Factor (CKN)
0314 Fuselage Multiplicative Drac Factor (CKF)
0315 Vert. Tail Multiplicative Drag Factor !CKVT)
0316 Horiz. Tail Multiplicative Drao Factor (CKHT)
0317 Lift Coefficient Values (TRCL!)THRU0324 For the table of wina profile draq coeff.
versus lift coefficients
0325 Mach Number Values ITBEM)THRU0329 For the table of compressibility arag as a
function of Mach number and lift coeff.(No input when DRGIND = 0).
S0330 Mean Reynold's No. per Foot for Misson (RECI)
0331 Two-Dimensional Lift Coefficient Slope (CSALF)
[PER RADIAN!
0332 Zero Lift Anole of Attack (ALPHL)
0333 Nondimensional Position AMona the ChorO (XCPS)
X/C (No input when DRGIND = 1).
0334 Nondimen. Pcsn. Alona Cnord at Max t/c (XCTCM)
(X/C)max t/c; (No input when DPGTNP = 1).
0335 Wir.n Profile Draq Coefficient Values (TRCPWI)THRU0342 Based on wino planform area at Revnol's
Number of I.OE+07 for the table of winaprofile drac coeff. vs lift coeff.
139
0343 Lift Coefficient Values (TBCL2)THRU0349 For the table of compressibility drao as a
function of Mach number and lift coeff.(No input wnen DRGIND = 0).
0350 Drag Increment (TBCDM)THRU0384 Increase in airplane draa due to Mach
number (compressibility effects). Tnputthis table as a function of Macb no. andlift coeff. hased on wing planform area.
0385 Supercritical Factor (SPACE4(l))
0.5 = 50% of technolocy
1.0 = 100% of technoloqv
0386 Max Lift Coeff. to Compute VMC (SPACEA(2))
0387 CLVRD CL of Vert. Tail & Rudder (SPACE4(3))
0388
THRU NOT USED0393
M. ROTOR, PROPELLER, AND GEARBOX WEIGHT
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0394 No. of Staces in Main Rotor Drive (SPACEA(10))
0395 Blade Fold Penalty (SPACE4(11))
Default = 1.0 (no folO)
0396 Hub Weicht Coefficient (SPACE4(12))
0397 Hub Material/Pevelooment Factor (SPACE4(13))
0398 Blade Weioht Coefficient (SPACE4114A)
C3r-9 Rotor Type Factor (SPACE4(15))
1.0 = Fully articulated
2.2 = Hinceless or teeterina
9.1 = X-Winn
140
N. AIRCRAFT WEIGHT INFORMATION
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0400 Operating Weight Empty (OWEl)
[LBS]; (No input when OPTIND = 1 or 2).
0401 Weight of Fixed Equipment (WFE)
rsi s.
0402 Weight of Fixed Useful Load (WFUL)
FLBS3
0403 Weight of Payload (WPL)
[LBSI: (No input when OPTIND = 2).
0404 Cockpit Controls Constant (SKCC)
0405 Fixed-Wing Controls Constant (SKFW)
0406 System and Hydraulics Constant (SKH)
0407 Factor for Stability Auqmented System (SKSAS)
Also includes mixing units
0408 Tiltina Mechanism Constant (SKrM)
0409 Upper Control Mechanisms Constant (SKUC)
LOCATIONS 0410 - 0415 ARE NOMINALLY SET EOUAL TO 1.0 *
0410 Cockpit Controls Weight Factor (CK1IS
0411 Upper Controls Weiaht Factor (CK16)
0412 Hvdraulics Weiaht Factor (rK17)
0413 Fixed Wing Controls Weioht Factor (CK]P)
0414 SAS Weight Factor 'CK19)
0415 Tilt Mechanism Weiqht Factor (CK20)
141
0416 Number of Temperature Pairs (THN)
Number of atmosphere temperature pairs inlocations 0440 - 0449 and 0466 - 0475 (Noinput if ATMIND is never set eaual to 2).
* LOCATIONS 0417 - 0419 ARE NOMINALLY SET EOUAL TO 0.0 *
0417 Fit Controls Group Incremental Weight (PELTFZ)
c. rLBSJ
* 0418 Propulsion Group Incremental Weight (DELWP)
-:. [L B S ]
•-0419 Structures Group Incremental Weight ý,DELWST)rLBS]
0420 Body Group Weight Adjustment Factor (SKP)
0421 Lift Engine Section Weight Factor (SKLES)
(No input when LFTIND = 0)
0422 Alightinq Gear Weight (SKLG)
Expressed as a percentace of Qross weiaht.
0423 Main Gear Weight to Gross Weight Ratio (SKMG)
0424 Tail Load Adjustment Factor (SKTLI
0425 Wina Bending Relief Moment Adj. Factor (SKWFI
0426 Wina Type Weioht Adjustment Factor (SFWWI
0427 Pitch Radius of Gyration (SKY)
rFTi
0423 Yaw Radius of Gyration (SKZ)
[FT?
0429 Primary Engine Section Weiaht Factor (SKPFSI
142
***************************************** ********************
NO INPUT TO LOCATIONS 0430 - 0432 UNLESS SKPES 0 (LOC 0429)
0430 Engine Nacelle Type Factor (SKTr)
0431 Engine Nacelle Adjustment Factor (SKNAC)
0432 Engine Attachment Point Ratio (SKLMT)
Distance between enqine center of gravityand closest structural attachment pointbetween nacelle and wing expressed as aratio to the length of the nacelle.
0433 Wing Weight Multiplicative Weiaht Factor (CKP)
0434 Hcriz. Tail Wt Multiplic. Weight Factor (CK9)
0435 VWrt. Tail Wt Multipli-. Weiaht Factor (CKlO)
0436 Fuselage Wt Multiplic. Weiaht Factor (CKli)
0437 Landinq Gear Wt Multiplic. Wt Factor (CK121
0438 Lift Enc. Section Multiplic. Wt Factor (CKI3)
0439 Prir-arv Enq. Sec. Multiplic. Wt Factor (CK141
0440 Non-Standard Atmosphere Altitude (TPH)THRU0449 FFTl1 Altitudes to be paired with ambient
temperature ratios (LOC 0466 - 0475) forthe non-standard atmosphere table.
0450 Cabin Differential Pressure Limit (DrLP!
rpsi))
0451 Weight of Concentrated Load (WC)
rLl3S1
0452 Concentrated Load Position (YC)
Distance of load out)'oard from aircraftcerterline. Expressed as a fraction of
"C. the wino semi-span.
143
S0453 Drive System Weight Adjustment Factor (SKDS)
0 = No gearbox weight.
1 - Ham, standard gearbox weiaht trend.
(No input when ENGIND = 1)
0454 Fuel System Weight Adjustment Factor (SKFS)
0435 Lift Engine Installation Weight Factor (SKLEI)
(No input when LFTIND - 0)
0456 Primary Engine Install. Weight Factor (SKPEI)
0457 Rotor or Prop Weight Adjustment Factor (SKRP)
-1 = Use HESCOMP rotor and drive weiohttrend. (LOC 0394-0399 rotor coeff.,LOC 0453 drive coeff. Also, LOC 0142if not X-Winc configurationS.
0 = No prop wt (Use if ETAIND=3 LOC 0200).
1 = 1970 Ham. standard prop weight factors
2 = 1980 Ham. standard prop weioht factors
0458 Drive Sys. Weight Variation Adj. Factor (SKVT)
Adjustment factor for variations in drive
system weight due to nonuniformities inhover tipspeed and transmission tipspeeeor the maximum power and the transmissiondesign power are not the same. Thenominal value is 1.0 when these parametersare similar. The value of SKVA' wi]l varywhen tip speed and ; •wer change asindicated by the following expression:
(Design Tipspeed) (Maximum Power)X
(Hover Tipspeed) MDesian Power)
Airplane category in Ham standard prop and
* gearbox weight can be used by the input ofa negative value of the catecory f-I, -2,-3, etc.1 (No input when ENGIND = 1).
144
4
0459 Propeller Group Multiplic. Weight Factor (CK2)
0460 Drive System Multiplic. Weight Factor (CK3)
0461 Lift Engine Multiplic. Weight Factor (CK4)
0462 Primary Engine Multiplic. Weight Factor (CK5)
0463 Lift Eng. Install. Multiplic. Wt Factor (CK6)
0464 Prim. Enq. Install. Multiplic. Wt Factor (CK7)
0465 Fuel System Multiplic. Weight Factor (CK21)
0466 Ambient Temperature Ratio Values (TBTHE)THRU0475 To be paired with alt. (LOC 0440 - 0449)
for the non-standard atmosphere table.
0. ENGINE ACOUSTICAL TREATMENT
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0476 Engine Acoustic Treatment Weight TrendCoefficient (SPACES51))
0477 Engine Weight Treatment Factor (SPACE5(2))
(= 0477 * 0478 * WEP)
0478 Multiplicative Factor (SPACE5(3))
0479 NOT USED
P. WEIGHT OF FIXED EOUIPMENT (WFE)/FIXED USEFUL LOAD (WFUL)
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0480 WFE/WFUL Indicator (SPACES(5))
0 = WFE and WFUL input in LOC 0401 and 0402.
1 = WFE and WFUL computed by proaram.
0481 APU Trend Coefficient (SPACFr5f6))
"0482 Instruments TrenA Coefficient (SPACES(7))
0483 Hydraulics Trend Coefficient (SPACE5(8))
145
0484 Electrical Trend Coefficient (SPACE5(9))
0485 Avionics Trend Coefficient (SPACE5(10))
0486 First Class Furnishings Coeff. (SPACE5(1l))
0487 Tourist Class Furnishings Coeff. (SPACE5(12))
0488 Air Conditioning Coefficient (SPACE5(13))
0489 Anti-Icing Coefficient (SPACE5(14))
"0490 Auxiliary Gear Coefficient (SPACF5(15))
0491 Crew Baggage (SPACES(16V)
rLBS/PERSONI
0492 First Class Passenger Service (SPACE5(17))
rLBS/PASSENGER1
0493 Tourist Class Passencer Service (SPACE5(]Pfl
rLBS/PASSENGERI
0494 Water Allocation (SPACE5(19))
"r"LBS/PERSON3
"0495 Emergency Eauipment (SPACE5(20))
(LBS!
0496 Crew Caterinc (SPACE5(21))
rLBS/CREW)
0497 First Clars Caterina (SPACF5(22))
rLBS/PASSENGFR10498 Tourist Class Caterina (SPACF5(23))
rLBS/PASSENGER1
0499 Unusable Fuel Factor (SPACF5(2d))
14r6
0. TAXI INFORMATION {SGTIND = 1I
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0501 Taxi Segment Atmosphere Indicator (ATMINI)THRU0510 0 = Standard atmosphere.
1 = Non-standard atmosphere. User inputs asingle point value for the increment inambient temperature above the standard-dayvalue.
0511 Incremental Time for Taxi (DELTT)THRU0520 EHRS]
0521 Ambient Temperature Increment (TIN1)THRU0530 [DEG FARENHEIT]; Used for engine sizing at
TAXI conditions (No input when ATMIND = 0or 2).
0531 Lift Engine Taxi Segment Factor (SKFL)THRU0540 0 = Lift engines .off during taxi.
1 = Lift engines operating during taxi.
0541 Power Turbine Speed Ratio (AN2MI)THRU0550 Ratio of operating power turbine speed to
maximum power turbine speed (input forboth primary and auxiliary independentengines in performance segment 3, TAXI!.See additional information in Ref. 1, page5--40. (No input when N2IND = 0 or 1).
R. TAKEOFF, HOVER, AND LANDING INFORMATION [SGTIND = 21
LOCATION DATA DESCRIPTION (FG7"'AN NAME)
0601 Takeoff, Fover, and Landing Indicator (TOLIND)THRU0610 1 = User inputs the required thrust-to-weight
ratio. Airplane will use:
a. Max power from litt engines beforeaugmenting with primary engines.
147
F
b. Only power from primary engines ifLFTIND = 0.
2 = User inputs the required thrust-to-weightratio. Airplane uses equal percentages ofpower from lift and primary engines. DONOT INPUT TOLIND = 2 IF LFTIND = 0.
3 = User inputs rea'd fraction of max power.
0611 Atmosphere Indicator for SGTIND = 2 (ATMIN2)THRU0620 0 = Standard atmosphere
1 = Non-standard atmos. User inputs sinalepoint value for increment in ambient temp.above the standard day value.
2 = Non-standard atmos. User inputs table ofambient temp. ratios as a function of alt.
0621 Primary Eno Power (or Thrust) Factcr (PFET2)THRU0630 Required when TOLIND = 3 (No input when
TOLIND = 1 or 2).
0631 Ambient Temperature Increment (TIN2)THRU0640 [DEG FARFNHEITI; Used for engine sizing at
Takeoff cnditions (No input if ATMIND = 0or 1).
0641 Lift Er,*ine Thrust Fraction (FLET2)THRU0650 Req'd if TOLIND=3 (No input if TOLIND=1,2)
0651 Thrust-to-Weiqbt Ratio for Takeoff (ENT)THRU0660 (No input when TOLIND = 3)
0661 Step Size for Hover (PELTrH)THRU0670 rHPSi
0671 Power Tarbine Speed Ratio (AN2M2)THRU0680 Ratio of operatina power turbine speed to
maximum power turbine speed finnut forboth primary and auxiliary independentengines in performaice seament 7, TAKEOFF,HOVER, LANDINGI (No input if N2IND = 0,1).
14P
0681 Incremental Time for Hover (STH)THRU0690 [HRS]
2321 Vertical Rate of Climb for Takeoff (VRCTO)THRU2330 [FT/MINJ
S. CLIMB INFORMATION [SGTIN4D = 3)
LOCATION DATA DESCRIPTION (FORTRAN NAIME)
0691 Climb Indicator (CLMIND)THRU0700 1 = Maximum rate of climb
2 = Climb at constant equivalent airspeed
3 = Climb at constant Mach number
4 = Climb at constant true airspeed
0701 Mach, Equiv. Airspeed or True Airspeed (EMACIS)THRTJ0710 [KTS.; (No input when CLMIND = I).
0711 Climb Segment Atmosphere Indicator (ATMIN3)THRU0720 0 = Standard atmosphere
1 = Non-standard atmos. User inputs singlepoint value for increment in ambient temp.above the standard day value.
2 = Ncn-standard atmos. User inputs table ofambient temp. ratios as a function of alt.
0721 Step Size for Climb Seament (DELH3)THRYT0730 FFTI
0731 Ambient Temperature Increment (TIN3)THRU
S0740 rDEG FARENH7IT 1 ; Used for sizina enoinedurinq climb (No input when ATMINP = 0,2).
0741 Max Altitude for Climb or Alt. Transfer (HMAXITHRU0750 fFT]
14ý
0751 Climb Segment Power Indicator (POWCLI)THRU0760 0 = Maximum power
1 = Military power
2 = Normal power
0761 Max Body Attitude Angle for Climb (THEMAX)THRU0770 [DEG]
0771 Power Turbine Sneed Ratio (AN2M3)THRU0780 Ratio - operating pwr turbine speec to max
pwr turbine sraeed f input for both primarvand auxillary independent enq. in sec. 3,CLIMB). (No input if N2IND = 0 or 1).
0781 Profile Drag Increase Durina Climb (DCLIMB)THRU0790
0791 Incremental Normal Load Factor (ENCLIMP)TH P0800 For enervy-maneuverabilitv calculations
(Nominally set ecual to 0.0).
T. CRUISE INFORMATION fSGTIND = 41
LOCATION DATA DESCRIPTION (FORTRAN NAM2)
0801 Cruise Indicator (CRSIND)"THRU0810 1 = Cruise at cruise power.
2 = Cruise at constant true airspeen.
3 = Cruise at speed for best specific rtnae.
A = Cruise at speed for 99% of best specifit"rance.
5 = Cruiqe-climb (constant acft wt to ambientoress ratiol at speed for best snecificrange.
6 = Cruise-climb (constant airplane weiaht toambient pressure ratio) at speeO for 90%of best specific rance.
150
'I*
0811 True Airspeed or Headwind (VIN)THRU0820 rKTS]; Input true airspeed if CRSIND = 2;
Input headwind when CRSIND = 3 thru 6.
0821 Cruise Segment Atmosphere Indicator (ATMIN4)THRU0830 0 = Standard atmospihere
1 = Non-standard atmos. User inputs sinalepoint value for increment in ambient temp.above the standard day value,
2 = Non-standard atmos. User inputs table ofambient temp. ratios as a function of alt.
0831 Step Size for Cruise Seament (DELR)THRU0840 FNAUTICAL MILESI
0841 Ambient Temperature Increment (TIN4)THRU0850 [DEG FARENHEIT3; Used for sizina encine
durinq cruise (No input if ATMIND = 0,21.
0851 Ranoe at End of Cruise. Seqment (RMAX)THRU0860 rNAUTICAL AILES1
0861 Cruise Segment Power Inlicator (POWCRU)THRU0870 0 = Maximum power
I = Military power
2 = Normal power
0871 Number of Pr-imary Encines Shutdown (INPSN)THRU0880 Durinq cruise seament
0881 Power Turbine Sreed Ratio (AN2M41THRU0890 Ratio of operatinn power turbine speee to
maximum power turbine speeO Jinput forboth primary and auxiliary independentengines in performance seament 4, CRUISE].(No input if N21NP = 0 or 1).
151
0891 Profile Drag Increase (DLCDCR)THRU0900 [SO. FT.]; Drag increase during cruise due
to enqines beina ahut down (based on winaplanform area).
U. DESCENT INFORMATION fSGTIND = 51
LOCATION DATA DESCRIPTION (FORTRAN NAME)
0901 Descent Indicator (DESIND)THRU0902 1 = Descend at maximum speed, terminal range
specified.
2 = Descend at maximum speed, terminal ranaenot specified.
3 = Descend at idle power, terminal ranaespecified.
4 = Descend at idle power, terminal rance notspecified.
5 = Descend at constant eauivalent airspeed,terminal range specified.
6 = Descend at constant eauivalent airspeed,terminal rance not specified.
7 = Descend at constant Mach number, terminalrange specified.
S= Descend at ccnstant Mach number, term inalrance not specified.
0911 Mach, Equiv. Airspeed or True Airspeed (FMACH')TIPU0920 [KTS]; (No input if DESIND = 1,2,3,or,4).
0921 Descent Seament Atmosphere Indicator (ATMIN91THRU0930 0 = Standard atmosphere
I = Non-standard atmos. User inputs sinalepoint value for increment in ambient temPx.above the standari day value.
2 = Non-standard atmos. User inputs tabhe ofambient temp. ratios a3 a function of alt.
152
0931 Minimum Body Attitude Anqle, Descent (THEMIN)THRU0940 CDEG)
0941 Ambient Temperature Increment (TIN5)THRTI0950 (DEG FAFENHEITJ; Used for sizing engine
during descent (No input if ATMIND = 0,2).
0951 Step Size for Cescent (DELH5)THRU0960 rFTi
0961 Range at End of Descent (RMAX5)THRU0970 rNmi; (No input when DESIND = 2,4,6,P).
0971 Minimum Altitude Durinc Descent (HMIN)THRU0980 [FT3
0981 Power Turbine Speed Ratio (AN2MS)THRU0990 Ratio of operatina power turbine speeO to
maximum power turbine speed [input forboth primary and auxiliary indeneneentengines in performance sea. 5, DESCENT)fNo input if N2IND = 0 or ]).
0991 Profile Drac Increase Durinq Descent (CLCr'DS)THRU1000 Used to simulate drag brakes.
V. LOITEr INFORMATION {SGTIND = 61
1001 Loiter Indicator (DNIRTL)THRU1010 0 = Loiter mission is used in reserve fuel
calculation (gross wt reset after loiter).
1 = Loiter mission used as nart of basicmission profile (aross weight not reset).
1011 Step Size for Loiter (DELST)THRU1020 rFRS]
1021 Atmosphere Indicator for SC-TIND = 2 (•T'MIN21THRU1030 0 = Standard atmosphere
153
1 = Non-standard atmos. User inputs singlepoint value for increment in ambient temp.above the standard day value.
2 Non-standard atmos. User inputs table of
ambient temp. ratios as a function of alt.
1031 Incremental Time for Loiter (STL)THRU1040 [HRS]
1041 Ambient Temperature Increment (TIN6)THRU1050 [DEG FARENHEIT; Used for engine sizLx.- at
LOITER conditions (No input when ATMIND0 or 2).
1051 Number of Primary Engines Shutdown (PNPSPL)THRU1060 During loiter seqment
106] Power Turbine Speed Ratio (AN2M6)THRU1070 Ratio of operatina powe- turbine spee: to
maximum power turbine speed finput Inrboth primary and auxiliary independentengines in performance segment 6, LOITER).(No input if N21ND = 0 or 1).
1071 Increase in Planfcrm Drag (DLOITR)THRU1080 During loiter segment.
1081 Wing Area Increase (RSW)THRU1090 Due to flap extension. This is the ratio
of the wina loading of the wing and flapto the winc loadino of the wino alone.
W. CHANGE IN FUEL WEIGHT (SGTIND = 71
LOCATION DATA DESCRIPTION (FORTRAN NAMEt
1101 Fuel Weight Increment (DLTAWF)THRU1110 rLBS1
1321 Incremental Time for Fuel Weiaht Chance (STFV)THRU1130 rHRSI
154
X. CHANGE IN PAYLOAD WEIGHT fSGTIND = 8}
LOCATION DATA DESCRIPTION (FORTRAN NAME)
1131 Payload Weight Increment (DELWPL)THRU1140 [LBS]
1141 Incremental Time for Payload Wt Chance (STPW)THRU1150 [HRS3
Y. TRANSFER ALTITUDE [SGTIND = 9)
LOCATION DATA DESCRIPTION (FORTRAN NAME)
1111 Final Altitude (HFIN)THRU1120 rFT]; Final alt. if HOPTIND = 0 (LOC 001A)
or max altitude if HOPTIND = 1.
Z. CHANGE FUEL OR CHANGE PAYLOAD
LOCATION DATA DESCRIPTION (FORTRAN NAME)
U151 Weight Indicator (WGTIND)
0 = Restriction on maximum airriane weiaht.Weight cannot exceed aross weiqcht.
1 = No restriction on airplane weiqht (willonly apply when runnina performancecalculations).
AA. GENERAL PERFORMANCE INFORMATION (SGTIND = 1I)
LOCATION DATA DESCRIPTION :FORTRAN NAME)
2201 Gross Weiaht Indicator (GWIND)THRU2210 1 = User inputs the incremental chance in
cross weiqht into LOC 2211.
2 = User inputs cross weiaht into LOC 2211.
2211 Increment in Gross Weicht/Gross Weiaht (GWP)THRU2220 rLBS'; For GWIND = 0, input the increment
155
in gross weight; For GWIND 1 1, input thegross weight value.
2221 General Perform. Atmosphere Indicator (ATMIN7)THRU2230 0 = Standard atmosphere
1 - Non-standard atmos. User inputs sinalepoint value for increment in ambient temr.above the standard day value.
2 = Non-standard atmos. User inputs table ofambient temp. ratios as a function of alt.
2241 Ambient Temperature Increment (TIN7)THRU2250 [DEG FARENHEIT); Used for enqine sizino at
GENERAL PERFORMANCE conditions (Wo inputwhen ATMIND = 0 or 2).
2251 Profile Drag Increase (DLCDCR)THRU"2260 [SO. FT.); Drag increase durina cruise due
to engines beina shut down (based on wingplanform area).
2261 Altitude (AHOP)THRU2270 rFT1
2271 Thrust-t-5-Weight Ratio for Takeoff (ENT')THRU2280
2281 Power Turbine Sneed Ratio (AN2M7)THRU2290 Ratio of operatinq power turbine speed to
max. power tu-bine speed f input for bothprim. and auxiliary independent enaines inperformance seament 11, TAKEOFF - GENEPALPERFORMANCE].
2291 Velocity Increment (DT.LVP)THRU2300 FKTS1
2301 Maximum Velocity (VMAXP)THRU2310 [KTS1
1 56
2311 Power Turbine Speed Ratio (AN2V8)THRU2320 Ratio of operating power turbine speed to
max. power turbine speed. (input for bothprim. and auxiliary indepetndent engines inperform. seg. 11, CRUISE - GFN PFRFORM).
BB. ENGINE CYCLE DATA, NON-STANDARD PERFORMANCE
LOCATION DATA DESCRIPTION (FORTRAN NAME)
1201 Primary Engine Fuel Flow Indicator (WDTIND)
0 = No primary encine fuel flow cutoff
1 = Primary engine fuel flow cutoff
1202 Primary Engine Nl Indicator (ANIIND)
0 = No primary engine limit for the aasgenerator shaft speed (Nl)
1 = Primary engine limit for the cas ceneratorshaft speed (NI)
1203 Primary Engine Referre6 Nl Indicator (AN3INN)
0 = No primary engine referred N' limit
I = Primary engine referred N1 limit
1204 Primary Engine N2 Indicator (ANIND)
0 = No primary encine N2 limit. Prima!'v eno.operates at optimum power turbine sneed(N2) value.
1 = Limit imposed on primary enqine N2. Ena.operates at optimum power turbine speed(N2) value.
2 = Limit imposed on primary enaine N2. Fna.operates at known value of N2 (in aeneral,a non-optimum value).
1205 Torque Limit Indicator (OIND)
S0 = No torque limit
I = Torque limit
157
1206 Reynold's No. Correction Indicator (RNOIND)
0 = No Reynold's no. corrections
1 = Reynold's no. corrections
1207 Reynoli°s Number Correctiv• Ff-tor (PRN)THRU1216 Reynold's no. correction for qas aenerator
shaft speed (No input if RNOIND = 0).
1217 Lift Eng. Fuel Flow Indicator (VWDIND)
0 = No fuel flow limit on the lift enqine.
1 = Fuel flow limit imposed on the lift eno.
1218 Lift Engine Ni Indicator (VNlIND)
0 = No limit on the lift engine gas generatorshaft speed (Ni).
I = Limit imposed on the lift engine qasgenerator shaft speed (NI).
1219 Lift 9ng. Pwr Turbine Speed Indicator (VN2INP)
0 = No limit on the lift engine power turbinespeed (N2).
1 = Limit imposed on the lift encine nowerspeed (N2).
1220 Primary Eng. Referred Fuel Flow Cutoff (WMAX)
(No input if WDTIND = 0)
1221 Primary Engine Gas Generator RPM Limit (AIMAX)
Ratio of max gas generator RPM to RPM atmaf static puwer, standard sea level (Woinput when AN1IND = 0).
1222 Prim. Eng. Referred Gas Gen. RPM Limit (A3MAX)
Simulates a rest':iction on compressionspeed (N3 input when AN3IND = 0).
1223 Primary Pna. Power Turbine Speed Limit (A2MAX)
Ratio of max power turbine speed (N2) to
158
power turbine sreed at max static power,standard sea level conditions (No inputwhen AN2IND = 0).
1224 Torque Limit (OMAX)
Ratio of max torque limit to torquedeveloped at static conditions, standardsea level.
1225 Engine Power Correction Factor (RNE)THRU1234 To account for Reynold's number effects
(No input when RNOIND = 0).
1235 Lift Eng. Referred Fuel Flow Cutoff (WLMAX)
(No input when VWDIND = 0)
1236 Lift Engine Gas Generator RPM Limit (AIMAX)
Ratio of max gas generator RPM (Ni) to RPMat max static power, standard sea level(No input when VNlIND = 0).
1237 Lift Engine Power Turbine Speed Limit (AL2MAX)
Ratio of max power turbine speed (N2) topcwer turbine speed at max static power,standard sea level (No input if VN2IND=-O).
1238 Output Shaft Speed Correction Factor fA2NO)THRU1247 Ratio of operating power turbine speed to
optimum power turbine speed (input whenN2IND = 2 and non-standard correction isdesired). See additional info in Ref. 3,page 5-39. (No input if N2IND = 0 or 1).
1248 Output Power Correction Factor (PNZ)THRU1257 Input if N2IND = 2 and non-standard
correction is desiree. Ratio of poweravailable at specified power turbine speedto power available at the optimum powerturbine speed (No input if N2IND = 0,I).
159
CC. PRIMARY ENGINE CYCLE INFORMATION
* LOCATIONS 1301 - 1565 ARE NOT REOUIRED IF A STANDARD ** PRIMARY ENGINE CYCLE IS SELECTED *
LOCATION DATA DESCRIPTION (FORTRAN NAME)
1301 Cycle Number (CYCPRL)
MUST match engine cycle no. in LOC 0201.
1302 Primary Engine Weight Factor (SK3)
[LB/HP) if ENGIND = 0; rLB/L,-THRUST1 ifENGIND = I or 2.
1303 Primary Engine Weight Factor (SX4)
rLBS)
1304 Primary Engine Dimensional Factor (X14)
rFT/LB-THRUST' if ENGIND = I r~r 2;[FT/SQRT(SHP)i if ENGIND = 0.
1305 Ground idle Turbine Inlet Temperature (TGI)
rDEG RANKINEJ
1306 Flight Idle Turbine Inlet Temperature (TFI)
[DEG RANKINE]
1307 Normal Power Turbine Inlet Temp. (TNRP)
rDEG RANi'INE1
1308 Military Power Turbine Inlet Temp. (TMIL)
rDEG RANKINEl
1309 Maximum Power Turbine Inlet Temp. (TMA-X)
rDEG RANKINE1
1310 Number of Referred Temperatures (UNTS)
Number of values in LOC 1311 - 131P.
160
1311 Referred Turbine Temperatures (TSHP)THRU1318 [DEG RANKINE]: Ratio of turbine temp. to
ambient temperature ratio.
1319 Number of Mach No. (UMS)
Number of values in LOC 1320 - 1325.
1320 Mach Number Values (AMSHP)THRU1325 Referred thrust tbl values (LOC 1326-13731
1326 Referred Thrust or Horsepower Values (SHPAV)THRU1373 (Table must be at least 3 X 3 in size)
1374 Number of Referred Temperatures (UNTWI
Number of values in LOC 1375 - 1382.
1375 Referred Turbine Temperature Values (TWD)THRU1382 [DEG RANKINE]; Ratio of turbine temp. to
ambient temperature ratio.
1383 Number of Mach Numbers (UMW)
Number of values in LOC 1384 - 1389.
1384 Mach Number Values (AMWD)THRU1389 Values for the referred fuel flow table
(LOC 1390 - 1437).
1390 Primary Enq. Referred Fuel Flow Rate (FWDOT)THRU1437 (Table must be at least 3 X 3 in size)
438 Number of Referred Temperatures (UNTI)
Number of values in LOC 1439 - 1446.
1439 Referred Turbine Temperatures (TNI)THRU1446 rDEG RANKINEI; Ratio of turbine temp. to
ambient temperature ratio.
1447 Number of Mach No. (UNMI)
Number of values in LOC 14-8 - 1453.
161
ýs
1448 Mach Number Values (AM1)THRU1453 Values for the referred gas generator RPM
limit table (LOC 1454 - 1501).
1454 Referred Gas Generator RPM Speed Limit (AONE)THRU1501 (Table must be at least 3 X 3 in size)
1502 Number of Referred Temperatures (UNT2)
Number of values in LOC 1503 - 1510.
1503 Referred Turbine Temperatures (TN2)THRU1510 [DEG RANKINE); Ratio of turbine temp. to
ambient temperature ratio.
1511 Number of Mach No. (UNM2)
Number of values in LOC 1512 - 1517.
1512 Mach Number Values (AM2)THRU1517 Values for the referred power turbine RPM"
limit table (LOC 1518 - 1565).
1518 Referred Gas Generator RPM Speed Limit (ATWO)THRU1565 (Table must be at least 3 X 3 in size)
DD. LIFT ENGINE CYCLE INFORMATION
S************* ************************************************
* LOCATIONS 1601 - 1672 ARE NOT REQUIRED IF A STANDARD LIFT ** ENGINE CYCLE IS SELECTED OR IF LFTIND = 0 ** * ************** .. *********** **** * **** ******* ***************** *
16U1 Cycle Number (CYCLFL)
MUST match lift eng. cycle no. - LOC 0218.
1602 Lift Engine Weight Factor (SK1)
[LB/LB THRUST)
1603 Lift Engine Weiqht Factor (SK2)
[LBSJ
162
1604 Lift Engine Dimensional Factor (XIll
[FT/SQRT(LB THRUST) .i
1605 Lift Engine Dimensional Factor (X12)
rFT]
1606 Lift Engine Dimensional Factor (XI3)
rFT/SORT(LB THRUST) 1
1607 Ground Idle Turbine Inlet Temperature (TLG-I'
rDEG RANKI•I%
1608 Max Power Turbire Inlet Temperature (TLMAX)
[DEG RANKINE 1
1609 Referred Turbine Temperatures (TF)THRU1616 [DEG RANiKINEI; Ratio of turbine temp. to
ambient temperature ratio.
1617 Values of Referred Thrust (FAVL)THRU1624
1625 Referred Turbine Temperatures (TFW)THRU1632 rDEG RANKINEI: Ratio of turbine temp. to
ambient temperature ratio.
1633 Values of Referred Fuel Flow Rate (FWPOT)THRU1640
1641 Referred Turbine Temperatures (TFI)THRU1648 FPEG RANKINE1; Ratio of turbinp temp. to
ambient temperature ratio.
1649 Referred Gas Generator Speed Limit (FON'F)THRU1656
16 5 7 :'efe.-red "urbine Temperatures (TF?)THRU1664 [DEG RANKINE,; Patio of turbine temp. to
ambient temperature ratio.
163
1665 Referred Power Turbine Speed Limit (FTWO)THRU1672
EE. ECONOMILS
LOCATION DATA DESCRIPTION (FORTRAN NAME)
1675 Inflation Factor (SPAC15(3))
Base year is 1967.
1676 Profit Factor (SPAC15(4))
Expressed as a no. creacer than 1.0. Forex., 1.1 would represent a 10% profit.
1677 No. of Prototype Aircraft (SPAC15(5)!
Number in development program.
1678 No. of Prod' :tion Aircraft (SPAC1S(6))
1679 Avionics and Miscellaneous Costs (SPAC15(7))
Per prototype aircraft.
1680 No. of Ground Test Articlep (SPACl5(8))
1681 Nc. of Flicht Test Hours (SPAC15(0))
1682 Trainer and Misc. RDT&E Costs (SPAC15(10))
1683 Avionics Costs (SPAC15(ll)f
Per production aircraft in 1067 aollars.
1684 Cost of Fuel (SPACI5(12))
rDOLLARS/LBI1C85 Cost of Oil (SPAC]5(13))
rDOLLARS/LB]
168f Hull Insurance Rate (SPAC15(14))
1687 Maintenance Labor Rate (SPAC15(15))
r POLLARS/HRI
1 64
1688 Time Between Engine Overhauls (SPAC!5(16))
CHRS?
1689 Time Between Dynamic Systems Overhauls(SPACl5(17))
rHRS1
1690 Dynamic Systems Indicator (SPAC15(181)
1 = Dynamic system used durina entire fliaht.
2 = Dynamic system used during talzeoff/laneinaonly.
1691 Annual Interest Rate on Capital (SPACl5(19))
1692 Depreciation Period (SPACI5(20))
rYRS]
1693 RE-sidual Value (SPAC15(21))
rDOLLARS1
1694 Annual Utilization (SPAC15(72))
rHRS1
1695 Customer Aircraft Buy (SPACI5(23))
FF. HOVER PERFORMANCE MAP
LOCATION DATA DESCRIPTION (FORTRAN NAME)
2351 Number of Thrust Coefficient to PropellerSolidity Ratios (CTSGNO)
Number of values in LOC 2352 - 2361.
2352 Thrust Coeff. to Solid. Ratio Values (C"'OSI)THRU2361 (Input at least three (3) values)
2362 Number of Tip Mach N-imber Values (TPMNO)
Number of values in LOC 23A3 - 236P.
1A5
2363 lip Mech Number Values (TIPM)THRU2368 (Input at least three 13) values)
2369 Figure of Merit Table (FMER)THRU2428 Figure of merit values as a function of
thrust coefficient to prop soliRitv ratiosand tip Mach numbers.
GG. PROPELLER/FAN PERFORMANCE DATA
* LOCATIONS 1700 - 2142 ARE REOUIRED WHEN ETAIND 1 or 3 *
1700 Propeller/Fan Tahle Number (PROPCY)
MUST match value for CYPROP (LOC 0256).
-701 Number of Advance Ratios or Mach No. (XPXNO)
Number of values in LOC 1702 - 1721
1702 Prop Advance Ratios or Mach No. Values (XPJ)THRU1721 (Input at least three (3) values of prop
advance ratio or Mach number)
1722 Number of Propeller Thrust Coefficients orReferred Thrust Coefficients (CPPNO)
Number of values in LOC 1723 - 1742.
1723 Propeller Thrust Coefficients or PeferreOTHRU Thrust Coefficients (CTPROP)1742
(Input at least three (3) values' of propthrust coeff. or referred thrust copff.)
1743 Propeller or Fan Power Coefficients (CPPROP)THRU2142 Prop Power Coeff: Array input as a func.
of advance ratio and prop thrust coeff.
Fan Power Coefficients: Array irput as afunction of Mach number and referredthrust ccefficient.
I•6
HH. DESCRIPTION OF SAMPLE DATA VALUES
The following is an actual data file used to study an
eight passenger tilt rotor aircraft design. Each data va,.ue
used in the input data file is listed below ane •- • bed.
II. LISTING OF DATA LOCATION/VALUES
LOC VARIABLE VALUE SIGNIFICANCE OF DATA VALUE
0001 OPTIND 1.0 Sizing run
0002 TNIRPK 0.0 Standard output
0003 DRGIND 0.0 Procram calculates drag rise eueto compressibility effects
0004 OSWIND 1.0 Program calculates the OswalA'sefficiency factor
0005 PDMIND 3.0 Input diameter and thrust coeff.to soliditv ratio
0006 FDMIND 2.0 Input desired seatina capacitv,seat width and pitch, numher anewidth of aisles, number of seatsabreast for tourist and firstclass, galley and lavatory size;Program calculates fuselaae size
0007 WDMIND 0.0 Input winq loadina & aspect ratio
0008 HTIND 2.0 Input horizontal tail area
0009 VTIND 2.0 Input vertical tail area
0010 FIXIND 0.0 Input level of maximum power orthrust (fixed encine si7e)
0011 ENGIND 0.0 Turboshaft encine
0072 ESZIND 0.0 Encines sizeO for takeoff only
0013 LFTIND 0.0 No separate lift propulsionengine
0014 WGOO 13000.0 First auess at cross weial: rLRSI
167
0015 HOO 0.0 Start altitude [FT]
0016 ROO 0.0 Start range iN'Mll
0017 STOO 0.0 Start time [HPS]
0018 HOPTIN 0.0 Input desired cruise segment alt.
0019 VLMIND 0.0 Airspeed limited to 250 kts EASat altitudes of 10,000 ft or less
0020 EMMO 0.575 Max operating Mach number
0021 VMO 260.0 Max operatina equivalent airspeedrKTS]
0022 VDIV 300.0 Design dive speed rKTSS
0023 EMLF 4.0 Maneuver load factor
0024 CK1 Default = 1.0 (no reserve fuel)
0025 DELWF Default = 0.0 (no fixed fuel forreserves or other use)
0026 CKFF Default = 1.0 (use nominal enginefuel)
0027 SGTIND 1.0 Taxi
0028 SGTIND 2.0 Takeoff
0029 SGTIND 3.0 Climb
0030 SGTIND 4.0 Cruise
0031 SGTIND 5.0 Descent
0032 SGTIND 2.0 Land
0033 SGTIND 1.0 Taxi
0034 SGTIND 9.0 Transfer altitude
0035 SGTIND 6.0 Loiter
0036 SGTIND 100.0 End of case
0037 SGTINDTHRU Not used0076 SGTIND
168
0077THRU Not assigned for program0093
0094 SPACE1(18) 1.0 High wing location for 3-View
drawing
0095 SPACEI(19) Default = 0.0
0096 SPACE1(20) Default = 0.0
0097 SPACE1(21) Default = 0.0
0098 SPACE1(22) 0.0 Any value greater than 0.0 willgenerate the 3-View drawing: NOTavailable at NPS as of thiswriting
0099 Not assignel for program
0100 1.0 NPS modification; output will teabbreviated and a maximum of 80characters wide for compatibilitywith THESIS2.
0101 DAM2 6.6 Wing aspect ratio
0102 DAM3 WDMIND = 0 (LOC 0007), thereforeno input
0103 EYEW 3.0 Wing incidence angle CrEGrlmeasured with respect to fuse]age
0104 TCR 0.223 Winc root thickness-chord ratio
0105 TCT 0.223 Wina tip thickness-chord ratio
0106 DAM4 72.45 Wino loadina FLBS/SO FTI atdesign gross weight
0107 DLMC4 -6.5 Quarter chord mean sweep anglerDEG): (The XV-15 has a forwardswept wino for prop clearancedurina flapping)
0108 SLM 1.0 Taper ritio of winq (tip chord/
root chord)
0109 ARHT 3.27 Horizontal, tail aspect rato
0110 SAH 0.0 Horizontal tail is on vertical
169
"tail root chord
0111 ELTH 22.4 Horizontal tail moment arm [FT)
0112 TLCT 0.15 Horizontal tail mean thickness tochord ratio
0113 VBARH HTIND = 2 (LOC 0008), thereforeno input
0114 SLMH 1.0 Horizontal tail taper ratio
0115 AAW11 50.25 Horizontal tail planform arearso FT]
0116 SR .08 Prop blaee attachment distancemeasured from the centerline ofthe hub and expressed as a"fraction of the propeller radius
0117 ICL WDMIND = 0 (LOC 0007), therefore* no input
0118 ZETAl WDMIND = 0 (LOC 0007), thereforeno input
"0119 ZETA2 WDMIND = 0 (LOC 0007), thereforeno input
0120 DLSWSW Default = 0.0 (no protrusionssuch as landing gear)
0121 HF FDMIND = 2 (LOC 0006), thereforeno input
0122 DAM5 FDMIND = 2 (LOC 0006), thereforeno input
0123 ELPD 1.2 Nose section fineness ratio
0124 ELTD 2.5 Tail section fineness ratio
0125 ELC 18.8 Cabin section length of constantdiameter rFT]
0126 ELRW 0.0 Length of ramp well [FT)
0127 DAM6 FDMIND = 2 (LOC 0006), thereforeno input
0128 SWF FDMIND = 2.0 (LOC 0006); no input
170
0129 1 T 2.33 Vertical tail aspect ratio
0130 ELTV 23.2 Vertical tail moment arm [FT?
0131 TCVT .09 Vertical tail mean thickness tochord ratio
0132 VBARV VTIND = 2 (LOC 0009), thereforeno input
0133 SLMV 0.587 Vertical tail taper ratio
0134 AAW12 50.5 Vertical tail area [So FT]
0135 YMG 0.0 Position of main landing gearmeasured outboard from the sideof the body and expressed as afraction of wing semi-span
"0136 YP 1.0 Mean position of primary enginesmeasured outboard from airplanecenterline and expressed as afraction of wing semi-span
0137 YL LFTIND = 0 (LOC 0013), thereforeno input
0138 EPSLON LFTIND = 0 (LOC 0013), thereforeno input
0139 AZETA1 0.0758 Primary engine nacelle dimensionfactor
0140 AZETA2 0.0 Primary engine nacelle dimensionfactor
0141 AZETA3 0.233 Primary engine nacelle dimensionfactor
0142 SKIP(l) 0.25 Rotor thickness to chord rati" at0.25R
0143THRU Not assigned in program0150
0151 DNIlGN 0.0 Galley area calculated by program
""0152 AGLLEY 0.0 Area of galley (e.g. no galley)
"0153 ANPXI 0.0 No first class seats
171
a.
0154 ANABI Default = 0
0155 ANISLI Default = 0
0156 WSEAT1 Default = 0
0157 PSEATI Default = 0
0158 WAISLI Default = 0
0159 DNIVAL 1.0 User inputs number of lavatories
0160 ANLAVS 1.0 Number of lavatories = 1
0161 ANPXT 8.0 Passenger capacity in the touristsection
0162 ANAAT 2.0 No. of seats abreast in tourist
63 ANISLT 1.0 No. of aisles in tourist
0164 WSEATT 21.0 Width of seats in tourist FIN?
0165 PSEATT 41.0 Seat pitch [IN]
0166 WAISLT 17. Aisle width [IN]
0167 SPACE2(1) 0.0 No. in flight deck crew
0168 SPACE2(2) 0.0 No of flight attendants
0169THRU Not assigned in program01990200 ETAIND 2.0 Propulsive performance calculated
by program
0201 CYCPRP 1.78 Primary engine cycle number
0202 DAM7 2920.0 Primary engine maximum statichorsepower [HP]; total for allengines
0203 Not used for turboshaft engines
0204 ENP 2.0 Number of primary enqines
0205 Not used for turboshaft engines
0206 ETAT 0.95 Transmission efficiency
172
0207 HES 0.0 Takeoff altitude [FT]
0208 SENE 1.098 Thrust-to-weight ratio
0209 TINY 0.0 Ambient temperature increment fortakeoff [DEG FARENHEIT]
0210 AN2TO 1.0 Ratio of operating power turbinespeed to max power turbine speed
0211 ENPO FIXIND = 0 (LOC 0010), thereforeno input
0212 ENLO FIXIND = 0 (LOC 0010), thereforeno input
0213 POWESI FIXIND = 0 (LOC 0010), thereforeno input
0214 HC FIXIND = 0 (LOC 0010), thereforeno input
0215 VC FIXIND = 0 (LOC 0010), thereforeno input
0216 ATMIY FIXIND = 0 (LOC 0010), thereforeno input
0217 AN2CR FIXIND = 0 (LOC 0010), thereforeno input
0218THRU Not used for turboshaft encines0221
0222 Mot assigned for program
0223 ENR 2.0 Number oL propellers
0224 VT 817.7 Propeller tip speed rFT/SEC?
0225 WGA PDMIND = 3 (LOC 0005;, thereforeno input
0226 DI 25.0 Propeller diameter rFTJ
0227 CTSIG 0.123 Thrust coefficient to propellersolidity ratio
C228 AF 72.92 Activity factor per blade
0229 BLDN 3.0 Number of blades on propeller
173
0230 CLEYE 0.25 Propeller integrated design liftcoefficient
0231 £TAIND = 2 MLOC 0200), thereforeTHRU no input0233
0234 ETAP5 0.75 Descent propeller efficiency0235 eTAIND = 2 (LOC 0200), thereforeTHRU no input0256
0257 XMSND 0.0 Transmission sized at fraction ofinstalled power
0258 XMSMRT 1.0 Fraction of installed power forsizing transmission
0259 DSHPAC 15.0 Accessory horsepower rHPI
0260 SHPTO Ratio of eng. SHP to primary eng.max static HP (LOC 0202); usedfor sizing eng; default = 1.0
0261. VRCRC Takeofr vertical rate of climbiFT/MI9]
0262 CKRC Climb pwr multiplicative constant(default is 2.0)
0263. THRU Not assigned in program
0300
0301 CDVTTTHRU Default 0.00304 CDLUI
0305 DELCD 0.0139 Profile draq increment based onwing planform area
0306 DAM10 OSWIND = I (LOC 0004), thereforeno input
0307 DELFE 9.08 Eauuv. flat plate area [SO FTI
0308 TLLN 2.0 Number of pairs of values in thetable of wine profile 0raccoeff. (LOC 0335 - 0342) versuslift coeff. MLoC 0317 - 0324)
174
0309 TENN DRGIND = 0.0 (LOC 0003); no input
0310 TCLZN DRGIND = 0.0 (LOC 0003); no input
0311 CKLN Default = 0.0
0312 CKW 1.0 Wing multiplicative drag factor
0313 CKN 1.0 Primary nacelle multiplicativedrag factor
0314THRO Default = 0.00316
0317 TBCLI(1) 0.0 Wing lift coeff. value
0318 TBCL1(2) 4.0 Wing lift coeff. value
0319THRU Not used0324
0325THRU DRGIND = 0.0 (LOC 0003); no input0329
0330 RECI 0.2E+07 Mean Reynold's number per foot
for mission
0331 CSALF 6.28 Two dimensional lift coeff. slope
0332 ALPHL -1.0 Zero lift angle of attack [DEG]
0333 XCPS 0.3 Nondimensional position alonq thechord
0334 XCTCM 0.35 Nondimensional position along the
chord at maximum t/c ratio
0335 TBCDWI(l) 0.0 Wing profile drag coeff. value
0336 TBCDWI(2) 0.0 Wing profile drag coeff. value
0337THRU Not used0342
0343THRU DRGIND = 0.0 (LOC 0003); no input0384
175
0385THRU Not used0393
0394 SPACE4(10) 4.0 No. of stages in main rotor drive
0395 SPACE4(11) Blade fold penalty: default - 1.0(no fold)
0396 SPACE4(12) 61.0 Hub weight coefficient
0397 SPACE4(13) 0.12 Hub material/development factor
0398 SPACE4(14) 44.0 Blade weight coefficient
0399 SPACE4(15) 2.2 Hingeless rotor system
0400 OWEl OPTIND = 1 (LOC 0001); no input
0401 WFE 2477.0 Weight of fixed equipment [LBS]
0402 WFUL 716.0 Weight of fixed useful load rLBSJ
0403 WPL 0.0 Weight of payload [LBS]
0404 SKCC 15.67 Cockpit controls constant
0405 SKFW .016 Fixed-wing controls constant
0406 SKH 0.0 System and hydraulics constant
0407 SKSAS 165.0 Factor for stability augmentedsystem
0408 SKTM 0.0162 Tiltinq mechanism constant
0409 SKUC 0.779 Upper control mechanisms constant
0410 CK15THRU Default = 1.00415 CK20
0416 THN ATMIND is not set equal to 2.0during any segment; no input
0417 DELWFZTHRU Default = 0.00419 DELWST
0420 SKP 162.0 Body group weight adjustmentfactor
176
0421 SKLES LFTIND = 0.0 (LOC 0013); no input
0422 SKLG 0.04 Alighting gear weight expressed
0423 SKMG 0.80 Ratio of main gear weight togross weight
0424 SKTL 1.0 Tail load adjustment factor
0425 SKWF 0.6 Wing bending relief momentadjustment factor
0426 SKWW 350.0 Wing type weight adjustmentfactor
0427 SKY 0.195 Pitch radius of gyration EFT]
0428 SKZ 0.13 Yaw :adius of gyration [FTI
0429 SKPES 0.3422 Primary engine section weight0421 SKLES LFTIND = 0.0 (LOC 0013); no input
0422 SKLG 0.04 Alighting gear weight expressedas a percentage of gross weight
0423 SKMG 0.80 Ratio of main gear weight togross weight
0424 SKTL 1.0 Taii load adjustment factor
0425 SKAF 0.6 Wing bending relief momentadjustment factor
0426 SKWW 350.0 Wing type weight adjustmentfactor
0427 SKY 0.195 Pitch radius of gyration rFTI
0428 SKZ 0.13 Yaw radius of gyration rFTI
0429 SKPES 0.3422 Primary engine section weiqht0453 SKDS 345.0 Drive system weiaht adjustment
factor
0454 SKFS 0.10 Fuel sys. wt. adjustmer't factor
0A55 SKLEI LFTIND = 0.0 (LOC 0013); no input
0456 SKPEI 0.1654 Primary eng. installation weiqh'tfactor
177
0457 SKRP 15.77 Prop. weight adjustment factor
0458 SKVT 1.00 Drive system weigh' variationadjustment factor
0459THRU Default = 1.00465
0466 TBTHrTHRU THW = 0.0 (LOC 0416); no input0475 TBTHE
0476THRU Default = 0.00500
0501 ATMINI 0.0 Standard atmosphere for firsttax, segment
0502 ATMIN! 0.0 Standard atmosphere for secondtaxi segment
0503THRU Not usedJ5so
0511 DELTT 0.025 Incremental time for first taxisegment [HPSI
0512 DELTT 0.025 Incremental time for second taxisegment [HRS]
0513THPtJ Mot used0520
0521 TTN1 ATMINl = 0.0 (LOC 0501 - 0510);THRU no input0530 "1,'N1
053.1THRU LFTIND = 0.0 (LOC 0013); no input0540
0541 ANIM1 0.81 Ratio cf operatina Dower turbinespeed to max power turbine speedfor first taxi segment
178
L"0542 ANIMI 0.81 Ratio of operating power turbinespeed to max power turbine speed
for second taxi segment
0542THRU Not used0550
0551TIRU Not assigned in program0600
0601 TOLIND 3.0 User inL.ts required fraction ofmaximum power for takeoff segment
0602 TOLIND 3.0 User inputs required fraction ofmaximum power for '.anding segment
0603THRU Not used0610
0611 ATMIN2 0.0 Standard atmosphere for takeoffsegment
0612 ATMIN2 0.0 Fandard atmosphere for landingsegment
0613THRU Not used0620
0621 PFET2 1.0 Primary engine power (or thrust)factor for takeoff segment
0622 PFET2 1.0 Primary enaine power (or Thrust)factor for landing seament
0# 23THRU Not used0630
0631 ATMIN2 = 0.0 'LOC 061] - 0620);THRU no input0640
0641THRU Default = 0.00660
179
0661 DELTH 0.01667 Step size for hover duringtakeoff segment rHRSj
0662 DELTH 0.01667 Step size for hover durircylanding segment rHRS]
0663THRU Not used0670
0671 AN2M2 1.0 Ratio of operating power turbinespeed to max power turbine speedfor takeoff segment
0672 AN2M2 1.0 Ratio of operating power turbinespeed to max power turbine speedfor landing segment
0673THRU Not used0680
0681 STH 0.01667 Incremental time for hover duringtakeoff segment [HRS3
0682 STH 0.01667 Incremental time for hover durinqlanoino segment rHRSI
0683THRU Not used0690
0691 CLMIND 1.0 Maximum rate of climb durinoclimb segment
0692THRU Not used0700
070iTHRU CLMIND = 1.0 (LOC 0691); no innut0710
0711 ATMIN3 0.0 Standard atmosphere for climbsegment
0712THRU Not used0720
0720 DELH3 1000.0 Step size for climb secment FFT'
iPO
0721THRU Not used0730
0731THRU ATMIND = 0.0 (LOC 0711); no input0740
0741 HMAX 15000.0 Max altitude for climb segment[FT]
0742THRU Not used0750
0751 POWCLI 1.0 Climb at military power ratini
0752THRU Not used0760
0761 THEMAX 20.0 Max body attitude anqle for climbsegment [DEG]
0762THRU Not used07?0
0771 AN2M3 0.81 Ratio of operating power turbinespeed to max power turbine speedfor climb segment
0772THRU Not used0780
0781THRU Default = 0,00800
0801 CRSIND 1.0 Cruise nt cruise power
0802THRU Not used0810
0811THRU Default =0.00820
N87
0821 ATMIN4 0.0 Standard atmosphere for cruisesegment
0822THRU Not used0830
0831 DELR 20.0 Step size for cruise segment [NM2
0832THRU Not used0840
0841THRU ATMIN4 = 0.0 (LOC 0821); no input0850
0851 RMAX 300.0 Range at end of cruise seament
0861 POWCRU 2.0 Cruise at normal power rating
0862THRU Not used0870
0871THRU Default = 0.00880
0881 AN2M4 0.81 Ratio of operatin- .,ower turbirespeed to max powou- turbine speedfor cruise seqment
0882THRU Not used0890
0891THRU PefavIlt = 0.00900
09CI VESINE 1.0 Descend at maximum speed at tbeterminal rance specified
002THRU Not used
0911THRU VESIND 1.0 (LOC 0901); no input0920
182
i.=-7
0921 ATMIN5 0.0 Standard atmosphere for thedescent segment
0922THRU Not used0930
0931 THEMIN -20.0 Minimum body attitude angle
during descent segment [DEG]
0941THRU ATMINS - 0.0 (LOC 0921): no input0950
0951 DELH5 1000.0 Step size for descent segment""N FT_
0952THRU Not used0960
0961 RFAX5 300.0 Rance at end of eescent rw?-
0962THRU Not used0970
0971 HMIN 0.0 Minimum altitude durinq descentsegment rFT!
0972THRU Not used0980
0981 AN2M5 0.*P Ratio of operating power turbinespeed to max power turbine speedduring descent seament
0982
THRU Not used0990
0991THRU Default = 0.01000
1001 LTRIND 2.0 Loiter performeA for reserve fuel
calcul ation
1_P.3
*.* .%* C... %%
1002THRU Not used for this case1010
1011 DELST 0.1 Step size for loiter segrent [HRJ
1012THRU Not used for this case1020
1021 ATMIN6 0.0 Standat'd atmosphere
1022THRU Not used for this case1030
1031 STL 0.3333 Incremental time for loiter rHRS1
1032THRU Not used for this case1060
1061 AN2M6 0.81 Ratio of oreratinq power turbinespeed to maximum power turbinespeed during loiter seament
1062THRU Not used for thii case1110
1111 HFIN 19000.0 Final altitude for transferaltitude segment [FT1
1112THRU Not used for this case1151
1152THRU Not assigned for prcqram1200
1201 WD'TIND 1.0 Restriction will be applieO tcfuel flow
1202 A'IND 0.0 No restruction on referred N1limit
1204 AN2 T ND 2.0 Restriction wili 1e applied toN2; Engine will operate at aknown value of N2 (generallynon-optirum)
184
1205 QIND 0.0 No restruction on torque limit
1206THRU Not used for this ca.:e1219
1220 WMAX 1.11 Maximum fuel flow will be 1.1%greater than fuel flow at maximumstatic thrust, standard sea level
1221 AIMAX 1.04 Gas generator cutoff at 4% overmax. sea level gas generator RPM
1222 A31MAX 0.0 No cutoff of referred N1
1223 A2MAX 0.905 Power turbine cutoff at q0.5% ofmaximum sea level turbine power
1224 QMAX 1.446 Torque cutoff at 44.6* over max.torque developed at sea level,static, standard day conditions
1225THRU Not used for this case1257
1258THRTU Not assigned for program1300
1301 CYCFRL 1.78 Propulsion cycle number
1302 SK3 0.36 Primary engine weiahtMultiplicative factor
1303 SK4 0.00 Primary engine weiahts additionalfactor
1304 X!4 0.03? Primary engine dimensional factor
1305 TGI 1650.0 Turbine inlet temperature, aroundidle power settinq DFEG RANFINFI
1306 TFI 1P.00.0 Turbine inlet temrerature, flichtidle power setting
1307 TNRP 2180.0 Turbine inlet temperature, norralpower settino
1308 TMIL 2235.0 Turbine inlet temperature,military power settinq
I PS
1309 TMAX 2280.0 Turbine inlet temperature,
maximum power setting
1310 UNTS 7.0 Number of referred temperaturesin locations 1311-1318
1311 TSHP 1600.0 Values of rsferred turbine temp.1312 " 1800.0 for the referred thrust or H.P.1313 " 2000.0 tables1314 " 2200.01315 " 2400.03316 " 2600.01317 2800.0
1318 Not used for this case
1319 UMS 5.0 Number of Mach number values inlocations 1320-1325
1320 AMSHP 0.0 Values of Mach number for the1321 0.2 referred thrust cr H.P. tab]es1322 0.41323 0.61324 0.8
1325 Not used for this case
1326 SHPAV 0.035 Values of referred thrust or H.P.1327 0.075 corresponding to referred1328 0.125 temperature location 1:11 and1329 0.180 Mach numbers fo-,,,* on locations1330 0.240 1320-1324
1331 Not used for this case
1332 SHPAV 0.330 Values of referred thrust or H.P.1333 0.375 corresponding to referred terp.1334 " 0.425 location 1312 and Mach numb-ers1335 "0.410 found in locations 1320 - 13241336 0.534
1337 Yot used for this case
1338 SHPAV 0.630 Values of referred thrust or H.P.1339 0.670 correspondino to referred tenp.1340 0.72C location 1313 and Mach numbers1341 0.775 found in locations 1320 - 13241342 0.835
1343 Not used for this case
186
' .S , . • - - .. . . -. - - - • - , : . j , . . '
1344 SHPAV 0.920 Values of referred thrust or H.P.1345 " 0.960 corresponding to referred1346 " 1.010 temperature location 1314 and1347 t 1.065 Mach numbers found in locations1348 " 1.125 1320-1324
1349 Not used for this case
1350 SHPAV 1.200 Values of referred thrust or H.P.1351 1.245 corresponding to referred temp.1352 1.295 location 1316 and Mach numbers1353 1.350 found in locations 1320 - 13241354 1.410
1355 Not used for this case
1356 SHPAV 1.340 Values of referred thrust or H.P.1357 i 1.390 corresponding to referred temp.1358 1.440 location 1316 and Mach numbers1359 1.495 found in locations 1320 - 13241360 1.550
1361 Not used 'or this case
1362 SHPAV 1.400 Values of referred thrust or H.P.1363 1.450 corresponding to referree tern.1364 1.500 location 1317 and -.;h numbers1365 1.550 found in cation: 1320 - 13241366 K 1.600
1367THRU Not used for this case1373
1374 UNTW 7.0 Number of referreO tkzperaturesin LOC 1375-1382
1375 TWD 1600.0 Values of referreO turbine temp.1376 "o00.0 f:r thr r,ý.ferreA 'e 1 frew1377 " 2000.0 talle1378 " 2200.01379 2400.01380 " 2600.01381 2800.0
1382 Not used for this case
1383 UMW 5.0 Number of Mach number values inlocations 13P4-1389
187
1384 AMWD 0.0 Values of Mach number for the1385 0.2 referred fuel flow table1386 0.41387 0.61388 0.8
1389 Not used for this case
1390 WDOT 0.150 Values of referred fuel flow1391 0 0.170 corresponding to the referred1392 " 0.15C temperature location 1375 and1393 " 0.150 Mach numbers found in locations1394 " 0.150 1384-1388
1395 Not used for this case
1396 WDOT 0.277 Values of referred fuel flow1397 I 0.277 corresponding to referred temp.1398 " 0.277 location 1376 and Mach numbers1399 0.277 found in locations 1384 - 1388A400 0.277
1401 NTot used for this case
1402 WDOT 0.407 Values of referred fuel flow1403 0.407 corresponding to referred temp.1404 0.407 location 1377 and Mach numbers1405 0.407 found in locations 1394 - 13PS1406 0.407
1407 Not used for this case
1408 WDOT 0.535 Values of referred fuel flow1409 " 0.535 correspcndinq to referred1410 " 0.535 temperature 1 ocation 1378 and1411 o 0.535 Mach numbers found in locations1412 t 0.535 1384-13P8
1413 Not usec! for this casp
1414 WDOT 0.662 Values of referred fue' flow1415 0.662 corresponding to referred tern.1416 0.662 location 1379 and Mac1, numbers1417 0.662 found in locations 1384 - 138S1418 0.66?
1419 Not used for this case
1420 WDOT 0.750 Values of refeLreO fuel flow1421 t 0.750 correspondina to referred temp.1422 0.750 location 13P0 and Mach numbers
188
-. . . .. . . . . . . . . . ~ ~ ~ ~ .• . ° . . . . . . .° .- 0 . -. . - ° - - . . . . . . . ° -
1423 " 0.750 found in locations 1384 - 1388
1424 0. 1)
1425 Not used for this case
1426 WDOT 0.802 Values of referred fuel flow1427 " 0.802 corresponding to referred temp.1428 " 0.802 location 1381 and Mach numbers1429 0.802 found in locations 1384 - 13881430 0.802
1431THRU Not used for this case1437
1438 UNTI 7.0 Nimber of referred temperaturesin LOC 1439-1446
1439 TN! 1600 0 Values of referred turbine temp.1440 " 1800.0 for the referred gas generator1441 " 2000.0 RPM limit table1442 " 2200.01443 * 2400.01444 2600.01445 " 2800.0
1446 Not used for this case
1447 UNMI 5.0 Number of Mach number values inlocations 144P-i453
1448 AM1 0.0 "alues of Mach number for the1449 o 0.2 referred aas generator RPM limit1450 0.4 table1451 0.61452 0.8
1453 Not used for this case
1454 AONE 0.722 Referred aas cenerator RPM limit1455 " 0.735 values correspondina to referree1456 0.748 temp. location 1439 and Mach no.1457 0.766 locations 1448-14S21458 0.789
1459 Not used for this case
1460 AONE 0.840 Values of referred fuel flow1461 0.846 correspondina to referred temp.1462 0.853 location 1376 and Mach numbers1463 0.860 foune in locations 1384 - 13PP
189
146A i 0.871
1465 Not used for this case
1466 AONE 0.925 Referred gas generator RPM limit1467 " 0.927 values corresponding to referred1468 " 0.933 temp. location 1441 and Mach no.1469 I 0.939 locations 1448-14521470 " 0.950
1471 Not used for this case
1472 AONE 0.990 Referred gas generator RPM limit1473 " 0.992 values correspondinq to referred1474 " 0,997 temp. location 14412 and Mach no.1475 " 1.004 locations 1448-14521476 " 1.015
1477 Not used for this case
1478 AONE 1.045 Referred gas generator RPM limit1479 " 1.048 values corresponding to referred1480 i 1.052 temp. location 1443 and Mach no.1481 1.059 locations 1448-14521482 1.068
1483 Not used for this case
1484 AONE 1.097 Referred gas qenerator RPM livit1485 " LC29 values correspondina to referred1486 " 1.105 temp. location 1444 and Mach no.1487 o 1.059 locations 1448-14521488 i 1.068
1489 Not used for this case
1490 AONE 1.150 Referred qas generator PPP limit1491 1 1.154 values corresponding to referred1492 1.158 temp. location 1445 and Mach no.1493 1.111 locations 1448-14521494 1.119
1495THRU Not used for this case1501
1502 UNT2 7.0 Numrer of referred temperaturesin locations 1503-1510
1503 TN2 i600.0 Values of referred temperature1504 " IOO.0 for the referred power turbine
190
1505 " 2000.0 speed limit ratio table1506 " 2200.01507 " 2400.01508 2600.01509 " 2800.0
1510 Not used for this case
1511 UNM2 5.0 Number of Mach number values inlocations 1512-1517
1512 AM2 0.0 Values of Mach number for the1513 " 0.2 referred power turbine speed1514 " 0.4 limit ratio table1515 0.61516 " 0.8
1517 Not used for this case
1518 ATWO 0.445 Values of referred power turbineS1519 0.461 speed limit corresponding to
1520 0.500 referred temp. location 1503 and1521 0.557 Mach no. locations 1512-15161522 0.640
1523 Not used for this case
1524 ATWO 0.685 Values of referred power turbine1525 " 0.699 speed limit corresponding to1526 1° 0.734 referred temp. location 1504 and1527 " 0.789 Mach no. locations 1512-15161528 " 0.858
1529 Not used for this case
1530 ATWO 0.856 Values of referred power turl-ine1531 0.880 speed limit correspondinc to1532 0.908 referred temp. location 1505 and1533 0.940 Mach no. locations 1512-15161534 0.973
1535 Not used for this case
1536 ATWO 0.983 Values of referred power turbine1537 " 0.997 speed limit corresponding to1538 1.009 referred temp. location 1507 and1539 1.023 Mach no. locations 1512-15161540C 1.029
1541 Not used for this case
191
1542 ATWO 1.084 Values of referred power turbine1543 " 1.088 speed limit corresponding to1544 " 1.089 referred temp. location 1507 and1545 o 1.086 Mach no. locations 1512-15161546 " 1.076
1547 Not used for this case
1548 ATWO 1.178 Values of referred power turbine1549 o 1.169 speed limit corresponding to1550 1.158 referred temp. location 1508 and1551 1.145 Mach no. locations 1512-15161552 1.123
1553 Not used for this case
1554 ATWO 1.264 Values of referred power turbine1555 I 1.246 speed limit corresponding to1556 " 1.224 referred temp. location 1509 and1557 1.197 Mach no. locations 1512-15161558 1.161
1559 Not used for this case
JJ. EXAMPLE DATA FILE
The remaining pages of this chapter are a presentation of
the actual data file that was submitted on the MVS batch
network at the NPS computer center for the tilt rotor
aircraft Oescribed by the data 'bove.
The user will note that the data above was kept in a
sequential order for ease of reading. The data file,
however, does not need to follow any sequence as can be seen
in the sample file on pages 193-196.
192
00 0 (14
00
N0 z
11 (fM0 0 n 0 incn )4 0 -.0C,4 C) zEn) 0.9C 00000 .0 .. 0. Wi 0 0
H r10 4 r-CI 0 C%1r' E-4 0 cr)
aD 1% 1I..'
C4 Iit M cn r40411 C~4 " 0
C. U04 &.3Qz "MEU I1 0 cn c. cr, N z
0" 4 E-4IEN .DDC CIhcn W 0 0O'bO 0%0010 m0 00000 C1400c~ U)00W . ZNl 0 . .- "
CP ri¶J 110a4 Oe,400c') OOOHOO00 WOW"~1 "CI 'I N E* -I
00 w03:M>i0-~ 0 0
0 -iu ,m4
U).en0 0 '4U)* 004 ()H E-4 .. -0C.
rý3 U 0 0~cu -~a' U)O . . .Z . Z 0(tNNO O 003 .04 2CY)C4 .t-444W) 00000 0 NN An . .0 Cil0C .0 "r-i Or,-i-i 0 <000CAM U)(n 111: n S . . . . to . .0 ano0. 1%..v-4 . En. a-I.O~NO Cn)>4EEIE '.E000CeO M) 0.NNO -4NOC OH~,- cro - 4,,
4G1U2:t~)D' 11 4Z<Mr)-4~- 11 IOO0. H ix) E- 04LO rjlZZmzUtit4 I 0W
01-<M4>4 m, M 4-i>i~U) 9 <4
11 C40~ Ex Q z - r -
MCOOO C0 wi ix 0 Cx. wx 00ZP4QOQ~tr-1 W z <aD 0,-)[1 0 (0 NIn wr %MC 0
*0 0 1. NO u m3N"U) 0 L0u00 Ui-IQ) 0cr)W '-4.-du.-4 040000") C;OO~r'cv~nn U)0000 Z00a'0 r-O jZO0000OOCw0m 0C~ .O"....!. .1 .. . . n.. " -...0. X'-4 Oorio-lO 0oco,0Nrt-.r-lv 0r, 040 -rs~-NON0N<0u'-40oc'i
-<ZWOi- CxWxE-IC.ZJ rN N- N N*M-4 C4 k-O-10rCxLn OM"Mj'L'J- H ~ ~ t'~r)-N cv),'V1'eN- UWrInc4n-icr-~ r-i%4C(4C4r-I*C-0< 04W 00O Nor N- N- N-
<2:0EH HHO-4E-ic-- r-iaor-Iitnor r-icwt)Cwio1N ~-qmcnrr,- 0~o",ckrr)aN iLnrNer-0>jc OU) CxCxOMNr 0C')ti--4e'~r-- 000.- tc,4cr~cN v'u)LDtorN 0ooocqm\~unr Oor-t-4.W)
* K~>.$'~NJ.. OQoON .- 4' 4,-4 i-w i-Iir-f - -I -v- v- N r c'CNIc'4N~ mcnm(l)cn'0000><N74-- 0000Q0ON OO OOOO OQOON ODDOO 00000
193
0 0 0 0 0 0
q) 4 0 0 lo()er~ 6 U)r'-Ocv)%OU00%0 UO 044o 0I00 '0 NOWHU)MW0 '0 H 0~n%-OD0 %0v-I CV) . . % 1r . . ... '0.4 .e' .. . .
OH- w-4ir--e 0 000r-jw-4-4r-4N 0 0000000N 0
00 0 0OD %A -Ln~) L) O~-LNL C'40( l
U) C*4 0 %D N m~oo W~t-%)oot)0 m)N)omwmOOr) OH 0w-I 0 q30 4 -444t-0cV)44LA v4 mt~44umr-wcD 44
C40 ..... ..Ic.0 . . ..*
0 ('- OCN C'4 0 000r-r-iu-Iw-14N 0 OOOOOOON 0
N' 00 00 00%0 N .4 LALA U) .. Utnc,4('40N.
w~ 0 i-I v4 LAO CNr-1QO4 00 LAN-cJrfl'LO000cr) 3: - 0 0cv) 000 ocn0 v; r-144C-0N44U)ooc4 w-IN044LA'or-)DC0N-
v4 0 HO .-IOo 00404 0 000wf-4w- w-4w-INO 0000000w-1e40
1,
000 00 00...Y rOU L'I .. r-Lnc'40c ..
0
C14
1 0 N 44 000 0 0Z~ .0 tD 0 LA 0 LA LOA . . U Dr LAs 04 0N tN0
OD 0 Ot . r- H oto).0tD 1-40O000 )C#r)(cnN It 0 Lor-Oci#oU) 0NO HO4 .N'ONC'0w-On-4-OLAC)rONOOCO0(%O)N~t44OOODN 440vO-CCII.rý-I
qPe',- V-444Yr-4r-;LAnV-4Utc)U)LA,-imLAU)tn-LAn,)LAV Lu)r-u)Lnu-)U)LA(V,U)1uiLA~LAr-u)uLALALALAU)CIlAV~r.
1.94
0Qm,-4 LnCO~C ODMONc'l-4 0WNin,-4%D,-r--0 qtn-N-%r-0~a%00,-4,-0 O %D OCDCO~0-4r-I 00
. . (1~f 4 0t.O.O66Hfrq4~c~ 66~~rr4 0'-
r-MMO~OHHOi- do m-oo00-4.- 00
4.s
00
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r-MMOOHOON r-moor-N -
C;60,HH*lOHO 'OcOH~ C,,i 0r4r
tnGOr'C0D -.V- 00 HM Qhr1mot E" O0Q %D0%
0 Hr-
.... : 1h .01 . n n 4
O -11 ..... .0 0.. .0...000oQIrbi-"l 0 OOOOHHH. Nr 4 0 Orm 0* r .0
41 LO 4*r 4g MO MiM4 " UN %0 t -'r 4rr0 ~ r 0 0 4 '4 v
0f L qUInWMnW'f))Innn W ,r - 0o- -
r~44N0000000000000000000000000000
5,.. 195
- - - - - - - - - - - - - - -
0 ~0w
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0 .L~.....n
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Q~c7G~QO0O-4 0QrlD~N
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19 I
IV. V/STOL AIRCRAFT EXAMPLE
This chapter contains the output generated by running
VASCOMP II using the data file shown at the end of
Chapter III. The sequence of the results printed by the
program using the standard format is as follows:
1. Data Echo2. Size Data3. Passenger Sizing Data4. Weights Data5. Propulsion Data6. Aerodynamics Data7. Mission Performance Data8. Sizing and Performance Summary"9. Mission Data Summary
1.9
S19~7
0 0+ +
IV ~000 00 00 ,.0C0 00 00 00 0 MILO 0
to U*) .0 0 Hr
000
00O 0 000 0 0+++ +.
0 000 O0000 00 0C) 0000 t" 0 00 E4 00 < 00MV)0 . .O0Owl 0. z 00 E-1 > ..- iO -40 - 00 0 -04 .r- ý- 0
9 mHN . 14 z-*-c . r-1. C~'qw4 004 124 E-1 00 0 0 z 0;C4 -4-4 000
' 000 0 E-w 0000 z
N Z +++ 0 00 Oi-4'-H0 "-0a4 0-.. 000 00 00 No
'4 Z -IU + +.-g +1 1+ 6-4 t0wC000rz W 4Wo4 EooEO 0u. 0 <630z N' 00000 0000000 c00 gn - CIO04I& 0 . 0000. * 00W)O zO00 WOOC0 aNN wz XI0000 4 00000 C'400 .00(t 000. :D(%40U 0 W~ PE-4E-E-E- > 0 .000 C'40UOD)UNOC'4 <0 .r-4 E-4as
E- M)(I0000 0 00 000 000 rWO0 u) :DZZZZ U Z
9 u 0 0 ::)Mon0 wj 0 N>Z~ZZEE- U)~ IXz z0000 0000 0~ 1% 00 W W 04'-4404C04040 0000 0 00 0 0
N w0 nww ....U)U <+- +t An+ +0 0~+W4 Z$04rz1lr~z. wwrZ~r E-4 W i 4 r4 0&I> (t~ oz=4 H 00000 <~ OOLAOOO a0000 W~O Or*-"c~%= 0. 0000. Q mNOC'.ILAO Wl000 Zoo 0.
0 z 0000 K4 00000 C144 ~an -. 0 Z00.0 "-'C- or--E- z < W0I-0U0o > 0O000.O 0 N4 .0 .CV)0 "-0 .r40o . Or-I'-4 u >< 0..(4'< .CMLONN'4LA.=)r-lN .Z-4 .U-CN l~rzil 06000 z 0 0 00 0 0 0
w U)
xW<4<< 002:000-M>>>> ++I M+ + 0ý
a4 0E Qi~ W Wi Wi azl4g~~ z~~~ 00000 00000000 ZOO00 0000ar) Z- 000O0 0 O0O000 H-OOO00mW-U) WiOOOO E-40000< U) 0 wwwwdEI 4 00001-. WO~OOONMOO~~i 00000 niWOOLAO
> *.0OOLO' < . . .U~- O..*r tO E-E.l4 [ t-404 .4 . -. 4' i r-ir-4 . 1% 04 04 .(04co 0 0~D ~ 0000u0 OU2O 00 0IN, E- Z0 ) UE)Losa0L0 fn 1- I-:D<4<4-
N,
H - L,- -'D-UOei~'O r-4iONW)C- Ol %D cv)nJw v-i-qc4w-C 00O0-q',4,4c-4 tur)L D %0 00O Nb-z> > 0 r4iir4-it-ir-if- v4-lh-it-it ('4C1CN('
E--
198
0
0 010 .0 *0 adONLA000 0 00nui0 0" 000 0 0d'UWt-ivo0 0 0N4 WD A 0.1.eg tflD..4toND. .. N %or-4
0 0 0V0 00
0 40 0+8 I
0 0 0 0 0 OW 0660 pooLnoLnoO. o0C 0 0 0 0 0e40000 o0Lflkifomml0 000 in N' 0 0 f4N 0otM*000
0 %D * 0 '0NC v H-1't- -.... 0 vr4IN * C-4 NN . -. 1fI4f-i4CN .
00 6 0 o0 00
r-4 00 00 00 0
+4. +
cq0 0 0*0 0000.000 0i 0.0 0Ch 0 0 H-40 0 ON' 0000OLA00c)o (03C)rol 0 "0)00 6000
0 ko . ' 0 000 0cm0 0 NNNON410N'f000 it)r-.-I Z H '-. H1. . *C-4N% . . -rrr4l'
0 0 rz4 0 0 000 0 o000 0 0H3
00m0 0 H 0
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LIST OF REFERENCES
1. NASA CR 2544, Conceptual Desig- Study of 1985 CommercialTilt Rotor Transports, by Bell Helicopter Comnany, p. 7,May 1975
2. NASA CR 151987, Applications of Advanced V/STOL AircraftConcepts to Civil Utility Missions, by The AerospaceCorp., v. 2, pp. 25-144, February 1977.
3. Boeing VEPTOL Company, Report D8-0375, User's Manual forVASCOMP II, The V/STOL Aircraft Sizing and PerformanceCornuter Program, Volume VI, Third Revision, byA. H. Schoen et al, May 1980.
4. Boeing VERTOL :ompany, US Navy Contract N62269-79-C-0217,User's Manual fcr HESCOMP, The Helicopter Sizina andPerformance Computer Proqram, Second Revision, byS.J.D--avis et al, October 1979.
5. Zuk, J., "Civil Benefits of the JVX", Vertiflite, v. 30,No. 1, pp. 20-23, Jan/Feb 1984.
6. Jane's All The World's Aircraft 1971-72, pp. 262-264,Sampson, Low, Marstol & Co. Ltd., 1971.
7. Martin, S. and Peck, W. B., JVX Desian Update, paperpresented at the 40th Annual Forum of the AmericanHelicopter Society, Arlington, Virginia, May 16-IP, 1084.
214
BIBLIOGRAPHY
"Advanced VTOL Concepts to be Studied," Aviation Week &
Soace Technology, April 30, 1973, pp. 85-86.
"Bell Tilt-Rotor: The Next V/STOL?" Flicht International.February 9, 1980, pp. 381-386, 412.
"Bell XV-15 to Test Tilt-Rotor Concept," Aviation Week &Space Technology, November 1, 1976, pp. 20-21.
"Development of V/STOL Aircraft 1950 to 1970," aTntrnal of theAmerican He1Vcooter Society, June 1979, pp. 7-75.
"Flexibility is Offered by XV-15 Tilt-Rotor Concept,"Aviation Week & Space Technology, January 11, 1982,pp. 74- 76, 81, 83, 85.
"Flying the V-22 Predecessor: The NASA/Army/Bell XV-15,"Rtor &. Wing International, June 1985, pp. 44-50, 54-58, 93.
"Handling Qualities Evaluation of the XV-15 Tilt RotorAircraft," .i�,,rnal nf the Amerinan Helicopter Society, April
1975, pp. 23-33.
"Improving the Helicopter - What Next?" )Inara.in, October1972, pp. 1121-1123.
"Joint Services Vertical Lift Development (JVX) Program:Looking to the Future," La.rtiflie, Nov/Dec 1984, pp. 24-25.
"JVX: Entering Preliminary Design," Veriflitte. Nov/Dec1983, pp. 18-19.
"JVX: The World's First Production Tilt-Rotor?" .Lntae, via,July 1983, np. 755-757.
"JVX: Tilt-Rotor Sets Out to Succeed," Flight Tnt-rnatianal,November 5, 1983, pp. 1214-1216.
"JVX Program: Still a lot of Questions," Inleravia,, April
1983, pp. 348-349.
"Military Potential of the Tilt-Rotor Aircraft," ApronauticajJnurnal March 1983, pp. 99-103.
"Moving V/STOL front Technology to System," tistronautics andAP.nnautj.=, December 1977, pp. 26-27.
215
"Proposal Request for JVX Authorized for December," AviationWeek & Space Technolog=, November 22, 1982, pg 29.
"Structural Components, Design of Tilt-Rotor JVX NearCompletion," Aviation Weak &_Spe Technoog~v_ January 14,1985, pp. 84-87.
"The Promise of Tilt-Rotor," Aireraft Engineerino July1978, pp. 11-13.
"The XV-15 Experience From Wind Tunnel and Simulations toJoint Services V/STOL Aircraft," Presented at Army ,perationsResearch Symposium, Fort Eustis, Virginia, October 4-5, 1983.
"The XV-15 Tilt Rotor Research Aircraft," Snnietv ofIx•grimental Test Pilots Symposium Prtco•di~gz, Sept. 1980,
pp. 168-185.
"Tilt-Rotor Aircraft," Aj= Research,. Development andAcquisition Maagine, May-June !q80, pp 1-3.
"Tilt-Rotor Aircraft Nears Final Design," Aviation Week &
Spae TnchnnIlogy. May 20, 1974, pp. 51-53.
"Tilt-Rotor: Expanding Rotorcraft Horizons," 1ertiflite,May/June i980, pp, 8-11.
"Tilt-Rotor VTOL," EnIu.ar.cence, February 1978, pp. 68-70.
"V/STOL - Turning Promise to Reality," Astro autics ander.onaui•s. , September 1965.
"XV-15 Experience: Joint Service Operational Testing of anExperimental Aircraft," Society of Experimental Test PilotsSymposium Proceedings, September 1983, pp. 4-20.
Si•
|f
INITIAL DISTRIBUTION LIST
No. Copies
1. Defense Technical Infou'ation Center 2Cameron StationAlexandria, Virginia 22304-6145
2. Library, Code 0142 2Naval Postgraduate SchoolMonterey, California 93943-5002
3. Department Chairman, Code 67 5Department of AeronauticsNaval Postgraduate SchoolMonterey, California 93943-5000
4. Professor Donald M. Layton, Code 67Ln"Department of AeronauticsNaval Postgraduate SchoolMonterey, California 93943-5000
5. Captain Thomas P. WalshUS Army Avaition Engineering Flight ActivityEngineering Flight Test Division AEdwards Air Force Base, California 93523-5000
217
