NASA CR 175020

NfLg/ 

APPLIED TO GENERAL AVIATION

[N&SA-CR-175320) EVAIUAIICN CF _OPFAN

_OPULSICN APSLIE_ TO G_N£_AL AVIATION

[Beech Aircraft Corp.) |45 p HC A07/_F A01

CSCL 21E

by

R. W. Awker

G3/07

N86-24695

Unclas

43075

BEECH AIRCRAFT CORPORATION

WICHITA, KANSAS

Prepared For

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

Lewis Research Center

Cleveland, Ohio 44135

Contract NAS3- 24349
NASA CR 175020

I /ISA

EVALUATION OF

APPLIED TO

PROPFAN

GENERAL

PROPULSION

AVIATION

by

R. W. Awker

BEECH AIRCRAFT CORPORATION

WICHITA, KANSAS

NATIONAL

Prepared For

AERONAUTICS AND SPACE

Lewis Research Center

Cleveland, Ohio 44135

ADMINISTRATION

Contract NAS3- 24349
FORWARD

This report presents the results of a study conducted by The Beech Aircraft

Corporation with support from Hamilton Standard Division of United

Technologies Corporation (UTC) and Pratt and Whitney, Canada Division of

United Technologies Corporation (UTC) for the National Aeronautics and

Space Administration, Lewis Research Center (NASA-LeRC) under NASA contract

NAS3-24349, "Multiple Application Propfan Studies (MAPS) Category One (1)"

The study was conducted from August, 1984 to June, 1985. The Study Manager

for Beech was Randal W. Awker. The NASA Contract Manager for the study was

Susan M. Johnson, NASA-LeRC, Cleveland, Ohio.

The author wishes to express special thanks to the following UTC personel

who provided valuable technical assistance and propulsion system data that

made this report possible.

Derek Emerson Pratt and Whitney, Canada

Ira Kei ter Hamilton Standard

Richard Millar Pratt and Whitney, Canada

Bernard Palfreeman --- Pratt and Whitney, Canada

Additionally, the author wishes to express his appreciation to Pratt and

Whitney, Canada for supporting the study at no cost.

F_:IECEDING PAGE, BLANK NOT FILMED

iii
TABLE OF CONTENTS

PAGE

1.0 SUMMARY ............................. 1

2.0 INTRODUCTION ........................... 3

3.0 DESIGN APPROACH ..................... 7

3.1 Design Ground Rules' . ................... 7

3.2 Aircraft Synthesis ..................... 14

3.2.1 Carpet Plots .................... 14

3.2.2 Optimization .................. 16

3.3 Cost Models and Aircraft Pricing .............. 19

4.0 CONFIGURATION DEFINITION ..................... 21

4.1 Configuration Selection .................. 21

4.2 Cabin Selection ...................... 23

4.3 Engine Selection ..................... 25

4.3.1 Turbofan Engines .................. 26

4.3.2 Turboprop Engine .................. 27

4.3.3 Summary of Engine Characteristics ......... 29

4.4 Propfan Selection ..................... 31

4.5 Propulsion System Scaling ................. 36

5.0 CONFIGURATION EVALUATION ....

5.1

5.2

5.3

5.4

5.5

5.6

5.7

5.8

............... 41

Configuration Sizing and Optimization ........... 41

5.1.1 Propulsion System Performance ........... 41

5.1.2 Configuration Three Views ............. 45

Installation Concepts .................. 50

Aerodynamic Characteristics ................ 56

5.3.1 Low Speed Drag Characteristics ........... 56

5.3.2 Stability and Control Considerations ........ 58

5.3.3 High Speed Considerations ............. 60

Weight and Balance ..................... 63

Structural Desi gn ..................... 67

5.5.1 Structural Material ................ 67

5.5.2 FAR 25 Safelife Requirements ............ 67

Acoustic Effects ...................... 68

Performance ........................ 73

Cost ............................ 80

6.0 DESIGN SENSITIVITIES ....................... 89

6.1 Configuration Factors ................... 89

6.2 Cost Factors .................... 92

6.3 Configuration/Cost Sensitivity to Retrofit ......... 100

7.0 CONCLUSIONS AND RECOMMENDATIONS ................. 103

APPENDIX - COST ESTIMATING MODEL ................. 107

SYMBOLS AND ABBREVIATIONS .................... 133

REFERENCES ............................ 137

_'RECEDING PAGE BLANK NOT FILMED
LIST OF ILLUSTRATIONS

PAGE

FIGURE 1.

FIGURE 2.

FIGURE 3.

FIGURE 4.

FIGURE 5.

FIGURE 6.

FIGURE 7.

FIGURE 8.

FIGURE 9.

FIGURE 10.

FIGURE 11.

FIGURE 12.

FIGURE 13.

FIGURE 14.

FIGURE 15.

FIGURE 16.

FIGURE 17.

FIGURE 18.

FIGURE 19.

FIGURE 20.

FIGURE 21.

FIGURE 22.

FIGURE 23.

FIGURE 24.

FIGURE 25.

FIGURE 26.

FIGURE 27.

FIGURE 28.

FIGURE 29.

FIGURE 30.

FIGURE 31.

FIGURE 32.

FIGURE 33.

FIGURE 34.

FIGURE 35.

FIGURE 36.

The Four Aircraft Used to Evaluate Propfan Propulsion . . . 5

The Pusher Propfan Configuration .............. 6

Comparison of Propulsive Efficiency ............ 7

Design Mission ....................... 13

Cost Mission ........................ 13

Typical Carpet Plot with 16 Individual Aircraft Point

Designs .......................... 14

Carpet Plot with Lines of Constant Takeoff Field Performance 15

Carpet Plot with Lines of Constant Cruise Speed ...... 16

A Design Carpet Plot .................... 16

Configuration Optimization ................. 17

Aircraft Computer Synthesis Flow Diagram .......... 18

Configurations Examined .................. 22

Selected Cabin Layout ................... 24

1988 Technology Turbofan Schematic ............. 26

1988 Technology Turboprop Schematic ............ 28

ISO-Views of the Selected Propfan ............. 33

Effect of Propfan Disk Loading on Fuel Burn ........ 35

Effect of Propfan Disk Loading on Efficiency ........ 36

Turbofan Engine Size Effect on Propulsion System Weight . . 37

Turboprop Engine Size Effect on Propulsion System Weight . . 37

RPM Effect on Turboprop Propulsion System Weight ...... 38

Flat Rating Effect on Turboprop Propulsion System Weight . . 38

Propfan Diameter Effect on Propeller Weight ........ 39

Current Turbofan Carpet Plot ................ 42

Advanced Turbofan Carpet Plot ............... 42

Pusher Propfan Carpet Plot ................. 43

Tractor Propfan Carpet Plot ................ 43

Sized Engine Characteristics ................ 44

1984 Current Turbofan ................... 46

1990 Advanced Turbofan ................... 47

1990 Pusher Propfan .................... 48

1990 Tractor Propfan .................... 49

Turbofan Installation Concept ............... 51

Tractor Propfan Installation Concept ............ 52

Pusher Propfan Installation Concept ............ 53

Effect of Exhaust Stack in the Propfan Flow Field ..... 55

vii

pTTECEDING PAGE BLANK NOT FILMED
FIGURE 37.

FIGURE 38.

FIGURE 39.

FIGURE 40.

FIGURE 41.

FIGURE 42.

FIGURE 43.

FIGURE 44.

FIGURE 45.

FIGURE 46.

FIGURE 47.

FIGURE 48.

FIGURE 49.

FIGURE 50.

FIGURE 51.

FIGURE 52.

FIGURE 53,

FIGURE 54.

FIGURE 55.

FIGURE 56.

FIGURE 57.

FIGURE 58.

FIGURE 59.

FIGURE 60.

FIGURE 61.

FIGURE 62.

FIGURE 63.

FIGURE 64.

FIGURE 65.

FIGURE 66.

FIGURE 67.

FIGURE 68.

FIGURE 69.

FIGURE 70.

FIGURE 71.

Current Turbofan Low Speed Drag Characteristics ...... 56

Advanced Turbofan Low Speed Drag Characteristics ...... 57

Pusher Propfan Low Speed Drag Characteristics ....... 57

Tractor Propfan Low Speed Drag Characteristics ....... 58

Current Turbofan High Speed Drag Characteristics ...... 61

Advanced Turbofan High Speed Drag Characteristics ..... 61

Pusher Propfan High Speed Drag Characteristics ....... 62

Tractor Propfan High Speed Drag Characteristics ...... 62

Current Turbofan Loading Diagram .............. 65

Advanced Turbofan Loading Diagram ............. 65

Pusher Propfan Loading Diagram ............... 66

Tractor Propfan Loading Diagram .............. 66

Pusher Propfan Installation Acoustic Loading ........ 69

Tractor Propfan Installation Acoustic Loading ....... 69

Additional Acoustic Mass Treatment for Tractor Installation 71

Uncertainty in Noise Directivity .............. 71

Time to Climb ....................... 74

Rate of Climb ...................... 75

Cruise Speed ........................ 75

VFR Range ......................... 76

Design Mission Elements .................. 77

Cost Mission Elements ................... 78

Cost Mission Fuel Burn Comparison ............. 79

Relative Cost Summary for the Study Configurations ..... 84

Mission Sensitivity to Design Aspect Ratio ......... 90

Design Mission Impact on Propfan Selection ......... 91

Mission Fuel Sensitivity to Acoustic Treatment Weight 92

Cost Sensitivity to Chosen Cost Model and Annual Utilization 93

Operating Cost Sensitivity to Initial Aircraft Selling Price 94

Relative Operating Cost Comparison for 400 Hours Annual

Aircraft Utilization (Corporate Model) ........... 98

Relative Operating Cost Comparison for 3000 Hours Annual

Aircraft Utilization (ATA Model) .............. 98

Sensitivity of Operating Cost to Fuel Price at 400 Hour

Annual Aircraft Utilization ................ 99

Sensitivity of Operating Cost to Fuel Price at 3000 Hour

Annual Aircraft Utilization ................ i00

Steps to Evaluating the Cash Flow Utilized in the Corporate

Cost Model ......................... 122

Propulsion System Acquisition Cost Factors ......... 130

viii
LIST OF TABLES

PAGE

TABLE 1.

TABLE 2.

TABLE 3.

TABLE 4.

TABLE 5.

TABLE 6.

TABLE 7.

TABLE 8.

TABLE 9.

TABLE I0.

TABLE ii.

TABLE 12,

TABLE 13.

TABLE 14.

TABLE 15.

TABLE 16.

TABLE 17.

TABLE 18.

TABLE 19.

TABLE 20.

TABLE 21.

TABLE 22.

TABLE 23.

TABLE 24.

TABLE 25.

TABLE 26.

TABLE 27.

TABLE 28.

TABLE 29.

TABLE 30.

TABLE 31.

TABLE 32.

TABLE 33.

TABLE 34.

TABLE 35.

TABLE 36.

Current Small Turbofan Business Aircraft Characteristics . 8

Cabin Design Criteria .................... 23

Engine Characteristics Summary ............... 29

Hamilton Standard Propfan Blade Designation ......... 32

Comparison of Propfan Performance Estimates ......... 34

Horizontal Tail Characteristics ............... 59

Vertical Tail Characteristics ................ 60

Configuration Weight Summary ................ 64

Community Noise Estimates .................. 72

Performance Summary ..................... 73

Design Mission Fuel Summary ................. 77

Cost Mission Fuel Summary .................. 78

Acquisition Costs ...................... 81

Relative Propulsion System Acquisition Cost ......... 81

Total Cost of Operation for 400 Hour Annual Utilization . . . 82

Variable Operating Costs - Corporate Model ......... 83

Fixed Annual Operating Costs - Corporate Model ....... 83

Modified ATA Costs for 400 Hour Utilization ......... 86

Modified ATA Costs for 3000 Hours Utilization ........ 87

Relative Weight of Propfans Compared to Turbofans ...... 95

Relative Cost Factors of Propfans Compared to Turbofans . . 96

Relative Operational Factors of Propfans Compared to

Turbofans .......................... 97

Current Turbofan with Retrofit 1988 Propulsion Systems . . I01

Current Turbofan ATA Analysis at 400 Hours Utilization . 112

Advanced Turbofan ATA Analysis at 400 Hours Utilization . . 113

Pusher Propfan ATA Analysis at 400 Hours Utilization .... 114

Tractor Propfan ATA Analysis at 400 Hours Utilization .... 115

Current Turbofan ATA Analysis at 3000 Hours Utilization . . 116

Advanced Turbofan ATA Analysis at 3000 Hours Utilization 117

Pusher Propfan ATA Analysis at 3000 Hours Utilization .... 118

Tractor Propfan ATA Analysis at 3000 Hours Utilization 119

Operating Cost Elements ................... 120

Current Turbofan Corporate Cost Analysis .......... 125

Advanced Turbofan Corporate Cost Analysis .......... 126

Pusher Propfan Corporate Cost Analysis ........... 127

Tractor Propfan Corporate Cost Analysis ........... 128

ix
1.0 SUMMARY

This study examined the potential application and benefits of propfan

propulsion for General Aviation aircraft. The study's objective was to

examine various candidate aircraft configurations for propfan propulsion

and select a configuration which would allow a good comparison of propfan

propulsion to conventional propulsion for a similar mission cruise speed of

.7 to .8 Plach. The selected configuration was then used to compare

the environmental impact, cost, performance, and market potential of

propfan propulsion against conventional turbofan propulsion.

A generic, small turbofan designed with characteristics similar to those of

current aircraft in today's General Aviation fleet was chosen as the basis

for the propulsion system comparison in the study. To provide a common

configurational basis for comparison, four similar aircraft were designed,

each with the same basic aerodynamic characteristics but with four

different propulsion systems. The four propulsion systems consisted of a

current technology turbofan engine, an advanced 1988 technology turbofan

engine, and an advanced 1988 turboprop engine combined with both a pusher

and a tractor propfan installation. Each aircraft examined was designed

for eight passengers with 1800 NM NBAA IFR range.

Comparison of the propulsion systems in these four similar business class

aircraft shows that the single-rotation (SR) propfans evaluated in this

study can provide a 33% reduction in fuel consumption compared to current

day small turbofans and a 14% reduction compared to equivalent technology

turbofans.

The study showed, however, different cost trends than previous studies of

similar propulsion systems in large transport aircraft applications. The

propfan propulsion system (engine and propeller) utilized in this study had

a 35% higher acquisition cost than an equivalent turbofan. This resulted

in a significantly higher total aircraft selling price for the propfan.

Because of lower utilization rates, the total cost of operating General

Aviation aircraft was shown in this study to be more sensitive to change in

aircraft acquisition cost (or airplane selling price) than to a change in

fuel costs due to a change in fuel consumption. Thus the higher price of

the propfan in this study overshadowed the considerable improvement in fuel

consumption and resulted in the propfan having a 4% higher total cost of

operation than the current technology turbofan and a 10% higher cost of

operation compared to an equivalent technology turbofan.

I
2.0 INTRODUCTION

Advanced, high speed propeller systems or propfans have been under

development since before 1976, primarily due to the fuel crisis in mid-1974.

Propulsion systems utilizing propellers can show significant benefits in

fuel consumption over current turbofan aircraft. The main problem in the

past is that these propeller driven aircraft could not compete with the

mission capability or the passenger comfort and appeal of the turbofans.

Prior to the fuel crisis in 1974, fuel was not of major concern in the

overall cost of operation of most U.S. aircraft and, therefore, there was

little incentive to design competitive, high speed turboprops. However,

after the 1974 fuel crisis with the substantial increase in fuel prices and

the spectre of fuel shortages, fuel conservation became of major public

interest. Fuel costs soared to over 50% of the total cost of operation

(Reference i) for large commercial and military transports.

In response to the concern from the Senate Committee on Aeronautical and

Space Sciences, NASA implemented in 1975 the Aircraft Energy Efficiency

(ACEE) program to investigate areas of technology that could conserve the

fuel used by U.S. aircraft. Propulsion system technology was among those

areas identified for investigation, and consisted of three general

programs.

• Energy Component Improvement (ECI)

• Energy Efficient Engine (E3)

• Advanced Turboprop Program (ATP)

The high speed propeller technology utilized in this study developed out of

the Advanced Turboprop Program. The Advanced Turboprop Program, begun in

1978, was designed to be a three phase effort (Reference 2). Phase I was

to develop a fundamental data base using small scale models and to

establish the concept feasibility. Phase II was to establish the design,

fabrication, and ground testing of a Large-Scale Advanced Propeller (LAP)

with a nine-foot diameter. Finally, Phase III was to complete the

necessary system integration to perform a flight research program on a

commercial-type aircraft (a Gulfstream G-II was selected) and was referred

to as the Propfan Test Assessment (PTA) (Reference i).

_r_.EDI_'_G PAC_ BLANK NOT FILMED
The majority of effort in Phase I has been directed towards large

commercial and military transport class aircraft. Many of these studies

have indicated possible 15% to 30% reduction in fuel use for these aircraft

by using propfan propulsion instead of conventional turbofans. These

studies have also shown that the propfan propulsion system has only

slightly higher (+3%) acquisition cost. This, combined with the significant

fuel savings shown, results in a 6% to 10% lower direct operating cost

(DOC) for these transports.

To identify other potential areas of application for this high speed

propeller technology NASA initiated the Multiple Application Propfan

Studies (_APS) program. Under the MAPS program six potential aircraft

categories of interest were identified.

(1) Business Aircraft

(2) Military Light or Heavy Attack Aircraft

(3) Long Endurance Aircraft

(4) VTOL, STOVL, or STOL Aircraft

(5) Unpiloted or Remote Piloted Aircraft

(6) Unique/Other

This study addressed the potential application of propfan propulsion for

the first category, business aircraft. This study was performed as part of

the MAPS program under the NASA contract NAS3-24349.

The objectives of this study were to:

• Assess the potential of propfan propulsion for business class

aircraft.

• Compare propfan propulsion to conventional propulsion for an aircraft

performing the same mission.

• Identify areas requiring further development.

• Recommend application of the promising technology.

4
The Beech MAPS study was divided into four tasks to accomplish these goals.

Task I - Definition of Study Evaluation Procedures and Assumptions

Based on a preliminary configuration evaluation, the study

ground rules, design approach, and aircraft configuration

were selected.

Task II -Conceptual Design

Four aircraft were conceptually designed. The four

aircraft are shown in Figure i. Two of these aircraft

utilized conventional propulsion systems and two utilized

propfans. One conventional system was designed to current

technology practices while the other conventional system

and both propfan systems were designed utilizing 1988

propulsion technology. An artist's conception of the

pusher propfan installation is shown in Figure 2.

TURBOFAN PROPFAN

FIGURE 1. The Four Aircraft Used to Evaluate Propfan Propulsion

5
ORIGINAL PAGE IS 

OF PCOR OIJALTTY 

FIGURE 2. The Pusher Propfan Configuration 

Task I11 - Mission Evaluation 

The four aircraft designed in Task I1 were then evaluated 

for the same mission to determine cost and mission 

performance. 

Task IV - Conclusions and Recommendations 

The study results were summarized in Task IV. Also, 

specific areas requiring further research identified in 

Task I11 were presented and summarized in Task IV. 

6 
3.0 DESIGN APPROACH

This section presents the ground rules and design approach for tile study.

The ground rules were established after a preliminary configuration

evaluation. A detailed discussion of the preliminary evaluation is given

in Section 4. The four aircraft evaluated in this study were all sized and

optimized for minimum fuel burn utilizing the following design and mission

ground ruleso

3.1 DESIGN GROUND RULES

In the preliminary evaluation phase of the study both current turboprops

and turbofans were examined as potential candidates for the study's

baseline current technology propulsion system. As can be seen in Figure 3,

for a given engine size (or horsepower level) propfans begin to show an

advantage over turboprops only at speeds higher than .65 Mach. Figure 3

100-

9O

8O

INSTALLED

PROPULSION

EFFICIENCY

%

7O

6O

5O

KING AIR B200 STARSHIP DIAMOND II

{Z> {i> {Z>

i I I t I III II

TYPICAL CRUISE 500 1000 2000

HP/ENGINE

SR PROPFAN

TURBOFAN

.3 ' .4 " .'5 ' .'6 " .7 " .8

CRUISE MACH NUMBER

.9

FIGURE 3. Comparison of Propulsive Efficiency

7
shows a typical turboprop, propfan, and turbofan propulsion system compared

at each system's design cruise altitude. Figure 3 illustrates the effect

of cruise speed on propulsive efficiency for the chosen propulsion systems.

It does not, however, represent the envelope of optimum design for all

propellers. Further, Figure 3 illustrates how the required engine size

must increase for increased cruise speed capability, for the class of

aircraft examined in this study, regardless of the propulsion system.

Since it was desired to have a current technology aircraft performing the

same mission as the propfan aircraft in this study for comparison, the

turboprop was eliminated as a candidate primarily because there are no

current general aviation turboprop aircraft which cruise in the .7 to .8

Machspeed range. There are, however, several small business turbofans

which do cruise in this range, as shownin Table 1. These were chosen to

provide a basis for establishing the ground rules of the study. All the

performance characteristics of these turbofans are achievable using propfan

TABLE1. Current Small Turbofan Business Aircraft Characteristics

ITEM

PASSENGER CAPACITY

MAX CRUISE SPEED MACH

VFR RANGE NM

TAKEOFF FIELD LENGTH FT

@ S.L., STD

MAX TAKEOFF WEIGHT

FALCON LEAR CITATION DIAMOND

100 35 SII II

.84 .80

10

.70 .77

FUEL WEIGHT

1880

4200

2289

4224

1800

3240

1810

3950

LB

LB

WING LOADING LB/FT 2

THRUST LOADING LBT/LBM

SPAN LOADING LB/FT

19300

5909

74.4

.33

45O

17000

6237

67.1

.41

430

14700

5777

42.9

.34

282

15780

4904

63.3

.37

363
propulsion. Additionally, the small business jet aircraft also provides a

good baseline for comparison because it represents a large part of the

business market (in terms of units) and provides the greatest spread from

the current transport class aircraft being evaluated for propfan propulsion.

The results of the preliminary configuration evaluation and selection are

discussed further in Section 4.

The design ground rules for the study were selected, using the small

turbofan market as a basis, to provide a good comparison between three

propulsion systems: a current technology turbofan, a 1988 technology level

turbofan, and a 1988 technology level propfan. Later in the study the

propfan propulsion system was split into two installations, a pusher and a

tractor, and a total of four propulsion systems were compared. Further,

the design ground rules were selected so that a potential flight prototype

would be feasible by the 1990 time frame. These ground rules were then

used to design the four aircraft utilizing the four propulsion systems

examined in this study. These ground rules were submitted to and approved

by NASA in the initial phase of the study.

The design ground rules for the study were:

(1) TECHNOLOGY READINESS DATE

A 1990 aircraft technology readiness date was selected.

Aircraft technology readiness, as used in this study, is defined

to be the date at which a prototype would be ready for initial

flight test. The 1990 date was based on information from

Hamilton Standard indicating, for airline class propfans, the

earliest operational readiness date would be 1987 for a prototype

propulsion system. It was anticipated that an additional year

and one half would be required by both engine and propfan

manufacturers for development of a propfan/engine installation

for use on a general aviation aircraft. This resulted in a

propulsion technology readiness date of mid-1988. Subsequently,

an additional year and one half for design and construction of a

flying prototype would be required by the airframe manufacturer,

thus yielding a 1990 aircraft technology readiness date.

9
(2) AIRCRAFT AERODYNAMIC TECHNOLOGY

All of the study aircraft were designed to have comparable

aerodynamics. Each incorporated a wing design which allowed a

significant amount of laminar flow. The wing configuration was

selected to optimize cruise performance and to prevent any

adverse shock effects. The high lift system design was

sufficient to provide the necessary capability to achieve the

desired 4000 FT takeoff field length at sea level. The tail

sizing accounted for differences necessary to maintain

directional control and sufficient static stability for a typical

business aircraft center of gravity travel.

(3) AIRFRAME STRUCTURAL TECHNOLOGY

The structural design of the current technology aircraft

incorporated conventional aluminum construction techniques. The

structural design of the 1990 technology level aircraft

incorporated advanced composite materials and advanced

aluminum-lithium alloys. The fuselage design of the advanced

aircraft utilized filament wound graphite epoxy with a Kevlar

based core and integrally woven metal wire for lightning

protection. The wing utilized a combination of advanced aluminum

alloys and advanced composite materials. Although the scope of

this study did not warrant a detailed structural analysis,

the airframe weights and cost estimates discussed in Section 5

reflect the effects of these current and 1990 level structural

design approaches.

(4) PROPULSION TECHNOLOGY

To meet a total aircraft technology readiness date of 1990

requires an engine technology readiness of 1988, as previously

stated in (1). The engine for the propfan aircraft consisted of

a Pratt and Whitney, Canada PT6A cycle scaled to the necessary

horsepower level to meet the mission requirements with weight and

fuel consumption adjusted to be consistent with the 1988

propulsion technology readiness. The turbofan aircraft utilized

a JT15D-5 cycle also adjusted to the 1988 technology level. The

1988 level engine technology for both utilized advanced material

technology (single crystal turbine blades) and advanced air flow

technology (improved gas path in compressor, combustor, and

10
exhaust). Engine data and technical assistance was provided by

Pratt and Whitney, Canada to assure consistent levels of

technology for the current day and 1988 engines.

The study utilized a propfan designed for General Aviation

application by Hamilton Standard with a sweep distribution and

diameter selected based on the study's performance, cost, and

noise goals. Propfan performance, cost, and noise were evaluated

using methods supplied by Hamilton Standard.

(5) CERTIFICATIONREQUIREMENTS

Both study aircraft were designed to FAR 25 (Airworthiness

Standards: Transport Category Airplanes) requirements. Both

aircraft were anticipated to be over 12,500 LB and therefore,

required FAR 25 certification. In particular, there were two

areas which impacted the aircraft design for this study resulting

from FAR25 certification:

(1) Hot day, high altitude climb capability

(2) Engine out control (VMC)

(6) ENVIRONMENTAL NOISE

Both near field and far field noise were considered in this study.

The preliminary evaluation indicated that small business jet

interior noise levels range between 78 dBAand 85 dBA. A value

of 82 dBAwas chosen for the cabin noise level target in this

study. Each design incorporated the acoustical treatment and

associated weight penalty to achieve this maximum82 dBA noise

level for both the propfan and turbofan desi'gns. For far field,

all aircraft were required to meet FAR 36 (Noise Standards:

Aircraft Type and Airworthiness Certification) Stage III noise

requirements.

(7) AIRCRAFTPAYLOAD

Each of the configurations was designed to provide a cabin with

seating for a maximumof 8 passengers or a maximumpayload with

full fuel of 1200 LB. The assumedweight for each passenger and

crew was 200 LB including baggage.

11
(8) MISSIONPROFILES

The mission analysis for this study was done using two missions.

The first mission, shown in Figure 4, was used to size each

aircraft and is referred to as the design mission. The second

mission, Figure 5, was used to evaluate the operating costs using

the direct operating cost methods discussed in the Appendix and

is referred to as the cost mission. The cost mission represents

a weighted average of actual missions being flown by

owners/operators of turbofan and turboprop aircraft based on

current industry experience. Over 50% of General Aviation

missions flown are less than 400 NM. Only 15% of the missions

flown today require the aircraft's maximum design range

capability.

(9) TAKEOFFFIELDLENGTH

The design takeoff field length was selected to be 4000 FT at sea

level, standard day. This field length allowed the design of a

configuration with the high speed capabilities required to

provide a valid comparison between propulsion systems without

over penalizing the design for takeoff field length.

12
CRUISE {_WAX. CRUISE .75M

41,000 FT

/TAKEOFF

_ eT_W

S'L'--_START, TAXI

lO MIN

RAMP T.O.

ORIGIN

CRUISE BBEST RANGE SPEED HOLD,

DESCENT _-_ / MAXIMUM _,_ DESCENT

AORROA --A/.---'ISSED AOPRGACH

STY'. I 5 MIN I

LANDING RAMP LANDING RAMP

BLOCK FUEL RANGE AND18OOBLOCKNM TIME =I= ". 200 NM RESERVE FUEL _ I

DESTINATION ALTERNATE

FIGURE 4. Design Mission

33.000 FT

MAXIMUM /

_I_ _T;_Eo_

S.L.--I_

_ STA_g./_xI

RAMP T,O.

. F

ORIGIN

CRUISE @ .75M

/

DESCENT

5 MIN

LANDING RAMP

J

RANGE 400 NM -I

BLOCK FUEL AND BLOCK TIME

S.C.

DESTINATION

FIGURE 5. Cost Mission

13
3.2 AIRCRAFT SYNTHESIS

The aircraft synthesis technique used in this study is a multi-mission

parametric optimization method which utilizes several steps in the

synthesis process. This allows the flexibility of interfacing with the

system at various key points in the design process to allow custom

tailoring of the process to fit a specific class of aircraft or a specific

set of mission requirements.

3.2.1 CARPET PLOTS

As part of this process, the aircraft synthesis technique utilizes carpet

plots as a means of multiple dimensional parametric selection and

optimization. In Figure 6, each intersection on the carpet represents an

individual aircraft design (or point design). As shown in Figure 6, each

point design is characterized on the plot by a unique combination of wing

area and engine size, as indicated by the static sea level thermodynamic

thrust. Each point design is sized to meet the design mission which was

specified in the ground rules, and thus the weight and usable fuel varies

for each point design. All 16 point designs in Figure 6 are for the same

wing geometry such as aspect ratio, sweep, taper ratio, and wing thickness

18800,

18600,

18400,

18200 -

18OO0 -

17BOO -

17600

TRKEOFF

HEIGHT 174O0

17;=OO

17OO0

168OO

16600.

164OO.

16200.

16000.

T4 T - SLS THRUST

S - HING RRER

T2

Tl

$4

$3

FIGURE 6. Typical Carpet Plot with 16 Individual Aircraft Point Designs

14
IBBOO 1

IBBOO.

184OO TOFL - TRKEOFF FIELD LENGTH

_ T - SLS THRUST

18200

18000 1

17800 1

176001

TRKEOFFI74001

HEIGHT

172001

170001

ISBO01 TOFL3

166001

164001 TOFL2

FIGURE 7. Carpet Plot with Lines of Constant Takeoff Field Performance

ratio. Additionally, each point design has an individual set of

performance capabilities. The performance characteristics of each point

design, such as takeoff field length shown in Figure 7, can be interpolated

to yield lines of constant performance. Performance constraints are

minimum or maximum design requirements and are indicated by lines with

cross-hatching, as in Figure 7. The process can be repeated for each

design constraint, such as cruise speed, shown in Figure 8. Finally, as in

Figure 9, all of the design requirements can be applied to define the

design region, shown unshaded, in which any combination of engine and wing

area satisfies both the design mission requirements and the required

performance capabilities. Also, the sensitivity of other cost or mission

parameters, such as mission block fuel in Figure 9, can be examined.

15
18800

18600

18400

18200

18000

17800.

17600.

TRKEOFF

WEIGHT 17400.

17200.

17000,

16B00

16600

16400

16200

16000

Vcr - CRUISE SPEED

T - SLS THRUST

S - WING RRER

Vcr2

T

CONSTRRINT

S

FIGURE 8. Carpet Plot with Lines of Constant Cruise Speed

18800.

18600.

18400.

18200.

18000.

17800,

TRKEOFF 17600,

HEIGHT 17400,

17200'

17000

16800

16600

16400

16200

16000

TOFL = TRKEOFF FIELD LENGTH

Grad - THKEOFF CLIMB GRRDIENT

Vcr " CRUISE SPEED

F - MISSION BLOCK FUEL

T :

_i_::". T

Grad

R - DESIGN POINT TOFL

FIGURE 9. A Design Carpet Plot

16
3.2.2 OPTIMIZATION

In Figure 9, all the point designs in the unshadedarea satisfy the design

requirements but do not necessarily yield the most optimum design. Once

the design region is defined, then the optimization can be accomplished by

finding the point which provides the maximumor minimumvalue or values of

a given figure or figures of merit. For example, in Figure 9, Point A

represents the design which satisfies all the mission requirements, all

performance constraints, and provides minimummission fuel burn.

Point A, in Figure 9, then is the optimized point design for a specific set

of propulsion system characteristics and aircraft geometry. The next step

in the process is to modify the aircraft geometry and/or the propulsion

system, repeat the carpet selection process, and obtain a new optimized

point design for the new characteristics. As shown in Figure 10, this

process is repeated, varying each of these characteristics until an overall

_ TOFL - TAKEOFF FIELD LENGTH

/\ / / _ o \ Grad - TBKEOFF CLIMB GRRDIENT

_ _/_ _ \ Vcr - CRUISE SPEED

/ _ / _ / _ _ _ D - PROPELLER DIAMETER

_ _'_ / _ _ RR - WING ASPECT RRTIO

X4 TOFL

X3 X2 Xl

where X can be

BURN I

NOTE:

D'l _2 D_ D_

REQUIRES IB,384 POINT

DESIGNS FOR ONE CONFIGURRTION

FIGURE 10. Configuration Optimization

17
optimum is obtained, or until a constrained design region is defined. For

this study the parameters indicated in Figure 10 were considered. However,

as will be discussed in Section 6.1, not all the parameters have a major

impact on the design selection.

Figure 11 shows the top level flow diagram of the synthesis process. The

baseline aircraft characteristics used as a point of departure by the

synthesis method is typically defined in a preliminary evaluation phase

which yields an aircraft with approximately similar characteristics to

those of interest.

I Input Baseline 1

I

I Generate 4 X 4 Matrix I 4 Engines, 4 Wing Areas

I

IMissi°n Analysis I I

@-_ Adjust Fuel H Weight _0,o,, i i Weight -,us,!

I I

I PerfOrmance I Drag Drag

TOGW I I I

i0o,,_ iTail-,u,, Size i iTail_0,o,, Size i

4 Engines Complete

Next Wing Area

Ico, I

FIGURE 11. Aircraft Computer Synthesis Flow Diagram

18
3.3 COST MODELS AND AIRCRAFT PRICING

Two separate cost models were used to evaluate the cost factors for this

study. One cost model was based on the Air Transport Association of

America (ATA) 1967 cost model modified to reflect General Aviation

experience. The other was developed specifically for this study to better

model the cost environment experienced by most of today's business aircraft

owners. It was felt that utilizing both cost models would provide a better

understanding of the commuter aircraft market and the corporate aircraft

market combined. Thus, in the study, the modified ATA cost model is used

to represent commuter airline class General Aviation aircraft, and the

Corporate cost model is used to represent the typical business owned

aircraft. A complete, detailed description of both cost models is

contained in the Appendix.

19
4.0 CONFIGURATION DEFINITION

A preliminary configuration analysis was performed as part of the study to

select a suitable configuration to use for evaluation of propfan propulsion.

This section presents the results of this preliminary evaluation and

discusses the details of the configuration selected for the propulsion

system compari son.

4.1 CONFIGURATION SELECTION

As part of the preliminary evaluation, various configurations were

initially considered for the evaluation of propfan propulsion for this

study. Each configuration in Figure 12 was evaluated in terms of weight

and balance, mission capability, installation feasibility, stability and

control, and acoustic impact to confirm that propfan propulsion was

feasible for each of them. Each configuration layout used a similar

passenger cabin, discussed in Section 4.2. The .75 Mach cruise speed

requirement specified in the design ground rules required an engine that

was significantly larger than is currently available for General Aviation

class turboprops. As discussed in Section 4.3, a scaled PT6A engine was

selected. This larger engine, then, had a significant impact on achieving

a balanced configuration as well as increasing the basic operating weight.

Configurations 2 and 3 were similar aircraft configurations with different

engine installations. In Configuration 2, a pusher propfan was evaluated,

whereas in Configuration 3 a tractor propfan installation was evaluated.

Configurations 4 and 5 also utilized a common configuration with two

different propulsion system installations. Both configurations 4 and 5

were pusher installations. Configuration 5 had a straight-forward propfan

attachment to the engine/gearbox. Because of the engine weight and

subsequent balance requirements, the installation only allowed

approximately a 10 inch clearance between the wing trailing edge and the

propeller plane. This eliminated the possibility of having flaps on

Configuration 5, thus impacting the takeoff and landing field performance.

Configuration 4, on the other hand, had an engine installation which

incorporated a shaft between the gearbox and propfan, allowing an aft

extension of the propeller plane. With the propeller plane extended aft, a

flap system was incorporated and resulted in an improvement of field

performance.

As stated, each configuration was examined in sufficient detail to

establish that propfan propulsion was feasible in each case for the

mRIc_m_ PAR BLAm_ NOT Flta_D

21
1 2 5

4 5 6 7

FIGURE 12. Configurations Examined

selected mission. However, the following two requirements established by

NASA to provide a good propulsion system comparison narrowed the choice of

configurations which could be selected:

i. The requirement to assess the potential application of

propfan propulsion for business aircraft at .7 to .8 _lach

cruise speed.

. The requirement to compare the selected configuration to

a current day technology aircraft performing the same

mission.

The only configurations with current day counterparts capable of those

speeds were Configurations 2 and 3. Thus, Configurations 2 and 3 were

selected for those two reasons. Further, by selecting Configurations 2 and

3 the flexibility exists to compare current propulsion to both pusher and

tractor propfan installations without clouding the comparison by variation

of Configuration. Also, the selection provides a basis of comparison which

22
is representative of the currently established General Aviation fleet.

4.2 CABIN SELECTION

Based on the preliminary evalation, a typical small business jet cabin was

selected. It was designed using the configuration ground rules in Table 2.

Figure 13 shows the selected cabin features. The double club arrangement

shown does not provide the maximum seating capacity but, rather, provides a

typical comfort level which is representative of how most business aircraft

are configured. Special regulations under FAR 25 allow aircraft of this

class to have aisle widths between 9" and 12"; 11" was selected as typical.

TABLE 2. Cabin Design Criteria

• 57" Aisle Height

• 11" Aisle Width

• Flat Floor

• Lavatory

• Desired 56 Cubic Foot Baggage Volume

• Circular Cross Section on Fuselage Pressure Vessel

• Seating for Eight Passengers in a Double Club Arrangement

• 55" Between Seats

• 19" Seat Width

• Small Beverage Service Area

• All Fuel in the Wings

• Two Pilot Cockpit with Dual Instrumentation

23
FIGURE 13. Selected Cabin Layout

In Figure 13, the entry door is shown just aft of the cockpit with a small

baggage area across from the door. Some of the configurations required

that the door and baggage compartment be moved aft to facilitate wing and

propulsion system location. The head clearance, sidewall clearance, and

seat separation of the passengers are all shown for a 97.5 percentile man

(Reference 3), i.e. a height of 74".

24
4.3 ENGINE SELECTION

Pratt and Whitney, Canada (P&WC) developed three conceptual engine designs

in support of this study. Pratt and Whitney, Canada is a major engine

manufacturer of small General Aviation engines, such as those in this

study, and provides both turboprop and turbofan engines for numerous

aircraft in the General Aviation market (both business aircraft and

commuter airlines). This provided a single source of engine definition

assuring a consistent comparison of technology levels in this study. The

three baseline conceptual designs, a current technology turbofan, an

advanced technology turbofan, and an advanced turboprop for propfan

propulsion, were selected based on the following ground rules utilizing

P&WC's experience with this class of engines.

i. Each reference engine was sized to yield 650 LB of installed thrust

at .75 Mach and 41,000 FT altitude. The propfan efficiency for this

initial sizing was assumed to be 79%.

2. The current day turbofan was to utilize 1984 level engine

technology.

. The advanced engines were to incorporate technology consistent with

a 1988 propulsion readiness date.

. The advanced engines were to have equivalent technology to provide a

valid comparison between propfan propulsion and conventional

propulsion.

5. An engine scaling scheme was to be provided to allow each engine to

be sized for each study aircraft.

Because of the near term technology readiness date, P&WC did not expect a

radical departure from the current small, General Aviation engine design

practice. Rather, application of currently available technology would be

utilized and combined with the flexibility of optimizing the cycle on a new

design to yield substantial performance benefits while still achieving

reasonable cost levels.

Pratt and Whitney, Canada chose as a point of departure for the conceptual

designs the JT15D cycle for the turbofan engines and the PT6 cycle for the

propfan engine. The engine designs did not include the use of ceramics,

composites, or variable geometry because of the near-term technology

25
readiness date.

designs were:

However, key technologies that were considered in these

1. Use of advanced single crystal turbine blades permitting

increased cycle temperatures.

o Use of three-dimensional aerodynamic design methods yielding

improved pressure ratios, reduced pressure losses, and improved

component efficiencies.

o Use of advanced mechanical design technology giving a high degree of

structural optimization for minimum weight and complexity, more

efficient use of secondary air, better sealing techniques, and

improved tip clearance control.

4. Use of computer automated manufacturing technology to ensure

repeatability and improve cost.

4.3.1 TURBOFAN ENGINES

The 1988 turbofan engine selected was an advanced version of the JT15D

engine. A schematic is shown in Figure 14. The engine was a two spool

configuration similar to the JT15D-5 engine. The high pressure spool

consisted of a single stage centrifugal compressor driven by a single stage

uncooled axial turbine with the advanced single crystal blades. The

centrifugal compressor had a pressure ratio of 6.5:1 at 7.2 LB/SEC air flow

FAN & BOOST COMPRESSOR TURBINE

FIGURE 14. 1988 Technology Turbofan Schematic

26
with a rotor speed of _v,_n385 RPM. A reverse flow annular combustor,

designed for high efficiency (99%) and low emissions, was utilized. The

low pressure spool consisted of a high pressure ratio, high tip speed

single stage axial fan plus a core boost stage both driven by a two stage

fan turbine. The fan and core boost provided a pressure ratio of 2.73:1 to

the core at a rotor speed of 14,385 RPM. This resulted in a total

compressor pressure ratio of 17.8:1. The bypass ratio was 3.5 with a

bypass air flow of 25.3 LB/SEC at a pressure ratio of 1.88:1. The engine

also utilized a simple exhaust mixer so that core and bypass flows would

pass through a single nozzle thus providing an improvement in cruise

performance and reducing jet noise. The engine was 64 inches long with a

fan diameter of 24 inches. The engine weight was 695 LB.

The schematic in Figure 14 is also representative of the 1984 technology

turbofan used in this study. This current day engine was similar to the

JTI5D-5o The single stage centrifugal compressor had a rotor speed of

30,041 RPM and was driven by a single stage turbine. The core mass flow

was 9.1 LB/SEC. The fan was driven by a two-stage turbine at 16,104 RPM.

The combined compressor and fan pressure ratio was 16.5:1. The fan bypass

ratio was 1.93 with a bypass airflow of 17.5 LB/SEC. The 1984 turbofan

length was 63 inches long with a fan diameter of 21 inches and a weight of

650 LB.

Comparing the current and 1988 turbofan engine at the design point gives

the following: for the advanced engine, the bypass ratio is higher,

the pressure ratio is higher, the core air flow is lower, and the turbine

inlet temperature is higher. These were a direct result of the single

crystal blades and higher component efficiencies, which produced the same

thrust but at significantly improved SFC.

4.3.2 TURBOPROP ENGINE

The design of the 1988 propfan engine was based on the PT6 turboprop

configuration, upgraded with the new technology items, and reoptimized for

high altitude, high speed flight conditions. A schematic is shown in

Figure 15. The PT6 cycle was chosen, rather than the newer PWIO0 series,

because it was felt that this cycle was better suited for high altitude

optimization. Further, the PT6 configuration easily facilitates either

tractor or pusher installation with very little modification. The engine

utilizes a free power turbine approach. The gas generator consists of a

single spool compressor with three axial stages and one centrifugal stage,

driven by a single stage axial turbine with advanced single crystal

27
TOGEARBOX

COMPRESSOR TURBINE

FIGURE 15. 1988 Technology Turboprop Schematic

uncooled blades. It utilized a reverse flow, fully annular combustor for

high efficiency (99%) and low emissions.

The 1988 propfan engine was chosen to have a similar technology level to

the 1988 turbofan engine but was optimized slightly different. Because of

the additional compressor boost stage on the low pressure fan spool, the

turbofan provided a higher overall pressure ratio than did the

single-spool, multi-stage compressor design in the propfan engine. To

achieve a similar high pressure ratio in the propfan engine would have

required a two-spool design incorporating variable geometry due to turbine

loading and engine control requirements. This would have substantially

increased the cost and weight of the propfan engine as well as increasing

its technology level above the turbofan. As a result of the lower pressure

ratio of the propfan engine combined with its lower required rotor speed,

the propfan engine also provided the additional benefits of lower turbine

stresses and subsequent higher permissible turbine entry temperature.

The power section was comprised of a three stage axial power turbine which

drove the gearbox and a low pressure ratio exhaust nozzle. The gearbox

designed for the turboprop engine took into consideration the heavy flat

rating (sea level thermodynamic power of 2970 SHP flat rated to 1500 SHP).

This allowed a much lighter gearbox and gearing system to handle the torque

and SHP loads. The flat rating was selected for several reasons, as stated

in Section 5.1.1. The gear system was a growth step from the standard PT6

gearing, and utilized a two-stage, in-line planetary reduction geartrain to

give an output RPM of 1800.

The air flow of the 1988 turboprop engine was 5.8 LB/SEC at a pressure

ratio of 14.5:1 and a rotor speed of 28,136 RPM. The power turbine rotated

at 17,000 RPM and drove a gearbox which provided the 1800 RPM to the

28
propfan. The engine was 86 inches long (including accessories) and 26

inches in diameter. The engine weight was 725 LB including gearbox.

4.3.3 SUMMARYOF ENGINECHARACTERISTICS

Table 3 summarizes the characteristics of the engines selected for this

study. Tables 3 provides both the charactistics at the design point

(41,000 FT altitude and .75 Mach cruise speed) and the maximum

thermodynamic characteristics at the sea level, static condition. Each of

TABLE3. Engine Characteristics Summary

FLIGHTCONDITION S.L, STATIC,STDDAY

ITEM

FAN

FAN AIRFLOW (WA) LB/SEC

FAN BYPASS RATIO

FAN SPEED RPM

FAN THRUST LB

COMPRESSOR

CORE AIRFLOW (WA) LB/SEC

CORE THRUST LB

TURBINE INLET °F

TEMPERATURE

TOTAL AIRFLOW (WA) LB/SEC

UNINSTALLED THRUST LB

UNINSTALLED POWER* SHP

1984

TURBOFAN

59.4

2.09

15,896

1756(53%)

28.4

1533(47%)

1900

87.8

3289

1988

TURBOFAN

78.7

3.7

13,480

2393(73%)

21.4

887(27%)

2005

i00.I

3280

1988

PROPFAN

_B

_w

mm

18.1

317

2062

18.1

2970*

*Flat Rated to 1500 SHP installed

29
TABLE 3. Engine Characteristics Summary (cont'd.)

DESIGN POINT FLIGHT CONDITION .75M, 41,000 FT, STD DAY

ITEM

FAN

FAN AIRFLOW (WA) LB/SEC

FAN BYPASS RATIO

FAN SPEED RPM

FAN THRUST LB

COMPRESSOR

CORE AIRFLOW (WA) LB/SEC

PRESSURE RATIO-TOTAL

CORE THRUST LB

COMPRESSORTURBINE

PRESSURE RATIO

ROTOR SPEED RPM

INLET TEMPERATURE OF

HORSEPOWER

POWER TURBINE

PRESSURE RATIO

ROTOR SPEED

HORSEPOWER

TOTAL AIRFLOW (WA)

UNINSTALLED

RPM

LB/SEC

1984

TURBOFAN

17.5

1.93

16,104

236(35%)

9.1

16.5:1

445(65%)

3.03

30,041

1820

1,608

2.30

16,104

1,062

26.6

1988

TURBOFAN

25.3

3.5

14,385

356(51%)

7.2

17.8:1

343(49%)

2.82

30,385

1905

1532

3.56

14,385

882

32.5

1988

PROPFAN

m_

D_

DB

5.8

14.5:1

23

3.52

28,136

1962

1243

4.61

17,000

1170

5.8

THRUST LB

TSFC LB/LB/HR

FUEL FLOW LB/HR

INSTALLED

THRUST LB

TSFC LB/LB/HR

FUEL. FLOW LB/HR

681

.862

588

650

.882

573

699

.714

499

650

.736

479

722*

.605

438

650*

.634

412

*NOTE: Assumes 79% propeller efficiency at 41,000 FT, .75 Mach and

includes effects of exhaust thrust

30
these engines has been sized to produce the desired 650 LB installed thr,-st

necessary at the 41,000 FT, .75 Machcruise condition as shown in Table 3.

Table 3 also shows the uninstalled thrust levels. The assumedinstallation

losses were as follows:

• 8 LB/MIN bleed air flow

• 8 HP accessory loss

• 0.4% inlet total pressure loss for the turbofan engines

• 3.0% inlet total pressure loss for the turboprop engine

Comparison of the installed fuel consumption of each engine, in Table 3

indicates a 28% reduction for the propfan engine compared to a current

turbofan and a 14% reduction when compared to an equivalent technology

turbofan. Once again, these engines are matched to give the same thrust at

the design point. As the engines were scaled to meet the design mission

for each aircraft, the reduction in fuel consumption of the propfan

relative to the turbofan improves slightly. This is detailed in Sections

4.5, 5.1.1, and 5.7

4.4 PROPFAN SELECTION

Hamilton Standard provided the basic propfan parametric data for the study,

References 4 and 5. Previous propfan blade designs developed by Hamilton

Standard for single-rotation (SR) applications are given in Table 4. The

propfan blade selected for this study is illustrated in Figure 16 and is

described in Table 4. It was a six-bladed, 8 FT diameter, single-rotation

(SR) propfan. In the preliminary evaluation of the study it was apparent

that General Aviation propfans would be highly sensitive to propulsion

system acquisition cost. Because of this, every attempt was made to select

the propfan design which would provide the lowest possible cost without

compromising the design goals. Only six-bladed propfans were examined in

this study. Eight and ten-bladed propfans, as well as counter-rotation

propfans, were not examined due to their much higher acquisition cost.

Additionally, on the recommendation of Hamilton Standard, the blade sweep

was reduced from 350 to 150 to further lower the propfan acquisition cost

while not significantly impacting the performance.

31
TABLE4. Hamilton Standard Propfan Blade Designations

SR-1 a first cut design, 8 blades, 22o sweepat the tip,

AF of 204, 37.5 SHP/D2, .8M, 800 FPStip speed, 35,000 FT.

SR-2 a rectangular shaped, straight blade with 0o sweep, AF of

203, 8 blades, designed for 37.5 SHP/D 2, .8M, 800 FPS tip

speed, 35,000 FT.

SR-3 33o sweep, AF of 235, 8 blades, designed for 37.5

SHP/D 2, 800 FPS tip speed, .8M, 35,000 FT.

SR-4 new airfoil design/dropped by NASA.

SR-5 acoustic design, 10 blades, AF of 210, 46.5o sweep,

designed for 35,000 FT, 26 SHP/D 2, .8M, 600 FPS tip speed.

SR-6 designed by NASA, similar to SR-1 with a larger hub, 10

blades, AF of 204, 390 sweep, 30 SHP/D 2, 700 FPS, .8M,

35,000 FT.

SR-7 blade design for the Large-Scale Advanced Propeller (LAP)

for flight test on the Propfan Test Assessment (PTA)

Gulfstream II jet aircraft; 9 ft diameter, 8 blades, AF

of 227, 360 sweep, 800 FPS tip speed, 32 SHP/D 2,

.8M, 35,000 FT.

SR-8 blade design that incorporates new, thicker airfoil

sections; designed for .8M, 35,000 FT, 31SHP/D 2,

850 FPS tip speed; 9 blades, AF of 195, 430 sweep.

GA

STUDY

PROPFAN

blade design for General Aviation applications; designed

for .75M, 41,000 FT, 16 SHP/D 2, 750 FPS, 6 blades, AF

180, 15° sweep.

32
fl/ \\\

FIGURE 16. ISO-Views of the Selected Propfan

33
The performance was estimated utilizing the data package from Reference 5.

The data in Reference 5 represents an SR-3, ten-bladed propfan and was used

to approximate the performance of the selected six-bladed propfan. Several

check cases were made by Hamilton Standard for key flight conditions to

assure that this approach would yield valid results for the six-bladed

propfan used in this study. Table 5 shows that the data package generated

slightly higher propeller efficiency in the climb condition and slightly

lower efficiency in the cruise condition. The higher propeller efficiency

in climb resulted in slightly lower estimated fuel use in climb. This was

compensated, however, by the slightly higher estimated fuel use in cruise

due to the lower cruise efficiency such that, overall, the data package of

Reference 5 gave a valid fuel utilization estimate for the six-bladed

propfan.

The propfan disk loading and tip speed were selected based on the study's

mission requirements and ground rules. Hamilton Standard provided propfan

information for the study for tip speeds from 600 FPS to 800 FPS. As

stated previously, the 82 dBA goal was expected to be difficult to achieve

for the propfan. Propfan noise is reduced as the tip speed is reduced, but

with a subsequent increased cost and weight of the gearbox. On the other

hand, higher tip speeds allow a reduction in gearbox weight and cost plus

improved aerodynamic performance, but with significantly increased noise

levels. The minimum cost and highest performance would be achieved

TABLE 5. Comparison of Propfan Performance Estimates

FLIGHT CONDITION

SEGMENT

CLIMB

CRUISE

CRUISE

ALTITUDE SPEED

(FT) (MACH)

15,000 .40

30,000 .70

41,000 .80

PROPFAN

EFFICIENCY

SPOTA82

DATA PACK

SELECTED SIX

BLADE PROPFAN

.81 .80

.83 .84

.82 .83

34
utilizing an 800 FPS tip speed. At this tip speed, however, the cruise

noise levels were potentially high enough that the airframe structure would

be subject to significant sonic fatigue. Due to the uncertainty of the

noise calculations, a tip speed design margin was introduced. A tip speed

of 750 FPS was selected as a compromise to lower the potential of sonic

fatigue problems while still providing the higher performance and lower

cost of a high tip speed design.

The propfan disk loading (or propeller diameter) was selected for minimum

block fuel for the cost mission, Figure 17. Figure 18 shows the

performance of the selected propfan for the cruise condition design point.

It is interesting to note that the minimum mission fuel burn occurs near

the maximum efficiency of the propeller at about an equivalent disk loading

(SHP/D2). Other studies have indicated much higher disk loadings for

propfans. This is discussed further in Section 6.1.

COST

MISSION

BLOCK FUEL

LB

1500-

1460-

1420-

1380-

1340

10

CONDITIONS FOR SHP/D 2

41,000 FT, STD DAY,

.75M, 750 FPS 6.0

DIAMETER - FT //11

8.5 8 0 7.5

14' ' 1'8' 22' ' 26' '

SHP/D 2 FOR MAXIMUM CONTINUOUS - HP/FT 2

I

34

FIGURE 17. Effect of Propfan Disk Loading on Fuel Burn

35
84-

PROPFAN

EFFICIENCY

%

82

80  ̧

78

i0

CONDITIONS FOR SHP/D 2

41,000 FT, STD DAY,

• 50 FPS

i_4 l 118 ' 22' ' 26' ' 30' ' 3'4

SHP/D 2 FOR MAXIMUM CONTINUOUS - HP/FT 2

FIGURE 18. Effect of Propfan Disk Loading on Efficiency

4.5 PROPULSION SYSTEM SCALING

Propulsion system weight data provided by Pratt and Whitney, Canada and

Hamilton Standard was used to develop the propulsion system weight scaling

shown in Figures 19 through 23. The propulsion system weight includes the

weight of the engines, gearboxes, and props• Figure 19 shows the relative

weight effects of engine size for both the current technology and 1988

technology turbofans. Figure 20 shows the relative weight effects of

engine size for the 1988 technology turboprop at a gearbox output speed of

1800 RPM. The weight effect of change of RPM and impact of flat rating is

given in Figures 21 and 22. The propulsion system weight in Figures 19

through 22 assumes a constant propfan diameter• Figure 23 shows the effect

of propfan diameter change on propfan propeller weight•

The RPM of the propfan was selected to yield the chosen rotational blade

tip speed of 750 FPS. The optimum blade diameter selected as discussed in

Section 4.4, was eight feet and the RPM required to achieve the 750 FPS tip

speed was roughly 1800 RPM.

36
CHANGEIN

PROPULSION

SYSTEMWEIGHT

%

30.

20

10

1988 TECHNOLOGY

TURBOFAN

1984 TECHNOLOGY

TURBOFAN

-_o ' -lO 1o

FIGURE 19.

CHANGE IN

PROPULSION

SYSTEM WEIGHT

%

-20

20 l 30

CHANGE IN SLS THRUST

%

Turbofan Engine Size Effect on Propulsion System Weight

30

1800 RPM

20

10

' -1'o 1"o ' 20' ' 30'

CHANGE IN SLS HORSEPOWER

(THERMODYNAMIC)

%

-20

FIGURE 20. Turboprop Engine Size Effect on Propulsion System Weight

37
,

CHANGE IN

PROPULSION

SYSTEM WEIGHT

%

-_o ' -1'0 '

-2

-4

'0 ' 2'0 ' 3'0RPM

GE%IN

FIGURE 21. RPM Effect on Turboprop Propulsion System Weight

REDUCTION IN

PROPULSION

SYSTEM WEIGHT

%

RPM

)00 RPM

RPM

, !

50 6'0 7'0 80 90 1oo

PERCENT FLAT RATING - %

FIGURE 22. Flat Rating Effect on Turboprop Propulsion System Weight

38
CHANGE IN

PROPFAN WEIGHT

%

20

10

FIGURE 23.

10 2'0

CHANGE IN PROPFAN DIAMETER

%

Propfan Diameter Effect on Propeller Weight

It was assumed that the engine envelope dimensions of the baseline engines

did not change with the small amount of engine scaling required for the

study aircraft. A nominal length of 64 inches was used for the turbofans

with a diameter of 31 inches. The turboprop engines (exclusive of the

propfan itself) had a nominal length of 82 inches overall (including

gearbox and accessories) and a diameter of 26 inches. These dimensions are

consistent with the current PT6 and JT15D series engines.

39
5.0 CONFIGURATION EVALUATION

Utilizing the selected general configuration of Section 4, this section

presents the detailed sizing and optimization of four aircraft; a 1984

technology turbofan aircraft, a 1990 technology turbofan aircraft, a 1990

pusher propfan aircraft, and a 1990 tractor propfan aircraft. The mission

performance and cost of operation results are presented.

5.1 CONFIGURATION SIZING AND OPTIMIZATION

Figures 24 through 27 are the carpet plots that were used to size and

optimize the final study configurations. The process to specify each

configuration (design point A in each figure) was similar to that described

in Section 3.2

Each design point was selected based on two criteria. First, the aircraft

design point had to meet all the performance capabilities specified in the

study ground rules. Then, second, if possible, the design point had to

provide the minimum fuel burn for the cost mission described in Section

3.1.

In Figures 24 and 25, both the current turbofan and advanced turbofan were

designed by performance constraints. Point A in each case was constrained

by a 2.4% climb gradient and a cruise Mach number of .75 at 41,000 FT

altitude. Point A, then, was the point which satisfied both design

criteria.

The pusher and tractor propfans, on the other hand, were only constrained

in performance by the cruise requirement. Thus the design point was the

minimum fuel point along the constant cruise Mach line shown in Figures 26

and 27. Point A, in addition to meeting the required cruise speed, also

exceeded all the other performance requirements while providing the minimum

fuel usage for the propfans.

5.1.1 PROPULSION SYSTEM PERFORMANCE

As stated in Section 4, the Pratt and Whitney PT6 engine cycle was used as

a baseline to provide scaled turboprop data for the study. The Pratt and

Whitney JT15D engine cycle was used as a basis for the turbofan data. A

flat rating of 1500 SHP was imposed on the turboprop engine at lower

altitudes. This was done to relieve some of the installation problems
18B00

18600

IB400

18200'

IBO00"

17800'

17600,

TAKEOFF

WEIGHT - LB 17400

17200

17000

16800

ISGO0

16400

IS200-

IBO00

TOFL = TAKEOFF FIELD LENGTH (FAR 25) - FT

GRAD = SECOND SEGMENT CLIMB GRADIENT - %

Mcr = CRUISE SPEED AT 41,000 FT - MACH

BLK FUEL = COST MISSION FUEL REQ'D - LB

5 ISSIOB "

g_SStOg

_- GRAD 2.4

...... _6_

16050

15900

15750.

15600.

15450.

15300,

15150,

TAKEOFF

WEIGHT - LB 15000,

14850

14700

14550

14400

14250

14100

13950.

FIGURE 24. Current Turbofan Carpet Plot

TOFL = TAKEOFF FIELD LENGTH (FAR 25) - FT

GRAD = SECOND SEGMENT CLIMB GRADIENT - %

Hcr : CRUISE SPEED AT 41,000 FT - MACtl

BLK FUEL = COST MISSION FUEL REQ'D - LB

\

%

\

TOFL 4000 GRAD 2.4

FIGURE 25. Advanced Turbofan Carpet Plot

42
16050-

15900-

15750-

15600-

15450-

15300

15150

TAKEOFF

WEIGHT - LB 15000

14850

14700

I4550

14400

14250

14100

13550

16650.

16500,

16350.

16200,

16050.

15900.

TAKEOFF 15750

_EIGHT - LB 15600

15450

15300

15150

15000

14850

14700

14550

FIGURE 26.

FIGURE 27.

TOFL : TAKEOFF FIELD LENGTH (FAR 25) - FT

GRAD = SECOND SEGMENT CLIMB GRADIENT - %

Mcr = CRUISE SPEED AT 41,000 FT - MACH

BLK FUEL = COST MISSION FUEL REQ'D - LB

516

Z930

Mcr .75

TOFL 4000 6. 0

Pusher Propfan Carpet Plot

TOFL = TAKEOFF FIELD LENGTII (FAR 25) - FT

GRAD = SECOND SEGMENT CLIMB GRADIENT - %

Mcr = CRUISE SPEED AT 41,000 FT - MACH

BLK FUEL = COST MISSION FUEL REQ'D - LB

Mcr .75

TOFL 4000 _ 6.0

Tractor Propfan Carpet Plot

43
THRUST

LB

1800

1600

140C

CONDITIONS

41,000 FT, STD DAY

MAXIMUM CRUISE SETTING

1200 !_FAN THRUST

i000 _

z-1988 TURBOFAN THRUST _ .....----"_/

200 ..... _ PROPFAN TSFC

i i ! I i I

o .'t .2 .5 .'4 ' "s ' :6 ' ;7

HACH NUMBER

•I .8

1.6

1.4

1.2

TSFC

LB/HR/LB

1.0

,8

,4

.2

0

.8

FIGURE 28. Sized Engine Characteristics

44
associated with the propfan such as high blade stresses, engine weight,

gearbox weight, cost, airframe weight, and aircraft tail size. Figure 28

compares the current 1984 turbofan engine, the advanced 1988 turbofan

engine, and the advanced 1988 propfan engine, all on a thrust basis. The

propulsion systems provided by Pratt and Whitney were all sized initially

to provide 650 LB of thrust at the design point. Wheneach aircraft was

then sized to meet the design mission, different thrusts were required at

the design point because of configuration weight and drag differences, as

shown in Figure 28. The propfan thrust shown is for the pusher

installation and includes the effect of the engine exhaust thrust as well

as the propeller thrust. The tractor installation is not shown, but

differs only slightly from the pusher data. The thrust for the tractor

installation was approximately 3%higher than that of the pusher but with

the sameTSFC. The TSFCcurves show that, at the cruise design point of

.75 Mach and 41,000 FT altitude, the propfan had 31%less fuel consumption

than a current turbofan and 17% less fuel consumption than an advanced

turbofan.

Propulsion system scaling utilized the baseline engine data provided by

Pratt and Whitney. The baseline thrust or power data was scaled linearly

while keeping the specific fuel consumption the sameas the baseline. This

approximation was valid over the narrow band of scaling required for this

study. The weight scaling was accomplished using the methods outlined in

Section 4.5.

5.1.2 CONFIGURATIONTHREE-VIEWS

Figures 29, 30, 31, and 32 show the 3-views for the final study aircraft.

The aircraft shownrepresent the aircraft characteristics of Design Point A

for each of their respective carpet plots in Figures 24 through 27. Each

aircraft satisfied all of the initial design ground rule requirements and

provided the minimum block fuel burn for the cost mission defined in

Section 3.1.

45
TOGW

OEW

PAYLOAD

USABLE FUEL

WING AREA

SPAN

ASPECT RATIO

WING LOADING

THRUST/ENGINE, SLS

17033 LB

8676 LB

1200 LB

7157 LB

349.0 FT 2

54.5 FT

8.5

49 LB/FT 2

3257 LB

54.5 FT

56.0 FT

FIGURE 29. 1984 Current Turbofan Aircraft

46
TOGW

OEW

PAYLOAD

USABLE FUEL

WING AREA

SPAN

ASPECT RATIO

WING LOADING

THRUST/ENGINE, SLS

14234 LB

7426 LB

1200 LB

5608 LB

248. i FT2

45.9 FT

8.5

57 LB/FT 2

3067 LB

45.9 FT

_) (_) - _:

54.4 FT

FIGURE 30. 1990 Advanced Turbofan Aircraft

47
TOGW

OEW

PAYLOAD

USABLE FUEL

WING AREA

SPAN

ASPECT RATIO

WING LOADING

SHP/ENGINE, SLS

14230 LB

8274 LB

1200 LB

4756 LB

252.2 FT 2

46.3 FT

8.5

56 LB/FT 2

2714 THERMODYNAMIC

1500 FLAT

46.3 FT _-

55.5 FT =-

FIGURE 31. 1990 Pusher Propfan Aircraft

48
TOGW

OEW

PAYLOAD

USABLE FUEL

WING AREA

SPAN

ASPECT RATIO

WING LOADING

SHP/ENGINE, SLS

14805 LB

8802 LB

1200 LB

4803 LB

260.8 FT2

47.1 FT

8.5

57 LB/FT 2

2693 THERMODYNAMIC

1500 FLAT RATED

47.1 FT --

O0

9 FT---_

8 FT

_ DIAMETER

111

I- 55.5 FT

FIGURE 32. 1990 Tractor Propfan Aircraft

49
5.2 INSTALLATIONCONCEPTS

Since the study goal was to provide a good comparison between conventional

propulsion and propfan propulsion, an attempt was made to select similar

installation concepts. Figure 33 shows a typical conventional turbofan

installation for a Pratt and Whitney JT15D class engine. The major loads

were carried by the engine mount support beam. The engine was stabilized

by a "steady-rest" support tied to the aft portion of the engine. The

accessories were provided cooling air through a small duct and were mounted

below the engine.

Figure 34 shows the installation for the tractor version of the propfan

using a Pratt and Whitney PT6 engine. Figure 35 shows the pusher version

installation. Similar to the turbofan, the major loads were carried by the

mount support beam for both propfan versions. However, because of the

higher propulsion system weight and slightly modified geometry, the support

beamfor the propfan versions was larger than the beam required for the

turbofan. The propfan propulsion system weight was roughly 30%more than

the turbofan and was located further from the fuselage for propeller

clearance. To carry the additional loads, the propfan installation

required a spar cap of roughly twice the area of that required for the

turbofan. The larger spar, combined with the effect of a longer length

pylon, resulted in a significant pylon weight increase over the turbofan

installation. Combining this with the structural requirements for

gyroscopic loads, whirl induced loads, and sudden prop stoppage (FAR

25.361(b)(1)) resulted in a pylon/nacelle weight increase of roughly 34%

for the pusher installation and 78% for the tractor installation, as

comparedto the turbofan installations.

The engine in both the pusher and tractor versions was stabilized by a

"steady-rest." This steady-rest had to be capable of handling whirl

induced loads. Also, the engine case itself required additional strength

to handle the additional propfan loads.

De-ice was accomplished using engine bleed air on the engine inlet lips for

the turbofan. De-ice was also provided on the inlet lips for both

turboprop installations. Also, as illustrated in Figures 34 and 35,

anti-ice protection for the engine intake screen was provided by the

inertial separators. The inertial separators are a mechanical vane and

bypass duct arrangement which forces the heavier, super-cooled water and

ice particles to bypass the inlet plenum and thus prevents the engine inlet

50
"STEADYREST" SUPPORT

\

I • O

ENGINE

MOUNT

FORGING

VIBRATION

ISOLATOR

ACCESSORY GEARBOX

-_9_ FWO

STARTER-GENERATOR ,,_ _ _

COOLING AIR INLET _'""--__

/------l---

STARTER-GENERATOR I

FIGURE 33. Turbofan Installation Concept

51
OIL COOLER AIR INLET

OIL COOLER

EXIT

ENGINE STEADY

REST

ENGINE MOUNT

SUPPORT BEAM

STARTERGENERATOR

_E-PORT EXHAUST STACK

ENGINE MOUNT

FORGING _ STARTER-GENERATOR

\ /COOLING AIR INLET

_LER \ / I

VIBRATION/ _ _ " AIR INLET

/

ISOLATOR

--III-----FWD

R GOVERNOR

NI

FIGURE 34. Tractor Propfan Installation Concept

52
\

.,,J

j_

ENGINE MOUNT SUPPORT BEAM

OIL COOLER INLET

OIL COOLER

INERTIAL SEPARAT(

"STEADYREST" MOUNT

ENGINE INDUCTION

AIR INLET

REDUCTION GEARBOX

ENGINE A

MOUNT FORGINGI _ __

OIL _ \ \11 IV /

_YLO_#,.._ /_ \

VIBRATION ' / / \__ J.,/ \

iSC)LATOR_ _ _ )

J

_,_ FWO

ACCESSORY GEARBOX _

FIGURE 35. Pusher Propfan Installation Concept

53
screen from becoming blocked. It was assumed that the propfan blades in

the pusher installation would not require any ice protection because of the

hot exhaust gas flowing into the blades. Ice protection for the propfan

blades in the tractor installation was not examined in this study.

Two different exhaust stack concepts are illustrated in Figures 34 and 35.

The tractor version utilized a single port exhaust and an under nacelle

"smile" type inlet. The pusher version used a bifurcated inlet located in

the pylon and dual exhaust ports. It was felt that either of these inlet

and exhaust concepts could have been used interchangeably with either the

pusher or tractor installations. However, the dual exhaust port had

several advantages over the single port exhaust. The large single port

exhaust presented a much larger base drag than the dual port. Also, with

the larger exposed area, the single port exhaust's potential for shocking

was greater. In a pusher installation, the formation of shocks, in

addition to an increase in drag, could result in significant cyclic loads

on the propfan blades. Additionally, in the pusher configuration, the dual

port would provide better hot exhaust gas mixing and thus lower

temperatures prior to entering the prop plane. Detailed analysis of the

exhaust stack impact on the flow field was beyond the scope of this study.

However, a simplified potential flow model, Figure 36, was constructed to

verify that a fairing could be designed which would prevent shock formation.

Figure 36 shows a fairing which remains just slightly subsonic for the

design cruise condition, as illustrated in the graph in Figure 36. This

fairing was a rough cut, first attempt design for the pusher installation.

It was felt that with further analysis and tailoring a fairing could easily

be designed which would remain well below sonic and thus yield a shock-free

design.

54
M_ = .75

MACH = 1

X =1_

FIGURE 36. Effect of Exhaust Stack in the Propfan Flow Field

55
5.3 AERODYNAMIC CHARACTERISTICS

The selected configurations utilized conventional aerodynamic design. The

wing plan form for each was chosen based on the results from the aircraft

synthesis programs. The wing section for each design was based on an

airfoil section which eliminated any transonic effects for the aircraft's

mission requirements. The airfoil section at the root was 18% thick. The

airfoil section at the tip was 15% thick.

5.3.1 LOW SPEED DRAG CHARACTERISTICS

The low speed drag polars for each configuration are shown in Figures 37

through 40. Each includes the effects of propulsion system installation.

The high speed Mach effects are minimal and are discussed in Section 5.3.3.

The drag shown is for twin engine cruise and takeoff. Single engine

(engine-out) takeoff drag is 3% higher for the turbofan installations and

9% higher for the propfan installations than the twin engine takeoff drag.

The higher single engine drag accounts for the effects of yaw, trim,

feathered propeller, and windmilling turbine. Taking into account the

1.40

1.20

..J

__)

1.00

I

I--

Z

i,i .80

H

U

I--I

b..

b..

W .60

0

la_ .40

b.4

_J

.2O

0

.02 .03 .04 .05 .06 .07 .08 .09 .10

DRRG COEFFICIENT - C D

FIGURE 37. Current Turbofan Low Speed Drag Characteristics

56
1.40 1

1.20

._I

1.00

l

I--

Z

Ld .80

I--I

U

b..

b.

W .60

0

U

I--

h .40

IiI

._1

0

• 02 . 03 . 04 , 05 . 06 . O? . 08 . 08

DRRG COEFFICI_-NT - C D

.IO

FIGURE 38. Advanced Turbofan Low Speed Drag Characteristics

I .40

..J

(iJ

I

I--

Z

i,i

I--I

U

IiI

h

h

W

0

U

II

I,

I--I

.J

1.20

I. O0

.80

.60

.40

.2O

ol.02

• • , • , , • , , , , , , l ,

.03 .04 .05 .06 .07 .08 .08 .10

DRRG COEFFICIENT - C D

FIGURE 39. Pusher Propfan Low Speed Drag Characteristics

57
i .40

.J

U

I

tZ

W

H

U

H

h

h

W

0

U

h

H

.J

I .20

I. O0

.80

.SO

.40

.2O

0

• 02

l, | i i K , , | _ | , , | i |

.03 .04 .05 .OS .07 .08 .0@ . I0

DRAG COEFFICIENT - C O

FIGURE 40. Tractor Propfan Low Speed Drag Characteristics

differences in reference wing area, the pusher propfan showed a 4.7% higher

drag than the advanced turbofan. Also, the tractor installation was 2%

higher drag than the pusher installation, due to the larger pylon required

and the effect of the prop wash.

5.3.2 STABILITY AND CONTROL CONSIDERATIONS

The horizontal tail of each study aircraft was designed to provide adequate

center of gravity range for typical business aircraft payloads. Each

design incorporated a variable incidence horizontal tail. Because of the

aft fuselage mounted arrangement of the propfan configurations, the propfan

propellers provided additional longitudinal stability when compared to the

turbofans. At first glance this would seem desirable. However, the

stabilizing effects of the larger pylons, as well as the stabilizing

effects of the propellers, shifts the usable CG range aft with the same

horizontal tail as the turbofan. Further, the increased stability makes

the nose wheel liftoff the limiting factor rather than the stall in ground

effect or approach trim condition. This new aft limit is aft of the main

landing gear tip-back limit for the propfans, making some of the CG range

unusable. It was not possible to move the landing gear any farther aft and

58
remain in the structure. Also, lengthening of the fuselage was not an

option because a commoncabin configuration was desired. Therefore, a

larger horizontal tail was utilized for the propfans for increased control

power. This provided a CG limit far enough forward to give the desired CG

range while providing the required nose wheel liftoff capability.

Table 6 shows the horizontal tail characteristics for each configuration.

TABLE 6. Horizontal Tail Characteristics

CONFIGURATION

C.G. RANGE %

TAIL VOLUME

TAIL AREA FT2

CURRENT ADVANCED PUSHER TRACTOR

TURBOFAN TURBOFAN PROPFAN PROPFAN

14.0

.78

74.9

18.5

1.00

61.2

15.0

1.11

63.4

15.0

1.05

63.4

The larger vertical tail size of the propfan configurations resulted

primarily due to the requirement for engine-out control. The vertical tail

characteristics are shown in Table 7. Several steps were taken to reduce

the impact of the propfan installation on the vertical tail size. First, a

flat rating of 1500 SHP was chosen to help reduce the thrust available for

the engine-out case. Second, the vertical position of the thrust

centerline was adjusted separately for the pusher and tractor versions so

that the thrust line's horizontal position remained in the same position.

Third, a larger, more effective rudder was incorporated in the propfans'

vertical tails.

59
TABLE7. Vertical Tail Characteristics

CONFIGURATION

THRUST MOMENTARM IN

VERTICAL TAIL VOLUME

VERTICAL TAIL AREA FT2

RUDDER SIZE (Cr/C v)

CURRENT ADVANCED

TURBOFAN TURBOFAN

55

.06

60.8

.35

55

.09

60.8

.35

PUSHER

PROPFAN

84

.12

71.1

.45

TRACTOR

PROPFAN

84

.11

71.1

.45

5.3.3 HIGH SPEED CONSIDERATIONS

The design cruise speed for the study aircraft was .75 Mach. This speed

was well below speeds in which significant transonic drag interference

effects are typically encountered, provided the designs are relatively

clean. The conventional configurations selected for this study could

easily be tailored to prevent these undesirable drag effects. The

turboprop engine exhaust stacks were of particular concern. To verify that

it was possible to tailor the stacks to prevent shock formation, the

installation was examined in Section 5.2 . The potential flow analysis

indicated that no shock effects were present. The wing plan form sweep was

chosen to prevent any possibility of shock formation which would affect the

flow field into the propfan blades. The high speed drag characteristics

for each aircraft are shown in Figures 41 through 44.

6O
1

U

I

lZ

Ld

H

IJ

H

Is_

h

W

0

U

I--

h

I-,I

_1

.70

.60

.50

.40

.30

.20

.10

$ @

0

•020 .024 .028 .032 .036 .040 . 044 .048

DRRG COEFFICIENT - Cn

.052

FIGURE 41. Current Turbofan High Speed Drag Characteristics

_J

U

.50

I

lZ

b3 .40

H

U

H

b.

b.

L,.J .30

0

U

lb. .20

H

_I

0

.020 .024 .028 .032 .036 .040 .044 .048

nRRG COEFFICIENT - C D

• 052

FIGURE 42. Advanced Turbofan High Speed Drag Characteristics

61
.7O

J

U

I

IZ

W

I--I

u

t,,-,,I

h

h

W

0

ID

I--

h

I-,.t

._1

.6O

.50

.40

.30

.20

.I0

0

•020 .024 .028 . 032 .036 .040 .044 .048

DRAG COEFFICIENT - C D

.052

FIGURE 43. Pusher Propfan High Speed Drag Characteristics

0

•020 .024 .028 .032 .036 .040 .044 .048

DRAG COEFFICIENT - C D

.O52

FIGURE 44. Tractor Propfan High Speed Drag Characteristics

62
5.4 WEIGHT AND BALANCE

A weight summary is given in Table B for each of the configurations. The

propulsion system group total is shown in parentheses and a breakdown of

the major elements are given below. All of the configurations studied were

twin engine aircraft, and the propulsion system elements shown are the

totals for both propulsion units on each aircraft. Since both propfans

were assumed to have full reverse capability for landing, thrust reversers

were incorporated on the turbofans to provide equivalent capability.

Gearbox and engine weights were provided by Pratt and Whitney, Canada.

Propfan weights were provided by Hamilton Standard. More discussion of the

propulsion system weights and installation can be found in Sections 4.3 and

5.2.

The current turbofan aircraft represents current technology aluminum

construction. The 1990 technology level turbofan and propfan aircraft

reflect the application of advanced materials. As discussed in Section

5.5, advanced airframe materials and construction represent a 26% reduction

in airframe structural weights (References 6, 7, and 8). The turbofan

aircraft had sufficient fuselage structural damping material and insulation

to achieve the required 82 dBA cabin noise level. The propfans

required additional acoustic treatment to achieve this level. The weight

of this additional material is shown as a separate item in Table 8. This

material for the pusher propfan is 1.5% of the gross weight; 4% of the

gross weight for the tractor propfan. Further discussion of these numbers

and acoustic effects can be found in Section 5.6.

The loading diagrams for each study aircraft are shown in Figures 45 to 48.

A typical loading for both a maximum forward CG and a maximum aft CG at

maximum gross weight is shown. The envelope limits were set by the

criteria discussed in Section 5.3.2.

63
TABLE 8. Configuration Weight Summary

(LB)

GROUP

WING

TAIL

BODY

GEAR

NACELLE

PROPULSION

ENGINE (2)

GEARBOX (2)

PROPELLERS (2)

SYSTEMS (2)

THRUST REVERSERS (2)

FLIGHT CONTROLS

INSTRUMENTS

HYDRAULICS

ELECTRICAL

AVIONICS

FURNISHINGS

ADDITIONAL ACOUSTIC MATERIAL

ENVIRONMENTAL/ANTI-ICE

OTHER/MISC.

EMPTY WEIGHT

FULL FUEL

PAYLOAD

GROSS WEIGHT

CURRENT ADVANCED PUSHER TRACTOR

TURBOFAN TURBOFAN PROPFAN PROPFAN

1569

391

1165

681

254

(1901)

1299

402

200

293

135

121

366

484

883

359

74

8676

7157

1200

17033

927

288

806

600

189

(1902)

1300

m_

402

2O0

293

135

121

366

484

883

356

76

7426

5608

1200

14234

910

318

844

600

267

(2412)

1221

232

478

481

274

135

121

366

484

883

210

357

93

8274

4756

1200

14230

944

318

824

622

343

(2412)

1221

232

478

481

274

135

121

366

484

883

625

357

94

8802

4803

1200

14805

64
LEGEND

FORHRRD(FHD)

1 CREW

2 FRONT PRSS

3 MID PRSS

4 BRGGRGE

5 FULL FUEL

RFT

6 CREW

7 FULL FUEL

m

i

I

FT

O

I.-I

I,I

I

19000

18000

17000

IBO00

15000

14000

13000

12000

II000

I0000

8000

BOO0

7000 .... ' ......... ' .... "

F'IdD RFT

E......... ' ........... l" ............. ' ......... ]

I

I

l_

..........I........)_._.

305 310 315 320 325 330 335 340 345 350 355

CENTER OF GRRVTTY------F.S.--IN

FIGURE 45. Current Turbofan Loading Diagram

LEGEND

FORHRRD(FHD)

I CREW

2 FRONT PRSS

3 MID PRSS

4 BRGGRGE

5 FULL FUEL

RFT

6 CREH

7 FULL FUEL

m

J

I

-r

L9

H

Ld

3[

IGO00

15000

14000

13000

12000

II000

I0000

9000

8000

7OO0

6000

FHD RFT

.....'...'........t......•....._..................

I I

I I

!/ r!

l/ i!

............i.__,_

305 310 315 320 325 330 335 340 345 350 355

CENTER OF GRRVITY------F.S.--IN

FIGURE 46. Advanced Turbofan Loading Diagram

65
LEGEND

FORHRRD(FHD)

I CREW

2 FRONT'PRSS

3 HID PRSS

4 BRGGRGE

5 FULL FUEL

RFT

6 CREW

7 FULL FUEL

g:l

_1

I

I-

-r

(3

H

W

.-¢

16000

15000

14000

13000

12000

11000

10000

9000

FHI3 RF'T

! !

I

J!

f!

BOO0

,ooo.................... I......... ,1.......)_."....

295 300 305 310 315 320 325 330 335 340 345

CENTER OF GRRVITY----F. S.--IN

FIGURE 47. Pusher Propfan Loading Diagram

LEGEND

FORHRRD(FWD)

I CREW

2 FRONT PASS

3 MlO PRSS

4 BRGGRGE

5 FULL FUEL

RFT

6 CREW

7 FULL FUEL

m

J

I

FI

O

H

W

".r

16000

15000

14000

13000

12000

II000

I0000

SO00

8000

FHD RFT

........... •"' ....'1........... i.....'.........

I I

I I

 ooo ....................... ,........... L.... ?:i..

295 300 305 310 315 320 325 330 335 340 345

CENTER OF GRRVITY------F.S.--IN

FIGURE 48. Tractor Propfan Loading Diagram

66
5.5 STRUCTURAL DESIGN

The weight and cost estimates for the study were based on the use of two

levels of material technology. The structural design criteria for all the

study aircraft was based on FAR 25.

5.5.1 STRUCTURAL MATERIAL

The current technology turbofan's airframe utilized an all aluminum

construction. The advanced 1990 level technology aircraft assumed the use

of aluminum, aluminum-lithium alloys, and composite materials to acheive

the best possible weight reduction. A detailed structural analysis was

beyond the scope of this study. However, based on current NASA and

industry understanding (References 6, 7, and 8) weight reductions of up to

30% (Reference 8) are possible on major structural components. It was

assumed, for this study, that utilizing these advanced materials in the

structural design would provide a 26% reduction in airframe weight. For

this class of aircraft, the airframe structural weight ranges between 38%

to 42% of the total aircraft empty weight. Thus the resultant improvement

due to this 26% reduction in airframe weight would be approximately, a 10%

reduction in total un-resized empty weight.

The propfan installations required additional structural beef-up to carry

the higher loads and prop effects. The installations shown in Section 5.2

all have a similar mount designs. However, the higher propfan loads

required a spar cap with roughly twice the area of that required for the

turbofan installations. The fuselage frames also required strengthening to

handle the additional loads, and the fuselage skin was thickened for

protection from possible sonic fatigue in the region of the prop plane.

5.5.2 FAR 25 SAFELIFE REQUIREMENTS

An area of concern which was not evaluated in this study was the potential

structural hazard of catastrophic failure of the propfan propulsion system.

Hamilton Standard's structural criteria for both the propfan blade spars

and hub is to design for infinite component life, but work needs to be done

relative to shed blade damage to the aircraft structure. Recently

certified General Aviation turbofans, certified under FAR 25, have had to

meet a special rotor burst condition. This condition states that a

projectile from a rotor burst traveling forward at a 5 degree angle

relative to the rotor face cannot penetrate either the fuselage pressure

67
vessel or aircraft fuel tank. Additionally, FAR 25.901 requires that no

single failure or malfunction can jeopardize the safe operation of the

aircraft, including potential structural damage from shed propfan blades.

FAR 23 (Airworthiness Standards: Normal, Utility, and Acrobatic Category

Airplanes) does not include this structural provision. Rather, it only

requires that a propulsion system failure be isolated from the other

propulsion systems of the aircraft so that it will not prevent the

continued safe operation of the remaining engines. This is an area which

requires further work before propfan propulsion can be implemented on

General Aviation aircraft.

5.6 ACOUSTIC EFFECTS

In the initial phase of the study it was clear that, for propfan propulsion

to be competitive with turbofan propulsion, the propfan powered aircraft

would have to provide an equivalent level of comfort in the cabin and

cockpit. The interior noise level significantly impacts the perceived

level of comfort, and thus, the propfan interior noise level would need to

be as close as possible to the level of comfort of the turbofan. Recent

work indicated that the annoyance threshold of passengers to noise due to

the boundary layer occurs at 82 dBA (Reference 9). The annoyance

threshold, as defined in Reference 9, is the A-weighted level at which 50%

of the cabin occupants were annoyed. This annoyance threshold level of 82

dBA was chosen as the maximum acoustic level design point for this study.

Additionally, 82 dBA is a typical interior noise level for this class of

small turbofans. All of theaircraft interior and structural weights

were estimated based on this average cabin noise level.

Two primary sources of aircraft interior noise were considered in this

study; the propulsion system and boundary layer. The fuselage structural

treatment was based on the mass law technique of Reference 10. The

boundary layer noise was estimated using the method defined in Reference 11.

The propfan external noise distribution was predicted using the methods

contained in Reference 4 and 5.

Figures 49 and 50 show the predicted external sound pressure levels for the

propfan aircraft as a function of fuselage station. The blade passage

frequency for the 8 FT diameter propfan examined in this study was 179 HZ.

The boundary layer noise shown in these figures is for the design cruise

point of .75 Mach at 41,000 FT. Cabin interior noise data available on the

Beech Model FJ400 test vehicle (an experimental prototype turbofan tested

by Beech in 1974) was used, combined with the method of Reference 11, to

predict the dominant boundary layer effect and the resultant exterior

68
EXTERIOR

OASPL

(riB)

150F

140

130

120

Ii0

10 1 00 250I 300I 350I 400

FUSELAGE STATION (IN)

TREATMENT APPLIED FOR

POSSIBLE SONIC FATIGUE_

ADDITIONAL TREATMENT APPLIED /

FOR STRUCTURE BORN NOISE /

BOUNDARY LAYER N_OISE _ /

IN 250 HZ, 1/3 OCTAVE BAND/_ /

AFT PRESSURE BULKHEAD

41,000 FT,

I i I

150 200 450

.75 M

560

FIGURE 49. Pusher Propfan Installation Acoustic Loading

EXTERIOR

OASPL

(dB)

150

140

130

120

110

ACOUSTIC TREATMENT

OF 9,54 LB/FT2 APPLIED

AT PEAK NOISE LEVEL

BOUNDARY LAYER NOISE

IN 250 HZ, 1/3 OCTAVE BAND

i° 'oo

FUSELAGE SECTION

WITH ADDITIONAL

ACOUSTIC TREATMENT

AFT PRESSURE BULKHEAD

41,000 FT, ,75 M

2_o 360 3_o 460 4_0 5bo

FUSELAGE STATION (IN)

FIGURE 50. Tractor Propfan Installation Acoustic Loading

69
boundary layer sound pressure levels, as shown. The boundary layer noise,

together with the propeller noise first harmonic, was then plotted versus

fuselage station in Figures 49 and 50. Figure 49 shows the acoustic

loading for the pusher propfan installation. The location of the prop

plane in this installation lies well aft of the cabin pressure vessel

bulkhead. The passenger cabin environment receives no directly radiated

exterior noise input from the propfan blades. The only acoustic treatment

necessary then, is that required to dampout the boundary layer noise and

the additional structural born noise generated by the propfan and engine.

Both the turbofan and propfan fuselage furnishing weight estimates shown in

Section 5.4 incorporate the necessary acoustic treatment to reduce the

boundary layer noise level and structure born noise of the turbofan to

achieve the 82 dBA goal. The additional weight required for the pusher

propfan's additional structure born noise was estimated using Reference 10

and data from the FJ400 test to be 210 LB or 1.5% of the maximumgross

weight. This weight also incorporated the treatment for protection from

sonic fatigue in the region of the prop plane.

The tractor version of the propfan required muchmore cabin treatment than

the pusher. The prop plane in this installation was farther forward and

was located at the pressure bulkhead. Because of this the passenger cabin

exterior was exposed to higher noise levels and more treatment was

necessary. The fuselage location requiring additional treatment is

illustrated in Figure 50 by the shaded area. Figure 51 shows that a

maximum of 9.54 LB/FT2 was required for this installation. The additional

acoustic treatment required for the tractor propfan installation was 625 LB

or 4%of the maximumgross weight.

The directivity of the propfan noise was predicted using the methods of

Reference 4 and 5. These methods are based on data for tractor propfans.

Other methods such as Reference 12 and work done at NASA-Langley

(References 13 and 14) indicate that the shape of the directivity curve for

pusher propellers could be much different than tractor propellers. The

difference is illustrated in the sketch in Figure 52. If this is indeed

correct it would meansignificantly higher noise treatment for pushers than

is currently indicated in this study. However, the directivity of noise

for pushers has not been sufficiently examined at this time to provide a

reliable noise distribution estimate. More work needs to be done in this

area. Therefore, for this study, the methods in Reference 4 and 5 were

assumedvalid for both tractors and pushers.

70
0

9.54 LB/FT 2

ADDITIONAL

FUSELAGE

PANEL MASS

LB/FT 2

8

6 PROP

PLANE

4

2

0

300 320 340 360 ' 380 400 420 440

FUSELAGE STATION - IN

FIGURE 51. Additional Acoustic Mass Treatment for Tractor Installation

EXTERIOR

OASPL

dB

P

7 / _-TRACTOR

-----FORWARD

FUSELAGE STATION

FIGURE 52. Uncertainty in Noise Directivity

71
The far field noise estimates are given in Table 9. Table 9 shows the

community noise at the three standard certification locations (as defined

in FAR 36, Appendix C: Noise Levels for Transport Category and Turbojet

PoweredAirplanes Under FAR36.201):

eTakeoff Flyover - beneath the flight path 21,325 FT from start of

takeoff roll ;

• Takeoff Sideline - peak noise level at a lateral distance of 1,476

FT;

• Landing Approach - beneath aircraft when at an altitude of 394 FT on

a three-degree glide slope, 6,562 FT from the touchdown point.

The propfan noise levels shownare for the pusher configuration with the 8

FT diameter prop and are estimated using Hamilton Standard's computer

method from Reference 5. Examination of these noise levels shows that the

pusher propfan meets all of the required FAR36 Stage 3 noise requirements.

This is achieved without the need for a power cutback after takeoff.

TABLE9. CommunityNoise Estimates

ITEM

FAR 36 REQUIREMENT

PUSHER PROPFAN

NOISE LEVEL - EPNdB

TAKEOFF APPROACH SIDELINE

89

85.5

98

94.6

94

89

72
5.7 PERFORMANCE

Table 10 presents a summary of the performance capabilities of each study

aircraft. Table 10 presents the design requirements established for the

study and then shows how each aircraft satisfied those requirements. All

of the aircraft were specifically sized to meet the design mission which

had a range requirement of 1800 NM plus 200 NM IFR alternate and a cruise

speed capability of .75 Mach at 41,000 FT altitude.

The field length goal selected was 4,000 FT. But because each of the

designs was constrained by other requirements, the field length for each

design was less than required. As discussed in Section 5.1, both turbofan

designs were constrained by the hot day takeoff climb (second segment climb

gradient). Both propfan configurations were not constrained by either

field length or climb gradient. Therefore, these were selected on the

basis of minimum fuel burn capability using the costing mission.

TABLE 10. Performance Summary

ITEM

RANGE +

ALTERNATE

@ 41,000 FT

CRUISE SPEED MACH

@ 41,000 FT

DESIGN

RQMTS

NM 1800+200

.75

CURRENT ADVANCED PUSHER TRACTOR

TURBOFAN TURBOFAN PROPFAN PROPFAN

1800+200

.75

1800+200

.75

1800+200

.75

TAKEOFF FIELD LENGTH* FT

S.L., STD DAY

SECOND SEGMENT*

CLIMB GRADIENT

@ 5,000 FT, ISA+25oc

%

4O00

2.4

3260

2.4

3520

2.4

2980

6.7

1800+200

.75

3140

6.0

*FAR 25

73
The time, rate of climb, maximum cruise speed, and VFR range performance

for each of the study aircraft are given in Figures 53 through 56. (Note:

VFR range is with only a 45 minute fuel reserve, i.e., no alternate

capability.) The effect of the flat rating on the propfan engines can be

seen in Figures 54, 55, and 56. The point at which both the maximum

continuous and maximum cruise power settings come off the flat rating

occurs roughly around 30,000 FT. There is an additional break, as with

most turbofan engines, at the iso-thermal break, roughly 36,000 FT. The

range shown in Figure 56 is a maximum VFR range with no alternate

capability and is basically a climb - cruise to destination - hold 45

minutes mission. The maximum cruise speed shown in Figure 55 is for a

mid-cruise weight.

The rate of climb curves in Figure 54 show that at low altitude the propfan

has less maximum rate of climb than the advanced turbofan. This was

due to the effect of the flat rating. The speed for the maximum rate was

approximately 220 KTAS for the turbofan and 170 KTAS for the propfan. When

the speed was reduced to the second segment climb speed for both aircraft

(roughly 130 KTAS) there was a significant relative increase in thrust for

the propfan because of the different thrust lapse rates with speed between

50000

RLTITUDE

(FT)

40000

30000

20000

PUSHER PR(

)R PROPFAN

TURBOFAN

CED TURBOFA#_

10000

0 I'0 ' 2.O 3.O ' ..O ' 5.O 6.O

TIME TO CLIMB (MIN)

FIGURE 53. Time to Climb

74
50000 1

RLTITUDE

(FT)

40000

30000

20000

10000

CURRENT

ADVANCED

1000 2000 :3000 4000

RATE OF CLIMB (FPM)

PROPFAN

PROPFAN

5000 6000

FIGURE 54. Rate of Climb

50000

.?M ,75M .SM

40000

ALTITUDE

(FT)

30000

20000

10000

TRACTOR PROPFAN

PusHER

rRRENT TURBOFAN

TURBOFAN

0

220 260 300 340 380

TRUE AIRSPEED (KT)

420 460 500

FIGURE 55. Cruise Speed

75
RLTITUDE

(FT)

50OOO 1

40OOO30000 ADVANCED TURBOFAN--_

CURRENT TURBOFAN N

20000

10000

0

0 500 1000 1500 2000 2500 3000

RRNGE (NM)

FIGURE 56. VFR Range

the turbofan and propfan. This difference in thrust, then, yielded the

higher second segment climb gradients shown in Table 10 for the propfans.

The mission performance is presented in Tables 11 and 12. Fuel required

for each segment is given. Figures 57 and 58 depict each numbered element

of the design and cost mission as they relate to Tables 11 and 12

respectively. The design mission, once again, was used to size each of the

aircraft. Subsequently, the required total mission fuel in Table 11 was

also the maximum usable fuel capacity for each of the study aircraft.

Table 12 shows both the fuel required for each segment and its relative

percentage to the total required mission fuel for the 400 NM cost mission.

It is interesting to note that the descent, landing, and taxi elements for

the propfans show a higher percentage of the total fuel burn than the

corresponding elements for the turbofans. This was caused by a higher fuel

flow at idle for the propfan engine data provided by Pratt and Whitney.

The reason for this was that the turboprop engines assumed for this study

required higher RPM's for minimum (idle) operation than did the turbofans

and thus used more fuel at idle. This effect of higher idle fuel flow was,

however, minimal on the total fuel consumption results.

76
0/

Y

ORIGIN

41,ooo_T® _.Ts.

\ 2s,oooFT®

1800 NM -I_ 200 NM

DESTINATION ALTERNATE

FIGURE 57. Design Mission Elements

TABLE 11. Design Mission Fuel Summary

MISSION ELEMENT

TO DESTINATION (1800 NM)

1. START, TAXI, TAKEOFF LB

CURRENT ADVANCED

TURBOFAN TURBOFAN

230 180

PUSHER

PROPFAN

169

2. CLIMB

3. CRUISE

4. DESCENT/APPROACH

LB

LB

LB

692

4321

276

493

3426

212

259

2953

231

TO ALTERNATE (200 NM)

5. MISSED APPROACH/CLIMB

6. CRUISE

7. 45 MINUTE HOLD

8. DESCENT/APPROACH

9. LANDING, TAXI, STOP

TOTAL MISSION FUEL

LB

LB

LB

LB

LB

LB

288

663

414

217

56

7157

210

534

340

169

44

5608

121

411

363

187

62

4756

TRACTOR

PROPFAN

169

272

2968

239

127

402

370

193

63

4803

77
33,000 FT Q .75M

2/ L_

L _1

I- 400 NM -I

ORIGIN DESTINATION

FIGURE 58. Cost Mission Elements

TABLE 12. Cost Mission Fuel Summary

MISSION ELEMENT

TO DESTINATION (400 NM)

I. START, TAXI, TAKEOFF

2. CLIMB

3. CRUISE

4. DESCENT/APPROACH

5. LANDING, TAXl, STOP

TOTAL MISSION FUEL

LB

(%)

LB

(%)

LB

(%)

LB

(%)

LB

(%)

LB

(%)

CURRENT ADVANCED

TURBOFAN TURBOFAN

230

(11.0)

497

(23.8)

1017

(48.7)

289

(13.8)

56

(2.7)

2089

(I00 )

180

(Ii.i)

349

(21.5)

834

(51.3)

218

(13.4)

44

(2.7)

1625

(100 )

PUSHER

PROPFAN

169

(12.1)

194

(13.9)

736

(52.9)

231

(16.6)

62

(4.5)

1392

(100 )

TRACTOR

PROPFAN

169

(12.0)

203

(14.5)

727

(51.9)

239

(17.1)

63

(4.5)

1401

(i00)

78
The cost mission fuel burn data for the pusher propfan is summarized in

Figure 59. The General Aviation propfan shows a 33% reduction in fuel burn

over the current turbofan for the cost mission and a 14% reduction in fuel

when compared to the equivalent 1990 technology turbofan. As mission range

increases, such as in the design mission, the relative fuel burn improves

slightly. The propfan design mission fuel burn is 16% less than the

equivalent technology turbofan and 34% less than the current day turbofan.

100

100

COST MISSION

BLOCK FUEL

%

50

78

0

CURRENT ADVANCED PUSHER

TURBOFAN TURBOFAN PROPFAN

FIGURE 59. Cost Mission Fuel Burn Comparison

79
5.8 COST

The total aircraft price for each airplane, as discussed in the Appendix,

was evaluated using the same formulation to provide an equal basis of

comparison. The pricing formulation does not yield the exact price of any

particular aircraft in the market today, but rather, gives an estimate

based on a correlation of published pricing information for various

aircraft which is typically within 5% to 7% of the published price data.

Pricing strategy or market fluctuation can induce similar variations in

price. Thus, the formulation yields a simplified yet relatively good

approximation of actual acquisition cost while also providing a consistent

method for price comparison in this study.

A similar method was utilized by Pratt and Whitney to estimate the engine

acquisition cost. The only disadvantage of this approach for the engines

is that it requires a considerable extrapolation of the turboprop cost data

from current market prices, since there are no General Aviation engines in

this horsepower class (2900 SHP). The propeller and speed control pricing

provided by Hamilton Standard was based on scaled data which was available

for the airline class propfans. Unlike the airline class studies, there

was no propfan design tailoring done for this preliminary study. Thus the

cost and performance of the propfans represent the large, transport propfan

data scaled to meet the General Aviation mission requirements. This is an

area where further study could help reduce the acquisition cost while still

maintaining the propfan performance levels.

The acquisition cost for both the propulsion system and total aircraft is

given in Table 13 for each of the study aircraft. As discussed in the

Appendix, the airplane price consisted of three elements; an airframe

weight related price, a propulsion system price, and a fixed avionics cost.

As shown in Table 13, even with a 10% higher propulsion price and a higher

cost for the use of advanced materials, the reduction in airframe weight

due to the advanced materials and improved engine SFC resulted in an

overall 1% reduction in total acquisition cost of the advanced turbofan

compared to the current turbofan. On the other hand, the pusher propfan

acquisition cost was 18% higher than the equivalent technology advanced

turbofan. The primary cause of this was that the pusher propfan propulsion

system cost was 35% higher than the advanced turbofan. Also a factor in

the higher cost was the additional weight of the structure required to

carry the propfan loads. The tractor installation, requiring even more

structural weight for acoustical treatment, had a 21% higher acquisition

cost compared to the advanced turbofan.

80
TABLE 13. Acquisition Costs

COST ITEM

TOTAL AIRCRAFT

PROPULSION SYSTEM

$(MILLION)

$(MILLION)

COST

CURRENT ADVANCED PUSHER TRACTOR

TURBOFAN TURBOFAN PROPFAN PROPFAN

3.60

.455

3.58 4.22

.464 .625

4.33

.621

A comparison of the propulsion system cost, given in Table 14, showed a

significant disadvantage for propfans. When compared on an equivalent

thrust basis, the propfan propulsion system was 33% higher cost than the

equivalent technology turbofan. Compared to the current technology

turbofan the propfan system was almost 45% higher cost. The propfan

propeller and speed control alone was 22% the cost of the current turbofan.

When the propulsion systems were then sized to meet the design mission

requirements, this cost disadvantage increased to 35%, shown in Table 13.

TABLE 14. Relative Propulsion System Acquisition Cost

ELEMENT

ENGINE %

PROPELLER %

TOTAL %

RELATIVE COST

CURRENT ADVANCED

TURBOFAN TURBOFAN

I00

i00 109

PROPFAN

122.7

22.2

144.9

NOTE: Propulsion systems matched for equivalent thrust

@ 41,000 FT, .75M.

81
The effect of the higher aircraft price on the total cost of operation for

these General Aviation propfans overcomes the cost improvement which

normally could be anticipated with a 15-30% improvement in fuel consumption•

The pusher propfan showed an 11% higher cost of operation, utilizing the

Corporate Cost model, compared to the advanced turbofan, Table 15, for

typical General Aviation annual utilization. On the other hand, the ATA

method showed a 10% cost disadvantage for the propfan. The modified ATA

method is less sensitive to aircraft acquisition cost than the Corporate

model• It is more representative of commuter airline aircraft and is

provided for comparison to other studies• A breakdown of the variable and

fixed elements of the annual operational costs are given in Tables 16 and

17. The fuel costs utilized in this study were based on a typical average

1984 retail price of $1.90 per gallon (see Appendix, Section A2.0 for

further discussion of fuel price). Table 16 shows that the cost of fuel

was reduced 15% by using a propfan while the overall maintenance cost only

increased by 11%. This would have provided a definite cost advantage for

the propfan if the propulsion system acquisition cost of the propfans were

TABLE 15. Total Cost of Operation for 400 Hours Annual Utilization

COST MODEL

MODIFIED ATA

CORPORATE

$/SEAT-NM

$/SEAT-NM

COST

CURRENT ADVANCED PUSHER TRACTOR

TURBOFAN TURBOFAN PROPFAN PROPFAN

.857

• 484

.797 .878

.453 .501

.899

.514

82
TABLE16. Variable Operating Costs - Corporated Model

ELEMENT

FUEL $/HR

@ $1.90/GAL

AIRCRAFT MAINTENANCE

LABOR $/HR

MATERIAL $/HR

PROPULSION SYSTEM

MAINTENACE

LABOR $/HR

MATERIAL $/HR

COST

CURRENT ADVANCED PUSHER TRACTOR

TURBOFAN TURBOFAN PROPFAN PROPFAN

392.84

61.98

60.39

22.37

102.64

309.28

57.20

48.12

22.36

99.80

264.04

58.36

52.59

33.59

107.35

264.51

60.08

59.62

33.60

106.92

TABLE 17. Fixed Operating Costs - Corporate Model

ELEMENT

CREW

INSURANCE

TOTAL

$(THOUSAND)

$(THOUSAND)

$(THOUSAND)

CURRENT

TURBOFAN

76.8

36.0

ADVANCED

TURBOFAN

PUSHER

PROPFAN

COST

76.8

35.8

76.8

42.2

TRACTOR

PROPFAN

112.8 112.6 119.0

76.8

43.3

120.0

83
similar to the turbofans. This was, however, not the case. Figure 60

summarizeseach configuration's cost element utilizing the Corporate cost

model and graphically illustrates the impact of the propfan°s higher price

on depreciation and interest costs.

It is important to note, also, in this study, that there was no cost added,

based on information from Hamilton Standard, for propeller system overhaul.

Hamilton Standard's current design philosophy for blade spars and hub is

for infinite life componentsand thus the maintenance costs for the propfan

blades and hub were for scheduled inspections, and on-condition repairs

only. This corresponded to only about 1% of the total propulsion system

cost shown in Table 16. There were anticipated to be no life limited parts

and thus replacement would be required only for accident or significant

foreign object damage. If this infinite design life can not be achieved,

however, an additional maintenance cost penalty for prop overhaul would be

required for propfans, thus making the propfan maintenance cost higher than

shown. The maintenance costs shownin Table 16 do include an allowance for

engine overhaul based on an average expected engine life of 3000 hours

before overhaul.

i00-

::::::::::::::::::::::::::::::::::::::::

6O

RELATIVE

COST OF -

OPERATION 40-

%

80

IllllllHIIIIIIIIIIlIllll

:::::::::::::::::::::::

:.-.......:.........- -..

::...'.:.-.-:..-:.-.: -.:

:.-:-..-.-.:.-.-.:....:

:..- .::':-.'.'..'.:.: :.."

20 :::::::::::::::::::::::

:::::::::::::::::::::::

O _ ,-oO.-.,°-o-o-°-o-°-.°.

CURRENT

TURBOFAN

*CORPORATECOST MODEL 400 HOURS ANNUAL UTILIZATION

,%_ii_!_i.:...i._.z!!_i_:.i_i_!!!ii_ ! ........................................

_iiiiiiii i!_i _i_!i_iillli:_.-'._:,_-i_i_i-:_d _ii_;iii_i_ @:_!i!_ F UEL

.............. IL CRE 

......_ _ mm_mmm_'_,, AIRFRAME MAINT / IIIHHIIIIHHIIHIHIIII __----

___ I_ ENGINE MAINT

_ INSURANCE

i i_i!iiiiii!iii!:.-'.i_il

°o°°-°°°.°°.°°-°°.°°°,

,°°,°-°°°°,°,°.°°°°Oo°,

O°,o°°°°°.°,°°°.°.°.°,

o°,°°°.°,°.°°,°o°.O°°,

°°°°o°°°°..°°°-°°.o°°,

°°°.,°°°..°°,°-o-°-°-,

°oO,°.°°°°°.°°°o°°°°-,

°°-o°°-°°°°°°°-.°°°°°,

,°°°o°°°°-°°°o°,.°°°°o,

,°°,°°°°°°,°o°°°°°o°°,,

,........:....._-.-.-.-,

,.......:.-...-.-.-...,,

°.°,°,°°°°°°°°°°,°-°°,

,-°o,°°,°°°°.,°°°°°°°o,

,-°o,-°°°°°OoO°°°,°°°°,

,°°°°-o°°°,°o°°,°o,o°o,

,°°°°°°°°°°°o°,O°O°°°°,

,°°.°°.°o.°°°-°°°°°°,°,

• .........

ADVANCED

TURBOFAN

:::::::::::::::::::::: -.-...........-.-.-.-.

°°.,°°°°o°,°°°.°°°°o°°

:::::::::::::::::::::: ......-.....-.-.-...-.

°°°o°°°°°.°°,°°°°°°°-°

:::::::::::::::::::::: ......-.......-.-.....

°°°°o°°,°°°°°°°°°°°,°°

::::::::::::::::::::::iiiiiii!iii!i_iiiiiiii

:::::::::::::::::::::: :.:.:.1.:.:.1.:.:.:.:.

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...................-.- ::::::::::::::::::::::

........... ° ..........

°°°°°°°°°°°°°°-°°°°°°° °°°°°°°o.°°o,°°°.o°.°

°°°°,,,°°°°°°°°°°°°°°° ..........

........... ° ..........

°°.°,°°.,..°o°°°°°°°°. °°°°°°,°°°°°°°°o°,°°°

...-.........-........1:.:.:.:.:.:.:.:.:°

•:-:-:.:.:-:-:.:.:.:.: :::::::::::::::::::::

°°o°,.°°°°°°°°°°Oo°°°° °°°°.°°°°,°°°°°o°°.°,

o.°°°.°°°o°°°°,°°.-°°° °.°°°°°°°°°°°-o°°°°o,

PUSHER TRACTOR

PROPFAN PROPFAN

INTEREST

DEPRECIATION

FIGURE 60. Relative Cost Summary for the Study Configurations

84
The depreciation effects and cost of capital significantly affect the total

cost of operation values. A complete breakdown of the costs and

depreciation effects are given in the Appendix. For comparison, the

modified ATAcost model results for 400 hours annual utilization are given

in Table 18, and the results for 3000 hours annual utilization are given in

Table 19. The variable elements, given in Table 16, in the Corporate model

are the same as those used in the ATA model. Additionally, the fixed

elements of the Corporate model and the depreciation effects were evaluated

on an hourly basis in the ATA model. The ATAmodel, unlike the Corporate

model, does not provide a cost of capital element (interest cost). At 400

hours utilization the propfan's reduction in the cost of fuel whencompared

to the advanced turbofan was still 15% and the increase in overall

maintenance was 11%, similar to the Corporate model results. The hourly

cost of depreciation and insurance, however, can be seen directly using the

ATA model. In Table 18, the pusher propfan shows a 19% increase for

depreciation and insurance costs over an equivalent advanced turbofan.

Figure 60 shows a similar effect for depreciation and insurance using the

Corporate model. At 3000 hours utilization, as shown in Table 19, the

propfan's relative reduction in fuel cost and relative increase in

maintenance, depreciation, and insurance costs from the advanced turbofan

were the same as at 400 hours utilization. However, the depreciation and

insurance costs' percentage of the total cost of operation was

significantly less than at 400 hours. Because of this, the cost of fuel

exerts more influence on the total cost of operation at 3000 hours, whereas

at 400 hours the aircraft price related items such as depreciation and

insurance predominate the cost of operation.

85
TABLE 18. Modified ATA Costs for 400 Hours Utilization

COST ELEMENT

CREW $/HR

FUEL $/HR

@ $1.90/GAL

INSURANCE $/HR

DEPRECIATION $/HR

AIRCRAFT MAINTENANCE

LABOR $/HR

MATERIAL $/HR

PROPULSION SYSTEM

MAINTENANCE

LABOR $/HR

MATERIAL $/HR

,, .

TOTAL $/HR

COST

CURRENT

TURBOFAN

ADVANCED

TURBOFAN

PUSHER

PROPFAN

TRACTOR

PROPFAN

192.00

392.84

89.91

895.40

192.00

309.28

89.47

892.72

192.00

264.04

105.60

1065.29

192.00

264.51

108.36

1090.00

61.98

60.39

22.37

57.20

48.12

22.36

58.36

52.59

33.59

60.08

59.62

102.64

1817.53

99.80

1710.95

107.35

1878.82

33.60

106.92

1915.09

86
TABLE 19. Modified ATA Costs for 3000 Hours Utilization

COST ELEMENT

CREW $/HR

FUEL $/HR

@ $1.90/GAL

INSURANCE $/HR

DEPRECIATION $/HR

COST

CURRENT

TURBOFAN

ADVANCED

TURBOFAN

PUSHER

PROPFAN

TRACTOR

PROPFAN

192.00

392.84

11.99

119.39

192.00

309.28

11.93

119.02

192.00

264.04

14.07

142.05

192.00

264.51

14.45

145.33

AIRCRAFT MAINTENANCE

LABOR $/HR

MATERIAL $/HR

PROPULSION SYSTEM

MAINTENANCE

LABOR $/HR

MATERIAL $/HR

TOTAL $/HR

61.98

60.39

22.37

102.64

963.60

57.20

48.12

22.36

99.80

859.71

58.36

52.59

33.59

107.35

864.05

60.08

59.62

33.60

106.92

876.52

87
6.0 DESIGNSENSITIVITIES

Several design sensitivities were examined in this study to enhance the

comparison of propfan propulsion to turbofan propulsion. These

sensitivities can be grouped into two categories; configuration sensitivity

factors and cost sensitivity factors.

6.1 CONFIGURATIONFACTORS

Several important design sensitivities were examined which impact the

overall definition of the configuration. As discussed in Section 3.2.2,

variations in aircraft and propulsion system parameters such as wing aspect

ratio, wing sweep, wing taper ratio, wing thickness, propeller diameter, or

number of propeller blades can significantly affect the selection of an

optimum design depending on the criteria for optimization being used. The

configurations in this study were selected for minimumfuel where possible.

Minimumfuel burn was selected as a design criteria rather than minimum

cost of operation for two reasons. First, the study utilized two cost

methods both of which were valid for specific types of utilization but

yielded different levels of answers. Second, because of the high cost of

the propfan propulsion systems comparedto turbofans there was concern that

using cost of operation as the design constraint might yield unrealistic

design choices considering the preliminary nature of these cost estimates

for the conceptual propulsion system designs in this study.

The aircraft geometry was defined as discussed in Section 3.2.2. The

parametric effect of wing taper ratio and thickness ratio were minimal on

mission fuel burn. These parameters were thus selected to provide a wing

design which would provide sufficient volume for the required systems and

fuel for the design mission, while preventing any possible shock formation

which might interfer with the flow field entering the propfan propeller.

To further assure a shock free wing at the design speed the wing was swept

250. The parametric effect of aspect ratio on fuel burn was examined and

is shown in Figure 61. Figure 61 showsthat the minimummission fuel burn

for each design occurs above the maximumwing fuel volume capability. Each

aircraft was resized for each new aspect ratio according to the method

discussed in Section 3.2.2. The ground rules required all the fuel to be

located in the wing to allow a commonfuselage design (i.e. no additional

fuselage fuel tanks) to be utilized in the study. An aspect ratio of 8.5

was chosen for all of the designs. This provided wing designs which in

each case had sufficient fuel capacity to meet the design mission without
2200

2000

COST1800

MISSION

BLOCK

FUEL 1600-

LB

1400-

1200

FIGURE61.

CURRENT

TURBOFAN

ADVANCED

TURBOFAN

I SELECTED AR

PUSHER

- PROPFAN

VOLUME LIMIT FUEL

1'o 1'1 1'2 1'3

WING ASPECT RATIO

Mission Sensitivity to Design Aspect Ratio

requiring fuselage tanks. Since the relative fuel change was essentially

the same for any given aspect ratio, including the minimum values,

choosing 8.5 as the wing aspect ratio did not impact the relative

propulsion system comparison.

In the preliminary phase of the study it was apparent that weight and, more

importantly, cost were going to be major elements in the assessment of

propfan propulsion. To achieve the most favorable comparison (lowest

cost/weight), a six-bladed, single rotation propfan was chosen. The

propeller diameter chosen, however, was selected using the minimum mission

fuel criteria. It was found that variation in the design mission could

result in a significant change in the propfan diameter selection. As shown

in Figure 62, by changing just the design mission (1800 NM, .75 Mach,

41,000 FT) to a lower altitude and higher speed (1800 NM, .80 Mach, 35,000

FT), while leaving the costing mission used for the fuel analysis

unchanged, a significant increase in design disk loading for the propfan

and reduction in propeller diameter would result. This higher disk loading

was similar to disk loadings seen in other studies and emphasized the

sensitivity of the propfan design to the selected mission requirements.

a

90
1580-

1540

1500

1460

COST

MISSION

BLOCK FUEL

LB 1420

1380

1340

10

35.000 FT

SHP/D2FORMAXIMUMCONTINUOUS

FIGURE 62. Design Mission Impact on Propfan Selection

The propfan noise estimates for the study were made utilizing the Hamilton

Standard method defined in Reference 5. This method was developed based on

data for tractor propfans. A pusher propfan typically operates in a

significantly more disturbed flow field than a tractor propfan. As

discussed in Section 5.6, recent work has been done at NASA (Reference 14)

that indicates that pushers in the presence of a pylon or wing wake could

have a substantially different noise directivity distribution more forward

of the prop plane than that of a tractor installation. To assess the

impact of the possible higher noise levels for pushers the sensitivity of

the design to increased acoustic treatment weight was examined. The

results are shown in Figure 63.

91
CHANGE IN

BLOCK FUEL

%

!

-2

3

| ! !

-2

i 2

!

3

CHANGE IN ADDITIONAL

ACOUSTIC TREATMENT WEIGHT

% TOGW

NOTE: FOR RESIZED PUSHER

PROPFAN AIRCRAFT

FIGURE 63. Mission Fuel Sensitivity to Acoustic Treatment Weight

6.2 COST FACTORS

The impact on the cost of operation due to variations in propulsion system

cost, aircraft acquisition cost, annual aircraft utilization, and price of

fuel were examined in the study. Both cost models used in this study

consisted of similar operational cost elements. The ATA model's fuel cost,

insurance cost, crew cost, airframe and propulsion system maintenance cost

were all evaluated on a flight hour basis. The Corporate model utilized

these same operational costs. It merely separated them into a set of

annually fixed costs and a set of flight hour variable costs, where the

crew and insurance were the fixed cost and the fuel and maintenance were

the variable cost (see Appendix). The financial aspects, on the other

hand, were treated differently in each model. The ATA model used a

formulation for depreciation which included a 5-year residual value of the

aircraft and an allowance for engine spares. The Corporate model, in

addition to depreciation figured on current 1984 tax laws, also included

the interest cost of borrowing capital to finance the aircraft. These

differences between the models resulted in some interesting variation in

sensitivity.

92
Both models were significantly influenced by utilization hours. With the

utilization at the 400 hour level per year typical of business aircraft,

the financial aspect of each model was the primary driver of the cost of

operation. However, as the utilization increased up to commuter airline

levels, such as 3000 hours, fuel consumption became the major influencing

element in the operating cost. For example in Figure 64, using the

corporate cost model for 400 hours utilization the depreciation and

interest cost, which was directly influenced by the price of the aircraft,

was 75.4% of the total cost of operation, whereas the fuel use represents

only 9.3%. In the ATA model the influence of price was reduced because the

model contains no cost of capital factor, yet, even still, the depreciation

element represented 51.6% of the direct operating cost while fuel cost

represented only 18.4%. On the other hand, using the ATA model at 3000

hours utilization, the depreciation only represented 13.5% of the cost of

operation, whereas the fuel cost element increased to 36.1% of the cost.

Similarly, fuel cost at 3000 hours for the Corporate model was 33.2% of the

total cost.

i00-

80

60

RELATIVE

OPERATING

COST 40

%

20

0

MODIFIED ATA MODEL

3000 400

HOURS

i

ilittiliilllllll

?W

ilittiiiillllll

DEPRECIATION

ANNUAL UTILIZATION 400 3000

CORPORATE MODEL

FIGURE 64. Cost Sensitivity to Chosen Cost Model and Annual Utilization

93
The Corporate model, with its effects of tax and cost of capital, exhibits

a stronger sensitivity to aircraft purchase price than does the ATA model.

This is shown in Figure 65. Aircraft purchase price affects the basis of

depreciation, the amount financed and interest paid, the cost of insurance,

and the tax credit received. Since the typical corporate utilization is

400 hours per year and these price related items are over 75% of the total

cost of operation, the aircraft purchase price becomes the major factor in

the cost analysis of a corporate aircraft.

CHANGE IN

COST OF

OPERATION

%

15

10

-10 .f

.Y

-5°

-10

COSY METHOD

----- MODIFIED ATA

_CORPORATE

j....-

0 5 10 1'5

CHANGE IN AIRCRAFT ACQUISITION COST

%

FIGURE 65. Operating Cost Sensitivity to Initial Aircraft Selling Price

94
The propulsion system acquisition cost has a major impact on the initial

selling price of the aircraft. The propfan propulsion system alone makes

up over 25% of the aircraft's empty weight (see Table 8 of Section 5.4).

From a simplistic view this weight corresponds to the amount of material

used which corresponds to the cost. The higher the percentage that the

propulsion system's weight is of the total aircraft weight the greater the

influence of the propulsion system price on the total aircraft price. As

smaller aircraft designs are examined a definite scale effect becomes

apparent. For example, a DC-9 airliner's propulsion system weight is

approximately 10% of the total aircraft's empty weight, a Diamond I

business jet's propulsion system is 15% of the total empty weight, and a

King Air B2OO's propulsion system is roughly 22% of the aircraft empty

weight.

To better understand the sensitivity to these cost factors and compare

these results to other studies that have been done, a comparison was done

between the factors affecting the cost in this study and a study done by

the Pratt and Whitney Aircraft Group (Reference 15) on an airline class

aircraft. The impact of the scale effect can be seen in Table 20. The

higher weight of the propfan installation as compared to an equivalent

TABLE20. Relative Weight of Propfans Comparedto Turbofans

ITEM

PROPULSIONSYSTEMWEIGHTINCREASE

(% OFEMPTYWEIGHT)

AIRFRAMESTRUCTURALWEIGHTINCREASE

(% OFEMPTYWEIGHT)

ADDITIONALACOUSTICTREATMENTWEIGHT

(%OFTAKEOFFWEIGHT)

%

%

GENERAL AIRLINE

AVIATION CLASS*

+26.8

+4.1

+1.5

+7.6

+4.3

+1.9

*Study by Pratt and Whitney Aircraft Group

95
technology turbofan installation represented a larger percentage of the

total aircraft weight than an equivalent comparison of airline class

propulsion systems. The airframe and acoustic treatment were, however,

similar for both the airline class design and the General Aviation design.

With the higher sensitivity to propulsion system weight, the small business

class aircraft exhibited more sensitivity to propulsion system cost than

did the airline class aircraft. Combining this higher sensitivity to price

with the higher cost of propfan propulsion resulted in a significant

disadvantage for propfan propulsion, in spite of a potential 15%reduction

in fuel consumption. Table 21 shows that a General Aviation propfan

propulsion system had a 34.6% higher acquisition cost than a turbofan

engine with equivalent mission capabilities. This combined with the higher

airframe cost for propfans (due primarily to the higher weight) resulted in

an overall 18%higher total aircraft selling price for a General Aviation

propfan whereas an airline class propfan was only 2.2% higher than an

equivalent turbofan.

TABLE21. Relative Cost Factors of Propfans Comparedto Turbofans

ITEM

PROPULSION SYSTEM COST INCREASE %

AIRFRAME AND AVIONICS COST INCREASE %

TOTAL AIRCRAFT PRICE INCREASE %

GENERAL AIRLINE

AVIATION CLASS*

+34.6 +3.8

+ 4.2 +1.8

+18.0 +2.2

*Study by Pratt and Whitney Aircraft Group

96
The difference in operational characteristics from airline operation, as

shown in Table 22, also affected the comparison of General Aviation

propfans to turbofans. First, the fuel consumption reduction for this

TABLE 22. Relative Operational Factors of Propfans Compared to Turbofans

ITEM

ANNUAL AIRCRAFT UTILIZATION

CRUISE FUEL CONSUMPTION REDUCTION

GENERAL

AVIATION

HR 400

% 15

AIRLINE

CLASS*

3000-4000

20

*Study by Pratt and Whitney Aircraft Group

study was less aggressive than that shown for the airline study. However,

more reduction than this was not believed to be realistically achievable

for a General Aviation engine by 1988. Second, the cost factor differences

were amplified over the fuel benefits for General Aviation because of the

typically lower aircraft utilization rates as compared to airlines. A

comparison of the effect of these cost factors in Table 21 on aircraft cost

of operation is shown in Figures 66 and 67. Figure 66 shows that, with the

cost factors determined for this study (including the benefits of the lower

fuel consumption shown in Figure 59 in Section 5.7), a General Aviation

propfan would have a 4% disadvantage compared to a current day turbofan

capable of the same mission; an 11% disadvantage compared to an equivalent

technology turbofan. If it were possible to reduce the cost differential

between General Aviation propfans and turbofans such as those shown for the

airline class aircraft in Table 21 a 2% advantage could be shown for

propfans over turbofans with 400 hours utilization. As the utilization was

increased to 3000 hours, Figure 67, this improvement would increase to 3%.

97
COST OF

OPERATION

%

100'

5O

100

_#iii_i_iiiii_i_iiii_iiiii!iiiiiiii

CURRENT

TURBOFAN

94

_:.:_->>::: :::::::::::::::::::::

:)!:i:i:!:i:i:i:i:_:i:_:i;_:i:i:i:_:_:i:_:_:i:_:_:!:i:_

_il}ii!!ii!iii!i!ii!!iiiii!iiiiiliiiiiiiiiiiiiiiiiiiii

_._:_:_:!:_ii:!i!:i_i_iiiii_iiii!ii:ii!iii!iii!iiiii_:

::::::::::::::::::::::::::::::::::::::::::::::::::::::::

::::::::::::::::::::::::::::::::::::::::::::::::::::::::

104

ADVANCED. PUSHER

TURBOFAN PROPFAN

92

PROPFAN WITH

AIRLINE CLASS

COST FACTORS

FIGURE 66. Relative Operating Cost Comparison for 400 Hours Annual

Aircraft Utilization (Corporate Model)

COST OF

OPERATION

%

100

5O

I00

88 89

CURRENT ADVANCED PUSHER

TURBOFAN TURBOFAN PROPFAN

85

PROPFAN WITH

AIRLINE CLASS

COST FACTORS

FIGURE 67. Relative Operating Cost Comparison for 3000 Hours Annual

Aircraft Utilization (ATA Model)

98
The sensitivity of the cost of operation to the price of fuel is given in

Figures 68 and 69. Because of higher fuel consumption the turbofan

propulsion system is slightly more sensitive to change in fuel price than

the propfan propulsion system as seen in Figures 68 and 69. Also, as

annual utilization was increased in Figure 69, fuel use increased and both

propfan and turbofan exhibited more sensitivity to change in fuel price.

The base fuel price assumed for both these Figures was $1.90/gallon (see

Appendix, Section A2.0 for further discussion of fuel price).

CHANGE 4

IN

OPERATING

COST 2

%

w I I

-6

ADVANCED

TURBOF_

1.6%_

___I.3%

10 20 30

G/A PROPFAN

CHANGE IN PRICE OF FUEL

%

CORPORATE COST MODEL

400 HOURS UTILIZATION

FIGURE 68. Sensitivity of Operating Cost to Fuel Price at 400 Hour

Annual Aircraft Utilization

99
CHANGE

IN 8-

OPERATING

COST

4

%

ADVANCED

TURBOFAN

3

3.0%

-io 1'o

-12-

G/A PROPFAN

2'0 3b

CHANGE IN FUEL PRICE

%

MODIFIED ATA MODEL

3000 HOURS UTILIZATION

FIGURE 69. Sensitivity of Operating Cost to Fuel Price at 3000 Hour

Annual Aircraft Utilization

6.3 CONFIGURATION/COST SENSITIVITY TO RETROFIT

A common practice in General Aviation (as well as other industries) is the

technique of retrofitting a new system on a current design. This technique

frequently results in improved design versatility at substantially lower

cost. To examine the sensitivity of the designs in this study to a

retrofit approach the current turbofan design was retrofit with both the

advanced 1988 turbofan engine and the 1988 pusher propfan installation.

The results are shown in Table 23. Both 1988 engines were sized to provide

the thrust necessary to give a .75 Mach cruise at 41,000 FT. In addition

to the engine retrofit, the propfan also required the changes in tail size,

nacelle attachment structure, systems weight, and additional acoustic

treatment to accommodate the propfan propulsion system. These changes were

reflected in the propfan's increased empty weight. It was assumed that the

retrofit configuration could be balanced without major problems which could

resize the aircraft and diminish the potential benefits. Further, it was

assumed that the takeoff weight would remain unchanged, and, therefore, the

propfan had less usable fuel available. Nevertheless, the propfan still

increased the range significantly from the current turbofan's range. The

i00
cost mission showed a 28.9% reduction in fuel use relative to the current

day turbofan; a 15.4% reduction relative to the advanced turbofan. Once

again, however, because of the much higher propulsion system acquisition

cost and subsequent higher aircraft selling price the retrofit pusher

propfan showed a 9.3% cost of operation disadvantage to both turbofans.

TABLE 23. Current Turbofan with Retrofit 1988 Propulsion Systems

ITEM

WEIGHTS

MAX TAKEOFF WEIGHT LB

CURRENT

TURBOFAN

17033

RETROFIT

ADVANCED

TURBOFAN

USABLE FUEL LB

EMPTY WEIGHT LB

PERFORMANCE

VFR RANGE*

CRUISE SPEED @ 41,000 FT

COST MISSION BLOCK FUEL

*WITH 45 MIN RESERVE

NM

MACH

LB

COST

AIRCRAFT ACQUISITION COST $ MILLIONS

PROPULSION ACQUISITION COST $ MILLIONS

COST OF OPERATION $/SEAT-NM

7157

8676

2175

.75

2089

3.60

.465

.484

17033

7157

8676

2703

.75

1757

3.75

• 499

.484

PUSHER

PROPFAN

17033

6209

9624

2866

.75

1486

4.38

.651

.529

i01
7.0 CONCLUSIONS AND RECOMMENDATIONS

The results of this study indicate that propfan propulsion is technically

feasible for application to General Aviation. Propfan propulsion applied

to General Aviation aircraft provides comparable mission capabilities to

General Aviation turbofans. For the typical 400 hours annual utilization

that business aircraft experience, propfans can provide the equivalent

speed and payload/range capabilities of a business jet while also providing

substantial fuel burn reductions. The fuel consumption benefits by

utilizing propfan propulsion are:

• a 33% reduction in fuel consumption when compared to a 1984

(current) technology turbofan aircraft;

• a 14% reduction in fuel consumption when compared to a 1990

(equivalent) technology turbofan aircraft;

• a 29% reduction in fuel consumption (or 32% increase in

range) using a current aircraft retrofit with propfans.

The aircraft examined in this study show that, based on current

understanding, propfan propulsion can provide acceptable near field and far

field noise levels similar to or better than turbofans. For far field, the

propfan aircraft examined in this study achieved community noise levels

below the Stage Ill requirements of the Federal Aviation Regulation's Part

36, Appendix C. For near field, an average cabin interior noise level of

82 dBA was selected as the design criteria. To meet this goal the General

Aviation propfans examined in this study required an additional acoustic

treatment above the noise treatment used in the turbofans of 1.5% of the

takeoff gross weight (TOGW) for the pusher installation and 4% of TOGW for

the tractor installation.

In spite of the obvious marketable benefits of reduction in fuel

consumption with comparable mission capabilities to aircraft with

turbofans, propfans on General Aviation aircraft have a severe cost of

operation (DOC) disadvantage compared to turbofans. The General Aviation

cost of operation exhibits higher sensitivity to aircraft initial selling

price and the tax environment than does the cost of operation (DOC) for

large transport class aircraft, while showing less sensitivity due to

changes in fuel consumption. This is primarily due to the lower annual

utilization rates for business aircraft.

PRECEDING PAGE BLANK NOT FILMED

103
The lower aircraft annual utilization rates for General Aviation (400

hours/year), as comparedto airlines (3000-3400 hours/year), and the 18%

higher aircraft acquisition cost for the propfan combine to give a cost of

operation (or direct operating cost, DOC) that is much higher than a

turbofan. The results showedthat the cost of operation for a propfan was:

• 4% higher than a 1984 (current) technology turbofan;

• 11% higher than a 1990 (equivalent) technology turbofan.

The higher aircraft price is due primarily to the significantly higher

acquisition cost for the propfan propulsion system:

• 37.4% higher price than a 1984 (current) technology turbofan

engine;

• 34.7% higher price than a 1988 (equivalent) technology turbofan

engine.

Propfan propulsion poses serious technical and cost challenges before it

can be successfully implemented in General Aviation. The broad brush

approach taken in this study indicated that several installation concepts

utilizing propfans were technically feasible. However, further development

in key technical and cost areas is required. Because of the smaller scale

of General Aviation aircraft, as compared to transport aircraft, and the

relatively large turboprop engines required for propfan propulsion to match

turbofan mission capability, pusher configurations provide the best

apparent installation from overall weight, balance, noise, and safelife

design considerations. Analytical studies are needed, along with

verification by test, to resolve the technical uncertainties for a pusher

installation. In particular, the following items need further work:

• the acoustic levels for pusher propfans and their

distributions/directivity for typical flight disk loadings;

• the asymmetric wake effects of a body or lifting surface in

front of the prop plane on the blade structural dynamics and

performance;

• the flow field shock effects on the blade dynamics;

• the exhaust flow effects on the propfan blades;

104
• the design and certification aspects of propfan structural

fai Iure ;

mthe development of design codes for the General Aviation

class propfans.

Of higher priority than the technical aspects, before propfan propulsion

can become marketable for General Aviation, the propulsion system cost for

propfans must be competitive with turbofans. Once again, because of the

broad brush aspect of this study, the costs provided by the propulsion

manufacturers were based on sizable extrapolations. More work needs to be

initiated by NASA to provide better definition of the propulsion cost

elements involved and to study potential engine and propfan cost

reductions.

105
APPENDIX

COST ESTIMATING MODEL

Two cost models were developed for this study to compare the propulsion

system designs. The first method consists of the 1967 Air Transport

Association (ATA) (Reference 16) cost model modified to reflect General

Aviation. The second method takes a similar approach to operating cost, as

does the ATA model, but also includes the cost of capital and tax benefits

of ownership. In each case, 1984 was assumed for the dollar value and tax

environment.

AI.0 MODIFIED 1967 AIR TRANSPORT ASSOCIATION (ATA) MODEL

The Direct Operating Cost (DOC) model is based on definitions provided in

the 1967 ATA formulation and the approach used in Reference 17. The crew

costs, fuel costs, insurance, and labor rates are based on 1984 levels

typical for General Aviation. Maintenance costs contain both a cyclic and

a flight hour factor. Propulsion system maintenance costs were supplied

both by Pratt and Whitney, Canada and Hamilton Standard. Maintenance labor

cost factors were assumed to be the same for both the current day

propulsion system and the 1988 technology level propulsion systems.

Variation in maintenance costs between current day and 1988-1990 level

technology shown in the study occur from a difference in airframe material

cost and propulsion system price.

A1.1 MODIFIED ATA COST ESTIMATING FORMULAS

The DOC model formulation used in this study is as follows:

DIRECT OPERATING COST

DOC = Cc + Cf + Cd + Ci + Cal + Cam + Cpl + Cpm (S/flight)

CREW COSTS -

Cc = Nc (Kc) (tb)

teR_CE_ING PAC_ Bt.ANK NOT

107
where

Cc = crew cost (S/flight)

Kc = crew hourly rate ($/HR) = $96/HR

Nc = number of crew = 2

tb = block time (HR)

FUEL COST -

Cf = Fb (Pf) / 6.7

where

Cf = cost of fuel (S/flight)

Fb = block fuel (LB)

Pf = price of jet fuel ($/GAL) = $1.90 (Reference 18)

HULL INSURANCE COST -

C i = tb (Ct) IR / U

where

C i = cost of hull insurance (S/flight)

tb = block time (HR)

Ct = aircraft initial selling price ($)

IR = annual insurance rate (%) = 1%

U = annual utilization (HR) = 400 HR or 3000 HR

DEPREC IATION -

Cd = tb (.46C t + .15 Cp) / (Da U)

where

Cd = cost of depreciation (S/flight)

Da = depreciation period (YR) = 5 YR

Ct = aircraft initial selling price ($)

Cp = propulsion system price ($)

U = annual utilization (HR) = 400 HR or 3000 HR

tb = block time (HR)

108
ATDDDAM_ .......,_ M,.AIN,_NANvE ITW P LABOR _POS T, _

Cal = (KI (tf)+ K2) (LR) (B)

where

Cal = airframe maintenance labor cost (S/flight)

K 1 = labor manhours per flight hour (MHR/HR)

K 1 = .59 K2

K 2 = labor manhours per flight cycle (MHR)

K2 = (.05 (Wa) / 1000) + 6 - 630 / ( (Wa / 1000) + 120)

tf = flight time (HR) = tb - tg

tb = block time (HR)

tg = ground time (HR) = .16 HR

LR = labor rate ($/MHR) = $21.67/HR

B = overhead burden factor = 1.8

Wa = aircraft airframe weight (LB)

AIRFRAME MAINTENANCE MATERIAL COST -

Cam = (K3 (tf) + K4) Wa/lO00

where

Cam = airframe maintenance material cost (S/flight)

K3 = airframe material hourly factor ($/HR/LB)

K3 = 3.08 (Ca + Cae)/106

K4 = airframe material cyclic factor ($/LB)

K4 = 6.24 (Ca + Cae)/lO 6

Ca = complete airframe price ($)

Cae = price of typical optional avionics and equipment ($)

Wa = aircraft airframe weight (LB)

tf = flight time = tb - tg

tb = block time (HR)

tg = ground time (HR) = .16 HR

PROPULSION SYSTEM MAINTENANCE LABOR COST -

Cpl = Ne (K5 (tf) + K6) (LR)(B)

where

109
Cpl = engine or engine/propeller maintenance labor cost

(S/flight)

Ne = numberof engines

K5 = labor manhours per flight hour per engine (MHR/HR)

K6 = labor manhours per flight cycle per engine (MHR)

LR = labor rate ($/MHR) = $21.67/HR

B = overhead burden factor = 1.8

for turbofans

K5 = .30

K6 = .03

for propfans

K 5 = .45 + (1.57 x IO-4(NB)(Dia) + .0045)

K 6 = .03

where

NB = number of blades

Dia = propeller diameter (FT)

PROPULSION SYSTEM MAINTENANCE MATERIAL COST -

Cpm = Ne (K7 (tf) + K8)

where

Cpm = engine or engine/propeller maintenance material cost

(S/flight)

Ne = number of engines

K7 = propulsion system material cost per flight hour ($/HR)

K8 = propulsion system material cost per flight cycle (S/flight)

for turbofans

K7 = .75 (FN) .5

K 8 = .35 (FN) .5

F N = thermodynamic thrust at SLS (LB)

110
for propfans

K7 = (Pt)-5 + (6.11 x 10-3 (NB)(Dia) + .175)

K8 = .20 (Pt).5

Pt = thermodynamic horsepower at SLS (SHP)

N B = number of blades

Dia = propeller diameter (FT)

A1.2 DOC BREAKDOWNS

Tables 24 through 27 show the DOC element costs for the cost mission with

an aircraft annual utilization of 400 hours. Tables 28 through 31 show

these same elements only with an aircraft annual utilization of 3000 hours.

111
TABLE 24. Current Turbofan ATA Analysis at 400 Hours Utilization

CURR_T ..... (1984) TURBOFAN

MISSION STAGE LENGTH = 400.0 NM YEARLY UTILIZATION = 400 HOURS

NUMBER OF PASSENGER SEATS = 8 AIRCRAFT EMPTY WEIGHT = 8676.0 LB

_I TURBOFAN ENGINE _

_NGIN_ SL T_EnMODYs_JIC THRUST 3257.0 LB

BLOCK TIME = 1.5(_-)8 HR

BLOCK FUEL = 2089.0 LB

MISSION ELEMENT COST BREAKDOWN:

CREW COST = $ 289.54

FUEL COST = $ 592.40

INSURANCE COST = $ 13_5 .uo_

DEPRECIATION COST = $ 1350.26

AIRFRAME MAI._TEN_NC_ MATERIAL COST = $ 91.07

AIRFRAME MAINTENANCE LABOR COST = $ 93.47

PROPULSION MAINTENANCE MATERIAL COST = $ 154.77

PROPULSION MAINTENANCE LABOR COST = $ oo._ 7_._

TOTAL COST = $ 2740.83

SUMMARY:

AIRCRAFT FIRST PRICE = $ 3596221

PROPULSION SYSTEM PRICE/ENG = $

DIRECT OPERATING COST PER HOUR =

COST PER SEAT MILE =

455134

1817.53 $/HR

.857 $/seat-nm

112
TABLE 25. Advanced Turbofan ATA Analysis at 400 Hours Utilization

ADVANCED (1990) TURBOFAN

MISSION STAGE LENGTH = 400.0 NM YEARLY UTILIZATION = 400 HOURS

NUMBER OF PASSENGER SEATS = 8 AIRCRAFT EMPTY WEIGHT = 7426.0 LB

_ TURBOFAN ENGINE _

ENGINE SL THERMODYNAMIC THRUST 3067.0 LB

BLOCK TIME =

BLOCK FUEL =

1.490 HR

1625.0 LB

MISSTOj_ ELEMENT COST B_=AKDOWN:

CREW COST = $ 286.08

FUEL COST = $ 460.82

,_=ur,_- COST = $ 133.31

DEPRECIATION COST = $ 1330.17

AIRFRAME :o_.i:_,_u_ _^_ _'_P_ ...... MATERIAL COST = $ 71.70

AIRFRAME MAINTENANCE LABOR COST = $ 85._3

Fr_uFU._oION MAINTENANCE MATERIAL COST $ 148.70

PROPULSION MAINTENANCE LABOR COST = $ 33.31

TOTAL COST = $ 2549.3-

_-,; i._4M r, l_.V

AIRCRAFT FIRST PRICE = $ 3578783

F'ROPULSION SYSTEM PRICE/ENG = $

DIRECT OPERATING COST PER HOUR =

COST PER SEAT MILE =

464086

1710.95 $/HR

.797 $/seat-nm

113
TABLE 26. Pusher Propfan ATA Analysis at 400 Hours Utilization

ADVANCED (1990) PUSHER PROPFAN

MISSION STAGE LENGTH = 400.0 NM YEARLY UTILIZATION = 400 HOURS

NUMBER OF PASSENGER SEATS = 8 AIRCRAFT EMPTY WEIGHT = 8274.0 LB

_ TURBOPROP ENGINE _

PROP DIAMETER 8.000 FT # OF BLADES = 6

ENGINE SL THERMODYNAMIC SHP 2714.0

BLOCK TIME =

BLOCK FUEL =

1. 495 HR

I_,7_-. 0 LB

MI_-O,._ ELEMENT COST BR_Ar.DO.,N:

CREW COST = $ 287.04

FUEL COST = $ 394.75

INsURaNCE COST = $ 157.88

DEPRECIATION COST = $ 1592.6q

AIRFRAME MAIN,LJ_A,_uE MATERIAL COST = $ 78.62

AI_.rR._I_ MAINTENANCE LABOR COST = $ 87.24

PROPULSION MAINTENANCE MATERIAL COST = $ 160.48

PROPULSION MAINTENANCE LABOR COST = $ 50.22

TOTAL COST = $ 2808.83

SUMMARY:

AIRCRAFT FIRST PRICE = $ 4224155

PROPULSION SYSTEM PRICE/ENG = $

DIRECT OPERATING COST PER HOUR =

COST PER SEAT MILE =

624857

1878.82 $/HR

.878 S/seat-rim

114
TABLE27. Tractor Propfan ATA Analysis at 400 Hours Utilization

_L'_H,,U_U (1990) TRACTOR PROPFAN

MISSION STAGE LENGTH = 400.0 NM YEARLY UTILIZATION = 400 HOURS

NUMBER OF PASSENGER SEATS = 8 AIRCRAFT EMPTY WEIGHT = 8802.0 LB

_ TURBOPROP ENGINE _

PROP DIAMETER 8.000 FT # OF BLADES = 6

ENGINE SL THERMODYNAMIC SHP 2693.0

BLOCK TIME =

BLOCK FUEL =

1.502 HR

1401.0 LB

MISSION ELEMENT COST BREAKDOWN:

CREW COST = $ =_oo.38

FUEL COST = $ 397.30

INSURANCE COST = $ 162.75

DEPRECIATION COST = $ 1637.17

AIRFRAME MAINTENANCE MATERIAL COST = $ 89.54

AIRFRAME MAINTENANCE LABOR COST = $ 90.24

PROPULSION MAINTENANCE MATERIAL COST = $ 160.60

PROPULSION MAINTENANCE LABOR COST = $ 50.47

TOTAL COST = $ 2876.46

SUMMARY:

AIRCRAFT FIRST PRICE = $ 4334136

PROPULSION SYSTEM PRICE/ENG = $

DIRECI OPERATING COST PER HOUR =

COST PER SEAT MILE =

620967

1915.09 $/HR

.899 $/seat-nm

115
TABLE28. Current Turbofan ATA Analysis at 3000 Hours Utilization

CURRENT (1984) TURBOFAN

MISSION STAGE LENGTH : 400.0 NM ...... tEMr, L_ UTILI ZATION = 3000 HOURS

NUMBER OF PASSENGER SEATS = 8 AIRCRAFT EMPTY WEIGHT = 8676.0 LB

_I TURBOFAN ENGINE I_

ENGINE SL THERMODYNAMIC THRUST 3257.0 LB

BLOCK TIME = 1.508 HR

BLOCK FUEL = 2089.0 LB

MISSION ELEMENT COST BREAKDOWN:

CREW COST = $

FUEL COST = $

INouR_N_ COST = $

DEPRECIATION COST = $

AIRFRAME MAINTENANCE MATERIAL COST = $

AIRFRAME MAINTENANCE LABOR COST = $

PROPULSION MAINTENANCE MATERIAL COST = $

PROPULSION MAINTENANCE LABOR COST = $

289.54

592.4(}

18.08

180.04

91.07

93.47

154.77

.... 73

TOTAL COST = $ 1453.10

SUMMARY:

AIRCRAFT FIRST PRICE = $ 3596221

PROPULSION SYSTEM PRICE/ENG = $

DIRECT OPERATING COST PER HOUR =

COST PER SEAT MILE =

455134

963.60 $/HR

.454 S/seat-rim

116
TABLE 29. Advanced Turbofan ATA Analysis at 3000 Hours Utilization

ADVANCED (1990) TURBOFAN

MISSION STAGE LENGTH = 400.0 NM YEARLY UTILIZATION = 3000 HOURS

NUMBER OF PASSENGER SEATS = 8 AIRCRAFT EMPTY WEIGHT = 7426.0 LB

_ TURBOFAN ENGINE _

ENGINE SL THERMODYNAMIC THRUST 3067.0 LB

BLOCK TIME =

BLOCK FUEL =

1.490 HR

1625.0 LB

MISSION ELEMENT COST BREAKDOWN:

CREW COST = $

FUEL COST = $

INSURANCE COST = $

DEPRECIATION COST = $

AIRFRAME MAINTENANCE MATERIAL COST = $

AIRFRAME MAINTENANCE LABOR COST = $

PROPULSION MAINTENANCE MATERIAL COST = $

PROPULSION MAINTENANCE LABOR COST = $

286.08

460.82

17.77

177.36

71.70

85.23

148.70

33.31

TOTAL COST = $ 1280.96

SUMMARY:

AIRCRAFT FIRST PRICE = $ 3578783

PROPULSION SYSTEM PRICE/ENG = $

DIRECT OPERATING COST PER HOUR =

COST PER SEAT MILE =

464086

859.71 $/HR

.400 $!seat-nm

117
TABLE 30. Pusher Propfan ATA Analysis at 3000 Hours Utilization

ADVANCED (1990) PUSHER PROPFAN

MISSION STAGE LENGTH = 400.0 NM YEARLY UTILIZATION = 3000 HOURS

NUMBER OF PASSENGER SEATS = 8 AIRCRAFT EMPTY WEIGHT = 8274.0 LB

_ TURBOPROP ENGINE _

PROP DIAMETER 8.0©0 FT # OF BLADES = 6

ENGINE SL THERMODYNAMIC SHP 2714.0

BLOCK TIME =

BLOCK FUEL =

1.495 HR

1392.0 LB

MTSSION ELEMENT COST BR=Ar.DOe_N.

CREW COST = $

FUEL COST = $

NooR_N_E COST = $

DEPRECIATION COST = $

AIRFRAME MAINTENANCE MATERIAL COST = $

AIRFRAME MAINTENANCE LABOR COST = $

PROPUI_SION MAINTENANCE MATERIAL COST = $

PROPULSION MAINTENANCE LABOR COST = $

287.04

394.75

21.05

212.35

78.62

87.24

16 r).48

50. --"_

TOTAL COST = $ 1291.75

SUMMARY :

AIRCRAFT FIRST PRICE = $ 4224155

PROPULSION SYSTEM PRICE/ENG = $

DIRECT OPERATING COST PER HOUR =

COST PER SEAT MILE =

624857

864.05 $/HR

.404 S/seat-rim

118
TABLE 31. Tractor Propfan ATA Analysis at 3000 Hours Utilization

ADVANCED (1990) TRACTOR PROPFAN

MISSION STAGE LENGTH = 400.0 NM YEARLY UTILIZATION = 3000 HOURS

NUMBER OF PASSENGER SEATS = 8 AIRCRAFT EMPTY WEIGHT = 8802.0 LB

!I TURBOPROP ENGINE I_

PROP DIAMETER 8.000 FT # OF BLADES = 6

ENGINE SL THERMODYNAMIC SHP 2693.0

BLOCK TIME =

BLOCK FUEL =

1.502 HR

1401.0 LB

MISSION ELEMENT COST BREAKDOWN:

CREW COST = $

FUEL COST = $

INSURANCE COST = $

DEPRECIATION COST = $

AIRFRAME MAINTENANCE MATERIAL COST = $

AIRFRAME MAINTENANCE LABOR COST = $

PROPULSION MAINTENANCE MATERIAL COST = $

PROPULSION MAINTENANCE LABOR COST = $

288.38

397.30

21.70

218.29

89.54

9£).24

160.60

50.47

TOTAL COST = $ 1316.53

SUMMARY:

AIRCRAFT FIRST PRICE = $ 4334136

PROPULSION SYSTEM PRICE/ENG = $

DIRECT OPERATING COST PER HOUR =

COST PER SEAT MILE =

620967

876.52 $/HR

.411 $/seat-nm

119
A2.0 CORPORATE COST MODEL

The corporate model was developed for this study utilizing the cost

elements of the ATA model. These elements, however, were modified somewhat

to facilitate evaluation of the cost of capital and tax advantage typically

experienced by business aircraft owners. Further, the depreciation formula

used in the ATA model was replaced by a cash flow analysis for a five-year

period to generate an overall five year cost of operation. This simple

cash flow analysis provides a cost of operation analysis that is more

typical of General Aviation business aircraft and provides a better

understanding of the cost factors which impact the business aircraft

market, particularly interest rates, insurance rates, aircraft prices, and

tax Iaw.

The operating cost was broken into fixed and variable costs as shown in

Table 32. Each of the corresponding ATA elements used to calculate the

fixed and variable costs are shown in Table 32. The fixed costs were

assumed to include only crew cost and insurance cost determined on an

annual basis. Other types of fixed costs such as hangar cost, airport use

fees, aircraft cleaning cost, office supplies, charts, catering, etc., were

TABLE 32. Operating Cost Elements

OPERATING COST

ANNUAL FIXED

Crew Cost

Cc x U/t b

Insurance Cost

C i x U/tb

VARIABLE

Fuel Cost

Cf/tb

Maintenance Cost

/tb, CDm/tb,

al/tb, Cam/tb

120
not considered but did not impact the relative comparison in this study.

The variable costs, similar to the ATAmodel, were determined on a flight

hour basis. The fuel cost was based on an average fuel price, as in the

ATAmodel, of $1.90 per gallon for JET-A (Reference 18). Unlike airline

operators who can obtain bulk fuel at approximately $1/gallon, business

aircraft operators are forced to purchase fuel at higher retail prices that

range from $1.65/gallon to $2.40/gallon depending on location. The

airframe and propulsion system maintenance costs, as in the ATAmodel, were

derived from data provided by Pratt and Whitney, Canada and Hamilton

Standard and included cyclic and flight hour cost factors. Additionally,

the engine cost factors also included an allowance for engine overhaul.

The cash flow analysis which yields the cost of operation for the Corporate

Cost Model is broken into five steps and is illustrated in Figure 70.

Step 1. Based on the aircraft and propulsion system characteristics

the aircraft acquisition cost (or aircraft first price) is determined using

the method described at the end of the Appendix, Section A3.0. The tax

depreciation basis and monthly loan paymentare then determined for use in

Steps 3 and 4.

Step 2. The yearly operating cost (or ops cost) is the next element

determined. It is broken into two sub-elements: a variable cost based on

flight hours and a yearly fixed cost. The yearly variable cost is

determined using the aircraft annual utilization hours. The yearly ops

cost is input both to Step 3 and to Step 4.

Step 3. Next the tax savings is determined for each year with the

assumed50%tax bracket. The tax savings is input into Step 4.

Step 4. The yearly net cash flow is then calculated. Negative cash

flow is defined as dollars being spent by the aircraft owner.

Step 5. Finally, the sumof each year's bottom-line cash flow in Step

4 is calculated to yield the total cost of operation for 5 years. Based on

the design mission's length, block time required, and the aircraft's annual

utilization, the hourly cost of operation or the cost per seat per nautical

mile can be calculated.

121
AIRCRAFT

FIRST

PRICE

_3,596,221.4 c

STEP I !

CASH FLOW ANALYSIS

CASH FLOW OUT (-) YEARS

i 2 3 4

_DOWN PAYMENT 179,811 ............

Lm-ANNUAL PAYMENT 588,187 588,187 588,187 588,187

COST OF OPERATION 368,851 368,851 368,851 368,851

_REPAY OF LOAN BAL ................

TAX ON RESALE GAIN ................

CASH FLOW IN (+)

INVSTMT TAX CREDIT

TAX SAVINGS

RESALE

I NET CASH FLOW

$ OUT (-)

I

I

AIRCRAFT

FIRST

PRICE

$3,596,221.49

STEP 1L____

359,622 ............

640,573 748,203 717,662 702,497

136,654 208,834 239,375 254,541

5 YEAR NET COST = $2,055,853

TAX SAVINGS

YEARS

I 2 3 4

ANNUAL OPS COST 368,851 368,851 368,851 368,851

_DEPRECIATION 512,462 751,610 717,446 717,446

_INTEREST 399,833 375,945 349,028 318,696

TOTAL 1,281,145 1,496,406 1,435,324 1,404,993

TAX SAVINGS (50%)

STEP 3 I

5

588,187

368,851"_,_,_

2,203,497_

1,258,678_

685,407

2,517,355

1,216,449 /

/

/

ANNUAL

OPERATING COST

FIXED + VARIABLE

TOTAL = $368,851

STEP 2 1

368,851-4 "j

717,446

284,518

1,370,815

$640,573 $748,203 $717,662 $702,497 $685,407

FIGURE 70. Steps to Evaluating the Cash Flow Utilized in the

Corporate Cost Model

122
A2.1 COSTOFCAPITALANDTAXASPECTS

The cost of capital, for this study, was simply the interest paid each year

on the loan. A value of 12%was assumedas a typical 1984 loan rate, and

the loan was assumed to be a ten year direct reduction loan with a balloon

payoff at the end of five years. A 5%down payment on the loan was also

assumed.

The Tax Equity and Fiscal Responsibility Act of 1982 allows for two

alternative methods for depreciation of capital equipment such as aircraft.

The taxpayer may elect to take an investment tax credit of 8% of the

initial purchase price of the aircraft for the first year, using 100% of

the purchase price as the depreciation basis over five years, or he may

elect to take a 10%tax credit and then use 95% of the purchase price as

the depreciation basis. In the purchase of a General Aviation aircraft the

option selected, of course, depends on the tax situation of the purchaser.

However, the 10%option is most frequently selected and was chosen for this

study as typical. Using 95% of the purchase price as a basis of

depreciation, the depreciation for each year is calculated as follows: year

1 - 15%of basis, year 2 - 22%of basis, year 3, 4, and 5 - 21%of basis.

To obtain the resultant tax benefits of ownership on the overall cash flow,

the annual operations cost, which is considered a deductable expense, is

added to the depreciation and interest to yield the total deductions.

These deductions, depending on the taxpayer's tax bracket, will result in a

reduction of tax paid. Most taxpayers purchasing General Aviation aircraft

of the class examined in this study are subject to a 50% tax rate. This

rate was assumed for this study. Thus the resultant tax benefits of

ownership on the taxpayer's overall cash flow is 50% of the allowable

deductions.

A2.2 CASH FLOW ANALYSIS

The cash flow is then examined assuming ownership of the aircraft for five

years. The cash flow "out" is combined with the cash flow "in" to yield

the "net" cash flow, where a negative number represents dollars spent by

the aircraft owner. Several assumptions were made in this cash flow model

to simplify the analysis and to provide an equivalent basis of comparison

for the aircraft in the study. It was assumed that the aircraft was sold

at the end of five years at which time the resale value was 70% of the

original purchase price. The 70% resale value is typical of the price on

used General Aviation turbofan aircraft whereas turboprop aircraft have a

123
typically lower resale value around 60% after five years. Therefore, the

assumption made in this study was that propfans would behave more like

turbofans in the marketplace than like turboprops. Also, it was assumed

that the aircraft owner, at the end of five years, did not purchase another

more expensive aircraft. Thus the dollar gain on the resale of the fully

depreciated aircraft is taxable as ordinary income. This is not usually

the case for a business aircraft owner. Owners typically "trade up" in

equipment, thus modifying their next aircraft's depreciation basis by the

sale price of their present aircraft. This significantly complicates the

cash flow analysis and clouds the propulsion system comparison being made

in this study. The assumption that the owner would sell the aircraft at

the end of five years without a subsequent purchase of another aircraft

provided a simplification to the cash flow method which still allowed a

valid (yet conservative in cost per seat per nautical mile) comparison

between configurations without introducing unnecessary complication.

The resultant net cash flow for each year is summed for the five year

period, yielding a five year cost of operation. Then, depending on the

operator's utilization, a cost per hour or a cost per seat per nautical

mile can be determined.

A2.3 COST OF OPERATION BREAKDOWNS

Tables 33 through 36 show the cost spread sheets for each of the four study

aircraft. These costs are evaluated for an annual utilization of 400

hours.

124
TABLE 33. Current Turbofan Corporate Cost Analysis

CURRENT TURBOFAN

NUMBER OF PASSENGER SEATS = 8 ENGINE SL THERMODYNAMIC THRUST 3257.0 LB

TYPICAL MISSION LENGTH = 400.0 NM YEARLY UTILIZATION = 400.0 HR

BLOCK TIME FOR MISSION = 1.5080 HR BLOCK FUEL FOR MISSION = 2089.0 LB

AIRCRAFT FIRST PRICE = $ 3596221.49 PROPULSION SYSTEM PRICE(EA) = $ 455134.43

ASSUMED FINANCING: 12% INTEREST. 95% FINANCED. 10% INVESTMENT TAX CREDIT

TAX DEPRECIATION BASIS = $ 3416410.41 10 YEAR LOAN

YEARLY OPERATING COST BREAK DOWN

$_VARIABLE$_ *_FIXED$_

FUEL COST = $ 392.84/HR CREW COST = $ 76800.00

AIRFRAME MAINTENANCE MATERIAL COST = $ 60.39/HR INSURANCE = $ 35962.21

AIRFRAME MAINTENANCE LABOR COST = $ 61.98/HR

PROPULSION MAINTENANCE MATERIAL COST = $ 102.64/HR

PROPULSION MAINTENANCE LABOR COST = $ 22.37/HR

YEARLY VARIABLE COST = $ 256088.44 YEARLY FIXED COST = $ 112762.21

TAX SAVINGS BREAKDOWN YEARS

1 2 3 4 5

ANNUAL OPS COST .... 368851 368851 368851 368851 368851

DEPRECIATION .... 512462 751610 717446 717446 717446

INTEREST ...... 399833 375945 349028 318696 284518

TOTAL ...... 1281145 1496406 1435324 1404993 1370815

TAX SAVINGS (50%)... 640573 748203 717662 702497 685407

CASH FLOW OUT(-)

DOWNPAYMENT .......

ANNUAL PAYMENT .....

ANNUAL OPS COST ......

REPAY OF LOAN BAL ....

TAX ON RESALE GAIN ....

CASH FLOW IN(+)

1

179811

588187

368851

INVSTMT TAX CREDIT ... 359622

TAX SAVINGS ..... 640573

RESALE ....

NET CASH FLOW

$ OUT (-) ....... 136654

CASH FLOW

YEARS

2 3 4

588187 588187 588187

368851 368851 368851

748203 717662 702497

588187

368851

2203497

1258678

685407

2517355

208834 239375 254541 1216449

5 YEAR COST OF OPERATION = $ 2055853.31

BASED ON THE TYPICAL MISSION:

ESTIMATED MILES FLOWN

IN 5 YEARS = 530504.0 NM

ESTIMATED COST/HR = $ 1027.93/HR

ESTIMATED $/SEAT-NM = .484

125
TABLE 34. Advanced Turbofan Corporate Cost Analysis

ADVANCED TURBOFAN

NUMBER OF PASSENGER SEATS = 8 ENGINE SL THERMODYNAMIC THRUST 3067.0 LB

TYPICAL MISSION LENGTH = 400.0 NM YEARLY UTILIZATION = 400.0 HR

BLOCK TIME FOR MISSION = 1.4900 HR BLOCK FUEL FOR MISSION = 1625.0 LB

AIRCRAFT FIRST PRICE = $ 3578783.17 PROPULSION SYSTEM PRICE(EA) = $ 464086.05

ASSUMED FINANCING: 12% INTEREST. 95% FINANCED. 10% INVESTMENT TAX CREDIT

TAX DEPRECIATION BASIS = $ 3399844.01 10 YEAR LOAN

YEARLY OPERATING COST BREAK DOWN

**VARIABLE FIXED** 

FUEL COST = $ 309.28/HR CREW COST = $ 76800.00

AIRFRAME MAINTENANCE MATERIAL COST = $ 48.12/HR INSURANCE = $ 35787.83

AIRFRAME MAINTENANCE LABOR COST = $ 57.20/HR

PROPULSION MAINTENANCE MATERIAL COST = $ 99.80/HR

PROPULSION MAINTENANCE LABOR COST = $ 22.36/HR

YEARLY VARIABLE COST = $ 214699.01 YEARLY FIXED COST = $ 112587.83

TAX SAVINGS BREAKDOWN YEARS

1 2 3 4 5

ANNUAL OPS COST .... 327287 327287 327287 327287 327287

DEPRECIATION .... 509977 747966 713967 713967 713967

INTEREST ...... 397894 374122 347335 317151 283138

TOTAL ...... 1235158 1449375 1388589 1358405 1324393

TAX SAVINGS (50%)... 617579 724687 694295 679202 662196

CASH FLOW OUT(-)

DOWNPAYMENT .......

ANNUAL PAYMENT .....

ANNUAL OPS COST ......

REPAY OF LOAN BAL ....

TAX ON RESALE GAIN ....

CASH FLOW IN(+)

INVSTMT TAX CREDIT ...

TAX SAVINGS .....

RESALE ....

NET CASH FLOW

* CASH FLOW *

YEARS

1 2 3 4

178939

585335 585335 585335 585335

327287 327287 327287 327287

357878

617579 724687 694295 679202

5

585335

327287

2192812

1252574

662196

2505148

$ OUT (-) ....... 116103 187934 218327 233419 1190663

5 YEAR COST OF OPERATION = $ 1946446.26

BASED ON THE TYPICAL MISSION:

ESTIMATED MILES FLOWN

IN 5 YEARS = 536912.8 NM

ESTIMATED COST/HR = $ 973.22/HR

ESTIMATED $/SEAT-NM = .453

126
TABLE35. Pusher Propfan Corporate Cost Analysis

PUSHER PROPFAN

PROP DIAMETER = 8.000 FT # OF BLADES = 6

NUMBER OF PASSENGER SEATS = 8 ENGINE SL THERMODYNAMIC SHP = 2714.0 ESHP

TYPICAL MISSION LENGTH = 400.0 NM YEARLY UTILIZATION = 400.0 HR

BLOCK TIME FOR MISSION = 1.4950 HR BLOCK FUEL FOR MISSION = 1392.0 LB

AIRCRAFT FIRST PRICE = $ 4224155.31 PROPULSION SYSTEM PRICE(EA) = $ 624857.09

ASSUMED FINANCING: 12% INTEREST, 95% FINANCED, 10% INVESTMENT TAX CREDIT

TAX DEPRECIATION BASIS = $ 4012947.54 10 YEAR LOAN

YEARLY OPERATING COST BREAK DOWN

IIVARIABLEII _FIXEDIZ

FUEL COST = $ 264.04/HR CREW COST = $ 76800,00

AIRFRAME MAINTENANCE MATERIAL COST = $ 52.59/HR INSURANCE = $ 42241,55

AIRFRAME MAINTENANCE LABOR COST = $ 58.36/HR

PROPULSION MAINTENANCE MATERIAL COST = $ 107.35/HR

PROPULSION MAINTENANCE LABOR COST = $ 33.59/HR

YEARLY VARIABLE COST = $ 206370.91 YEARLY FIXED COST = $ 119041.55

TAX SAVINGS BREAKDOWN YEARS

1 2 3 4 5

ANNUAL OPS COST .... 325412 325412 325412 325412 325412

DEPRECIATION .... 601942 882848 842719 842719 842719

INTEREST ...... 469648 441589 409971 374344 334198

TOTAL ...... 1397002 1649850 1578102 1542475 1502329

TAX SAVINGS (50%)... 698501 824925 789051 771237 751164

CASH FLOW OUT(-)

DOWNPAYMENT .......

ANNUAL PAYMENT .....

ANNUAL OPS COST ......

REPAY OF LOAN BAL ....

TAX ON RESALE GAIN ....

CASH FLOW IN(+)

INVSTMT TAX CREDIT ...

TAX SAVINGS .....

RESALE ....

NET CASH FLOW

$ OUT (-) .......

! CASH FLOW

YEARS

1 2 3 4

211208

690890 690890 690890 690890

325412 325412 325412 325412

422416

698501 824925 789051 771237

690890

325412

2588248

1478454

751164

2956909

106593 191377 227251 245065 1374931

5 YEAR COST OF OPERATION = $ 2145217.00

BASED ON THE TYPICAL MISSION:

ESTIMATED MILES FLOWN

IN 5 YEARS = 535117.1 NM

ESTIMATED COST/HR = $ 1072.61/HR

ESTIMATED $/SEAT-NM = .501

127
TABLE 36. Tractor Propfan Corporate Cost Analysis

TRACTOR PRDPFAN

PROP DIAMETER = 8.000 FT # OF BLADES = 6

NUMBER OF PASSENGER SEATS = 8 ENGINE SL THERMODYNAMIC SHP = 2693.0 ESHP

TYPICAL MISSION LENGTH = 400.0 NM YEARLY UTILIZATION = 400.0 HR

BLOCK TIME FOR MISSION = 1.5020 HR BLOCK FUEL FOR MISSION = 1401.0 LB

AIRCRAFT FIRST PRICE = $ 4334135.65 PROPULSION SYSTEM PRICE(EA) = $ 620967.04

ASSUMED FINANCING: 12% INTEREST. 95% FINANCED, 10% INVESTMENT TAX CREDIT

TAX DEPRECIATION BASIS = $ 4117428.87 10 YEAR LOAN

YEARLY OPERATING COST BREAK DOWN

IIVARIABLEI! IIFIXEDII

FUEL COST = $ 264.51/HR CREW COST = $ 76800.00

AIRFRAME MAINTENANCE MATERIAL COST = $ 59.62/HR INSURANCE = $ 43341.36

AIRFRAME MAINTENANCE LABOR COST = $ 60.08/HR

PROPULSION MAINTENANCE MATERIAL COST = $ 106.92/HR

PROPULSION MAINTENANCE LABOR COST = $ 33.60/HR

YEARLY VARIABLE COST = $ 209895.13 YEARLY FIXED COST = $ 120141.36

TAX SAVINGS BREAKDOWN YEARS

1 2 3 4 5

ANNUAL OPS COST .... 330036 330036 330036 330036 330036

DEPRECIATION .... 617614 905834 864660 864660 864660

INTEREST ...... 481875 453086 420645 384090 342899

TOTAL ...... 1429526 1688957 1615342 1578787 1537595

TAX SAVINGS (50%)... 714763 844478 807671 789393 768798

CASH FLOW OUT(-)

DOWNPAYMENT .......

ANNUAL PAYMENT .....

ANNUAL OPS COST ......

REPAY OF LOAN BAL ....

TAX ON RESALE GAIN ....

CASH FLOW IN(+)

INVSTMT TAX CREDIT ...

TAX SAVINGS .....

RESALE ....

NET CASH FLOW

CASH FLOW

YEARS

1 2 3 4

216707

708878 708878 708878 708878

330036 330036 330036 330036

433414

714763 844478 807671 789393

$ OUT (-) ....... 107444 194436 231243 249521

708878

330036

2655636

1516947

768798

3033895

1408805

5 YEAR COST OF OPERATION = $ 2191448.94

BASED ON THE TYPICAL MISSION:

ESTIMATED MILES FLOWN

IN 5 YEARS = 532623.2 NM

ESTIMATED COST/HR = $ I095.72/HR

ESTIMATED $/SEAT-NM = .514

128
A3.0 AIRCRAFTACQUISITIONCOST

The formulation for the aircraft selling price consists of three simple

elements; an airframe price, a propulsion system price, and an avionics

package price. Each element includes the effects of manufacturing costs

and overhead. Each is fully burdened and the airframe and propulsion

systems include a price margin index. The price margin index is a

statistical approach to model such items such as factory profit, liability

coverage, typical option pricing, and dealer markup. The pricing method

does not incorporate any discounting or pricing strategy factors and does

not represent any specific manufacturers method of pricing (including

Beech). Rather, because of the difficulty in accurately modeling these

effects of market and pricing strategies as well as the amortization of

manufacturing costs, this formulation is a statistical, rough order of

magnitude (ROM) estimate of the aircraft's initial selling price.

Additionally, the pricing method has been tailored for this study utilizing

1984 published price data for the class of business aircraft examined in

this study and is not necessarily representative of any other class of

aircraft.

The airframe pricing for current day technology represents current 1984

manufacturing processes. The airframe pricing for the 1990 level

technology incorporates the higher material and manufacturing costs

associated with advanced materials. The airframe price calculation, which

incorporated everything except propulsion system and avionics, included all

other items such as electrical equipment, environmental systems, hydraulic

systems, interiors, tires, brakes, bearings, cables, etc. These have been

averaged for typically equipped aircraft and are included in the airframe

average price factor; $132/LB for current technology, $158/LB for 1990

technology.

The turboprop and turbofan engine pricing was provided by Pratt and

Whitney, Canada. The propfan pricing was provided by Hamilton Standard.

The pricing in all cases was related to propulsion system size. Maximum

thermodynamic capability at sea level was used as an indicator of size in

each case; static thrust for the turbofans, static horsepower for the

turboprops. Figure 71 shows the functional relationship of price to

thermodynamic capability. The propfan curve, in addition to engine price,

also includes the price of the propfan and gear box. The avionics cost

selected is representative of a typically equipped business turbofan with a

single electronic flight instrument system (EFIS).

129
800"

700"

600'

500-

PROPULSION

PRICE

FACTOR 400"

(Kp)

$1000 3O0"

200"

i00"

! | I '

0 zo'oo 20'00 30'o0 40'o0

MAXIMUM THERMODYNAMIC THRUST (FN IN LB FOR TURBOFANS) OR

POWER (PT IN SHP FOR PROPFANS)

FIGURE 71. Propulsion System Acquisition Cost Factors

All pricing is done in 1984 dollars. The pricing formulation is as

fol lows:

AIRCRAFT SELLING PRICE

Ct = C a + Cp + Cae

where

Ct = aircraft selling price ($)

Ca = complete airframe price ($)

Cp = total propulsion system price ($)

Cae = price of typical optional avionics and equipment ($)

130
AIRFRAME PRICE

where

Ca = Im Ka Wa

Im = price margin index = 1.75

Wa = aircraft airframe weight (LB)

Ka = average airframe price factor ($/LB)

for current materials (1984)

Ka = 132

for advanced materials (1990)

Ka = 158

PROPULSION SYSTEM PRICE

for turbofans

Cp = Im Kp F N Ne

where

Im = price margin index = 1.75

FN = total maximum thermodynamic sea level static thrust (LB)

Kp = propulsion system price factor (Figure 71) ($)

Ne = number of engines

for propfans

Cp = Im Kp Pt Ne

where

Im = price margin index = 1.75

131
Pt = total maximumthermodynamic sea level static power (SHP)

Kp = propulsion system price factor (Figure 71) ($)

Ne = number of engines

AVIONICS PRICE

Cae = price of typical optional avionics and equipment ($)

Cae = 520,000

132
SYMBOLSANDABBREVIATIONS

ACEE

AF

AR

ATA

ATP

B

BAL

BLK

Ca

Cae

Cal

Cam

Cc

Cd

CD

Cf

CG

Ci

CL

Cp

Cpl

Cpm

C r

Ct

Cv

D, DIA

Da

dB

dBA

DOC

E3

ECI

EFIS

EPNdB

Aircraft Energy Efficiency Program

blade activity factor

aspect ratio

Air Transport Association of America

Advanced Turboprop Program

overhead burden factor

balance

block

complete airframe price

price of typical optional avionics and equipment

airframe maintenance labor cost

airframe maintenance material cost

crew cost

cost of depreciation

drag coefficient

cost of fuel

center of gravity

cost of hull insurance

lift coefficient

propulsion system price

engine or engine/propeller maintenance labor cost

engine or engine/propeller maintenance material cost

rudder chord

aircraft initial selling price

vertical tail chord

diameter

depreciation period

decibels

A-weighted decibels

direct operating cost

Energy Efficient Engine

Energy Component Improvement

electronic flight instrument system

effective perceived noise level

133
F, Fb

FN

FAR

FPM

FPS

FS

FT

FWD

GA

GAL

Grad

HP

HR

HSD

Hz

IFR

Im

IN

INVSTMT

IR

ISA

K1

K2

K3

K4

K5

K6

K7

K8

Ka

Kc

Kp

KT

KTAS

LAP

LB

block fuel

thermodynamic thrust

Federal Aviation Regulations

feet per minute

feet per second

fuselage station

feet

forward

General Aviation

gallon

gradient

horsepower

hours

Hamiltion Standard Division

Hertz

Federal Aviation Regulation Instrument Flight Rules

price margin index

inches

investment

annual insurance rate

International Standard Atmosphere

labor manhours per flight hour

labor manhours per flight cycle

airframe material hourly factor

airframe material cyclic factor

labor manhours per flight hour per engine

labor manhours per flight cycle per engine

propulsion system material cost per flight hour

propulsion system material cost per flight cycle

average airframe price factor

crew hourly rate

propulsion system price factor

knots

knots true airspeed

Large-Scale Advanced Propeller

pound

134
LBM

LBT

LR

M

MAINT

MAPS

MAX

Mcr

MIN

NASA

NB

NBAA

Nc

Ne

NM

OASPL

OEW

ops

P&WC

Pf

Pt

PTA

ROM

RPM

RQMTS

S

SEC

SFC

SHP

SL

SLS

SR

STD

T

t/c

pounds mass

pounds thrust

labor rate

Mach

maintenance

Multiple Application Propfan Studies

maximum

cruise Mach number

minute

National Aeronautics and Space Administration

number of blades

National Business Aircraft Association

number of crew

number of engines

nautical mile

overall sound pressure level

operational empty weight

operations

Pratt & Whitney, Canada

price of fuel

thermodynamic horsepower

Propfan Test Assessment

rough order of magnitude

revolutions per minute

requirements

wing area

seconds

specific fuel consumption

shaft horsepower

sea Ievel

sea level static

single rotation

standard

thrust

thickness'ratio

135
tb

tf

tg

TO

TOFL

TOGW

TSFC

U

UTC

Vcr

VFR

VMC

WA

Wa

YR

0

oc

OF

n

block time

flight time

ground time

takeoff

takeoff field length

takeoff gross weight

thrust specific fuel consumption

annual utilization

United Technologies Corporation

cruise speed

Federal Aviation Regulation Visual Flight Rules

minimum control airspeed

actual airflow

aircraft airframe weight

year

degrees

degrees Centigrade

degrees Fahrenheit

inches

136
REFERENCES

1. Whitlow, J. B., Jr. and Sievers, G. K.: Fuel Savings Potential of the

NASAAdvancedTurboprop Program. NASATM-83736, Sept. 1984.

. Dugan, James F., Miller, Brent A., Graber, Edwin J., and Sagerser,

David A.: The NASA High-Speed Turboprop Program. SAE 801120, Oct.

1980.

3. Dreyfuss, Henry: The Measure of Man, Human Factors in Design.

Whitney Publications, 1967.

4. Hamilton Standard: 6, 8, and 10 Bladed Propeller Parametric Data

Package. HSD Report No. SPO8A82, May 1982.

5. Hamilton Standard: Ten Blade Prop-Fan Parametric Data Package. HSD

Report No. SPO7A82, May 1982.

. Griffin, C. F., Dunning, E. G.: Development of an Advanced Composite

Aileron for the L-1011 Transport Aircraft. NASA CR-3517, February

1982.

o Lockheed-California Company: Advanced Manufacturing Development of a

Composite Empennage Component for L-1011 Aircraft. NASA CR-165885,

May 1982.

. Boeing Commercial Airplane Company: Design, Ancillary Testing,

Analysis, and Fabrication Data for the Advanced Composite Stabilizer

for Boeing 737 Aircraft. NASA CR-3648, April 1983.

. Mixson, John S. and Powell, Clemans A.:

Interior Noise of Propeller Aircraft.

Conference, AIAA Paper 2349, Oct. 1984.

Review of Recent Research on

AIAA/NASA 9th Aeroacoustics

10. Beranek, L. L.:

York, 1971.

Noise and Vibration Control. McGraw Hill Co., New

137
11. Cockburn, J. A. and Jolly, A. C.: Structural-Acoustic Responses,

Noise Transmission Losses and Interior Noise Levels of an Aircraft

Fuselage Excited by Random Pressure Fields. AFFDL-TR-68-2, Aug. 1968.

12. Society of Automotive Engineers:

and Far-Field Propeller Noise.

AIR-1407, May 1977.

Prediction Procedure for Near-Field

SAE Aerospace Information Report

13. Block, P. J. W.: Installation Noise Measurements of Model SR and CR

Propellers. NASA TM-85790, May 1984.

14. Block, P. J. W.: Noise Generated by a Propeller in a Wake. NASA

TM-85794, May 1984.

15. Owens, Robert E. and Ferguson, W. Wade: Some Aspects of Prop-Fan

Propulsion System Analysis. SAE 821358, Oct. 1982.

16. Air Transport Association of America: Standard Method of Estimating

Comparative Direct Operating Costs of Turbine Powered Transport

Airplanes. December 1967.

17. Kraus, E. F., Mall, O. D., Awker, R. W., and Scholl, J. W.:

Application of Advanced Technologies to Small, Short-Haul Transport

Aircraft (STAT). NASA CR-152362, Aug. 1982.

18. Aircraft Owners and Pilots Association:

Frederick, Maryland, Nov. 1984.

AOPA Fuel Watch. AOPA,

138
1. Report No.'

NASA CR 175020

"4, Title and Subtitle

Evaluation of Propfan

to General Aviation

2. Government Accession No. 3. Recipient's C_talog No.

5. Report Date

I'IU I _ J. JUU Propulsion Applied

7. Author(s)

R. W. Awker

Perfuming Organization Name and Addre=

Beech Aircraft Corporation

9709 E. Central

Wichita, KS 67201-0085

12. Sponsoring Agency Name and Address

National Aeronautics and

Lewis Research Center

Cleveland, OH 44135

Space Administration

6. Performing Organization Code

8. Performing Organization Report No.

10. Work Unit No.

11. Contract or Grant No.

NAS3-24349

13. Type of Report and Period Covered

Contractor Report

"_14. Sponsoring Agency CoOe

15. Supplementary Notes

Contract Project Manager: Susan M. Johnson

Advanced Turboprop Project Office

NASA Lewis Research Center

Cleveland, OH 44135

16. Abstract

This report presents the results of a study to evaluate propfan propulsion on

business aircraft. Comparisons, in terms of cost and performance, were made

between propfan propulsion systems and conventional turbofan propulsion

systems on a typical business aircraft. In addition, configuration and cost

sensitivity studies were conducted to further assess the potential of propfan

propulsion.

17. Key Words (Suggested by Author(s))

PROPFAN, HIGH SPEED PROPELLERS,

MAPS, BUSINESS AIRCRAFT,

GENERAL AVIATION, TURBOPROP AIRCRAFT

18. Distribution Statement

Uncl assi fied-Unl imited

19. Security Classif. (of this report) 20. Security Classif. (of this page)

Unclassified Unclassified

21. No. of Pages

" For sale by the NationalTechnical InformationService.Springfield,Virginia 22161

NASA-C-168 (Rev 10-75)