N 84 - 21

N ASA CR-174633

GARRETT REPORT 21-4498

NASA

FINAL REPORT

FOR THE

ADVANCED TURBOCHARGER

DESIGN STUDY PROGRAM

BY

D.G. CULY AND R.W. HELOENBRAND

GARRETT TURBINE ENGINE COMPANY

PHOENIX, ARIZONA

AND

N.R. RICHARDSON

AIRESEARCH INDUSTRIAL DIVISION

GARRETT AUTOMOTIVE PRODUCTS COMPANY

TORRANCE, CALIFORNIA

PREPARED FOR

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

NASA-LEWIS RESEARCH CENTER

(CONTRACT NAS3-22750)
t Report No

NASA CR-174633

2 Government Accession No. 3 Recipient's Catalog No

4 Title and Subtitle

ADVANCED TURBOCHARGER DESIGN

STUDY PROGRAM - FINAL REPORT

5 Report Date

1 January 1984

6 Performing Organization Code

7. Author(s)

D.G. Culy, R.w. Heldenbrand, N.R. Richardson

8 Performing Organization Report No

GTEC 21-4498

10. Work Unit No

9. Performing Organization Name and Address

Garrett Turbine Engine Company

Phoenix, Arizona

11 Contract or Grant No

NAS 3-22750

13. Type of Report and" Period Covered

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546 14 Sponsoring Agency Code

15 Supplementary Notes

16 Abstract

The Advanced Turbocharger Design Study was conducted from April 1, 1981

to April 30, 1983 as a joint effort of the Garrett Turbine Engine Co. and

the AiResearch Industrial Division of The Garrett Corporation. The program

consisted of: (1) the evaluation of three advanced engine designs to determine their turbocharging requirements, and of technologies applicable to

advanced turbocharger designs; (2) trade-off studies to define a turbocharger

conceptual design and select the engine with the most representative requirements for turbocha.rgi.ng; C3) the preparation, of a, turbochargai; conceptual .

design for the 'Curtiss Wright RC2-32 engine 'selected in 'the trade-off studies;

and (4) the assessment of market impact and the preparation of a technology

demonstration plan for the advanced turbocharger. Supporting subcontract

effort was provided by Cessna Aircraft Co. for aircraft performance and cost,

and by Avco-Lycoming for baseline reciprocating engine performance.

17. Key Words (Suggested by Author(s))

Advanced Turbocharger

Design Study

18 Distribution Statement

Unclassified - unlimited

19 Security Qassif (of this report)

Unclassified

20 Security Classif (of this page)

Unclassified

21. No of Pages 22 Price

* For sale by the National Technical Information Service, Springfield, Virginia 22151

NASA-C-168 (Rev fi-71)

FOREWORD

The Advanced Turbocharger Design Study (ATDS), Contract NAS3-

22750, was conducted by the Garrett Turbine Engine Company (GTEC)

and the AiResearch Industrial Division (AID), divisions of The

Garrett Corporation, for the National Aeronautics and Space Administration, Lewis Research Center. Subcontract support was provided

by the Cessna Aircraft Co., Pawnee Division and by AVCO-Lycoming,

Williamsport Division.

The authors would like to thank the following NASA-Lewis personnel for their specific contributions to the program: K. L.

Abdalla, T. J. Wickenheiser, and W. F. Hady. Special appreciation

is also extended to Teledyne Continental Motors and to the

Curtiss-Wright Corporation for the essential information supplied

by them that made this program possible.

111
"Page missing from available version"
TABLE OF CONTENTS

1.0 SUMMARY 1

2.0 INTRODUCTION 3

2.1 Background 3

2.2 ATDS Program Scope 3

2.3 Related Work Effort 4

2.4 Significance of the ATDS Program 4

2.5 Purpose of the ATDS Program 5

3.0 PROGRAM APPROACH 7

3.1 Requirement Definition 7

3.2 Technology Survey 8

3.2.1 Component Technology 16

3.2.2 Cycle Technology 26

3.2.3 Design Opportunities 38

3.2.4 ATDS Technology Survey Results 41

3.3 Turbocharger Trade-Off Studies 41

3.3.1 Evaluation Criteria 41

3.3.2 Turbocharger System Configurations 43

3.3.3 Exhaust Pulse Utilization 53

3.3.4 Engine Vibration 59

3.4 Turbocharger Conceptual Design 59

3.4.1 Compressor 60

3.4.2 Turbine Aerodynamic Design 64

3.4.3 Turbine Rotor Mechanical Design 67

3.4.4 Ceramic Turbine Rotor Material and

Process Analysis 70

3.4.5 Turbocharger Design Details 70

3.4.6 Bearings 73

4.0 TECHNOLOGY IMPACT 83

4.1 Key Technology Requirements 83

4.2 Overall Market Impact 85

4.2.1 General-Aviation Market 88

4.2.2 Truck Market 80

4.2.3 Marine Market • 89

4.2.4 Off-highway Market 90

4.2.5 Automotive Market 90

4.2.6 Locomotive Market 92
TABLE OF CONTENTS (Contd)

Page

4.2.7 Ground Power Market 92

4.2.8 Market Impact Summary 94

4.3 ATDS - Related Programs 94

5.0 TURBOCHARGER TECHNOLOGY - FUTURE PLANS 99

5.1 Summary 99

5.1.1 Cycle Definition 99

5.1.2 Advanced Turbochargers 100

5.1.3 Required Technology Advances 100

5.2 Recommended Follow-On Effort 104

5.2.1 Technology Requirements Coordination 105

5.2.2 Advanced Component Turbocharger Testing 105

5.2.3 Component Technologies 107

6.0 CONCLUSIONS AND RECOMMENDATIONS 113

REFERENCES 115

APPENDICES

I CESSNA-PAWNEE SENSITIVITY STUDIES

II ATDS AIRPLANE COST EQUATIONS
LIST OF ILLUSTRATIONS

Figure Title Page

1 Advanced Technology Engines 8

2 Engine Contractors Turbosystem Requirements 10

3 Engine Temperature Rise Characterization 11

4 Characterization of Engine Indicated SFC 12

5 Characterization of Engine Friction Horsepower 13

6 Characterization of Engine Fuel/Air Ratio 13

7 Standardized Basis for Relative Technical

Risk Assessments 16

8 Centrifugal Compressor Industry Capability

Available in 1986 17

9 Turbine Industry Capability 17

10 Elements of the Maximum Complexity

Microprocessor System 26

11 Heat-Exchanger Core Configurations 27

12 Heat-Exchanger Pressure Drop Characteristics 28

13 Heat-Exchanger Airflow/Effectiveness

Characteristics 28

14 Conventional Design Intercooler Core Weight 29

15 Intercooler Core Weight for Minimal Cooling

Air Face Area Design 29

16 Heat-Exchanger Sensitivities 30

17 The Hyperbar System 32

18 Bottoming Cycle Schematic 35

19 Steam Bottoming Cycle Performance 36

20 Steam Bottoming Cycle Performance 37

21 Schematic of the Turbocharger APU Mode 39

22 Performance of the Turbocharger as an APU 40

vii
LIST OF ILLUSTRATIONS (Contd.)

Figure Title Page

23 Technical Survey Results 41

24 Cessna Airplane Sensitivities to BSFC 43

25 Cessna Airplane Sensitivities to BSFC 44

26 Cessna Airplane Sensitivities to Engine Weight 44

27 Cessna Airplane Sensitivities to Engine Weight 45

28 Cessna Airplane Sensitivities to Block Fuel 45

29 Turbocharger Component Sensitivities 46

30 ATDS Tradeoff Study Matrix 46

31 Port Opening/Closing Geometries 53

32 Port Effective Areas 54

33 Turbocharger Location Tradeoff for RC2-32 Engine

(25,000 Ft, 250 HP) 55

34 Exhaust Energy Recovery for RC2-32 Engine

(25,000 Ft, 250 HP) 55

35 Exhaust Configuration Influences Energy

Recovery by Turbocharger for RC2-32

Engine (25,000 Ft, 250 HP) 56

36 Turbocharger Location Affects Available

Engine Power for RC2-32 Engine (25,000 FT,

250 HP) 57

37 Effect of Exhaust Manifold Volume for

GTSIO-420/SC (25,000 Ft, 250 HP) 58

38 Exhaust Energy Recovery for GTSIO-420/SC

(25,000 Ft, 250 HP) 58

39 Exhaust Energy Recovery (GTSIO-420/SC 25,000 FT,

250 HP Divided Exhaust System) 59

40 Compressor Flow Path is Compact 62

41 High Efficiency and Pressure Ratio

are Major Advancements 63

viii
LIST OF ILLUSTRATIONS (Contd.)

Figure Title Page

42 Turbine Stator and Rotor Flow Path 65

43 High Turbine Efficiency Dictates Long Flow Path 65

44 Advanced Turbocharger Stage Inlet and

Exit Diagrams 66

45 Required Turbine Performance is Close to

State-of-the-Art 66

46 Turbine Blade Vibration Study Shows Basic

Concept Feasibility 68

47 Blade Natural Frequencies are Greater Than the

Minimim Requirement 68

48 Turbine Rotor Blade Taper Chosen for Low Stress 69

49 ATDS Turbocharger Cross Section 71

50 Advanced Turbocharger Weighs Half of

Conventional 71

51 ATDS Conceptual Design Outline 73

52 Compressor Performance RC2-32 Engine 74

53 ATDS Foil Bearing Load Capacity 78

54 ATDS Foil Bearing Power Consumption 79

55 Details of Typical Gas Lubricated Foil Bearing 80

56 Shaft Whirl Analysis of the ATDS Rotor on

Foil Bearings 80

57 ATDS Technology will have a Broad Market Impact 94

58 Foil Bearing Technology Roadmap 101

59 Ceramic Rotor Technology Roadmap 102

60 Integration of Engine and Turbocharger

Technologies/Development Activities 104

IX
LIST OF TABLES

Table Title Page

1 Engine Master Data Chart 9

2 ATDS Characterization of Engine Airflow 10

3 Engine Airflow Characterization 11

4 NASA Aircraft Parameters 14

5 Engine Parameters at 25,000-Foot Cruise

Condition (ISA, 250 BHP) 14

6 Selected Candidates for Technology Survey 15

7 Compressor Technology Industrial Capability 18

8 Turbine Technology Industrial Capability 18

9 ATDS Compressor Efficiency Technology Evaluation

Current Technology: 70-Percent at 2.6 PR

at 50 PPM 19

10 ATDS Turbine Efficiency Technology Evaluation

Current Technology: 70-Percent at 2.6 PR,

50 PPM) 19

11 Industrial Capabilities for Advanced Rotor

Material 21

12 Lightweight Housing Materials Industry

Capabilities 21

t

13 Advanced Rotor Materials Technology Evaluation 22

14 Lightweight Housing Materials Evaluation 22

15 Industrial Capabilities for Variable-Geometry

Technology 23

16 Variable-Geometry Compressor Technology

Evaluation Current Technology: GTEC Production

ATM's and Engine 24

17 Variable-Geometry Turbine Technology

Evaluation Current Technology: GTEC Production

Axial Flow 24

x
LIST OF TABLES (Contd)

Table Title Page

18 Electronic System Evaluation for Turbocharged

Engines 25

19 ADTS Microprocessor Control Technology

Evaluation 27

20 Charge Air Cooler Technology Evaluation

Current Technology: AID Production

Nonflying and Garrett Production Flying 30

21 Two-Stage Turbocharger Technology Evaluation

Current Technology: Series Connection or Two

Turbos 31

22 Mechanical Compounding Technology Evaluation

Current Technology: Production IN R-3350 32

23 Hyperbar Technology Evaluation

Current Technology: Limited French

Production 33

24 Exhaust Pulse Utilization Technology Evaluation

Current Technology: Well-Developed 34

25 Exhaust Pulse Tuning Technology Evaluation

Current Technology: Requires Analysis for

Each Engine 34

26 Intake Pulse Tuning Technology Evaluation

Current Technology: Well-Developed. Requires

Analysis of Each Engine 35

27 Bottoming Cycle System Operating Conditions 36

28 Bottoming Cycle Technology Evaluation

Current Technology: Laboratory Demonstration

Only 37

29 Gas-Lubricated Foil Bearings Technology

Evaluation. Current Technology:

Cold: GTEC Production

Hot: Laboratory Demonstration Only 38

30 Bearing Industry Capability 39

XI
LIST OF TABLES (Contd)

Table Title Page

31 APU-Mode Turbocharger Technology Evaluation

Current Technology: Related to Well-Developed

Hyperbar and APU Technologies 40

32 NASA CR-165564 Evaluation Scheme 42

33 ATDS Benefit Ranking 42

34 Individual Component Technology Trade-Offs

for ATDS 47

35 Optimized TIO-540 SI RECIP Engine,

Current Technology 47

36 Optimized GTSIO-420/SC Highly Advanced

SCSI RECIP 48

37 Optimized GTDR-246 Highly Advanced Diesel 48

38 Optimized RC2-32 Highly Advanced SCSI Rotary 49

39 Technology Trade-Off Evaluation Points 49

40 Turbocharger Requirement Commonality 50

41 Recommended Features for RC2-32 Turbocharger

Conceptual Design 52

42 ATDS Engine Vibration Estimates 60

43 RC2-32 Turbocharger Flow Requirements

(25,000 Ft., 250 BHP) 61

44 Component Design Points for 25,000-Foot

Condition (200 to 225 KTS, 250-Engine BHP) 61

45 Compressor Aerodynamic Design 63

46 ATDS Conceptual Design Bill of Materials 72

47 RC2-32 Ratings for Both Turbocharged and

Nonturbocharged Engines 74

48 Aircraft Operational Mode and Associated Loads 75

49 Bearing Capacity 76

XII
LIST OF TABLES (Contd)

Table Title Page

50 Key Turbocharger Technology Requirements 83

51 Turbocharger Features have Broad Appeal 86

52 Suggested Areas of Emphasis 86

53 Factors Needing Further Demonstration 86

54 General-Aviation Market 88

55 Truck Market 89

56 Marine Market 90

57 Off-Highway Market 91

58 Automotive Market 92

59 Locomotive Market 93

60 Ground Power Market 93

61 Overall Market Impact Summary 95

62 Existing Technology Programs 95

63 Concepts that must be Addressed by Future Programs 97

Kill
1.0 SUMMARY

The GTEC Advanced Turbocharger Design Study (ATDS) Program,

NASA Contract NAS3-22750, was conducted to determine the potential

of advanced turbochargers as candidates for application to advanced

technology engines in light, general-aviation aircraft capable of

moderately-high altitude operation. The primary objectives of this

program were to:

o Define the benefits expected from applying advanced turbocharger concepts to advanced technology intermittent

combustion (1C) engine concepts.

o Define the turbocharger technology programs needed to

realize the most significant advancements.

o Describe the scenario necessary to convert this technology into a cost-effective turbocharger family having

numerous applications.

Studies conducted by several airframe manufacturers were used

to define the mission profiles and aircraft performance for several

engine configurations. Garrett then selected four of these concepts for analysis in the ATDS Program.

At the onset of the ATDS Program, three advanced turbocharged

reciprocating engines were modeled for a 250-bhp cruise, 25,000-

foot altitude condition. The performance of these designs was then

compared to a current baseline configuration. The four engines

were:

o TIO-540 (Baseline Engine) - Current Technology SparkIgnition Turbocharged Reciprocating Engine

o GTSIO-420/SC - Highly Advanced Spark-Ignition Stratified-Charge Turbocharged Reciprocating Engine

o RC2-32 - Highly Advanced Spark Ignition Stratified

Charge Turbocharged Rotary Combustion Engine

o GTDR - Highly Advanced Two-Stroke Diesel Turbocharged

Radial Engine.

The RC2-32 engine was ultimately selected as the reference

engine for designing the advanced turbocharger, since its overall

needs were the approximate median of the turbocharger requirements

of all of the advanced engine designs under study. The results of

the ATDS Program showed that the ATDS turbocharger concept is, indeed, a viable candidate for such an application. Moreover, many

of the technologies satisfy the needs of nonaviation applications,

as well. However, it was also determined that additional component
development programs must be instigated immediately to enable the

ATDS turbocharger concept to be committed to production. These include such programs as compressor and turbine aerodynamics, advanced turbine materials, gas-foil bearing technology, and structural advancements (i.e., advanced lightweight materials).
2.0 INTRODUCTION

This document is submitted to NASA by Garrett Turbine Engine

Company (GTEC) and AiResearch Industrial Division (AID), divisions

of The Garrett Corporation. It presents the Final Report for the

Advanced Turbocharger Design Study (ATDS) Program as requested in

Exhibit A, Task IV of Contract NAS3-22750. The program was conducted from April 1, 1981 through April 30, 1983 for the NASA-Lewis

Research Center (NASA-LeRC).

AID performed over 40 percent of the ATDS effort, and is the

world's largest producer of turbochargers for both the general- and

nonaviation marketplace. Over 60,000 turbochargers are produced

each month—400 of which are used in light, general-aviation aircraft. AID has contributed major advancements in the areas of control systems and heat exchangers, as well as turbocharger aerodynamics, all of which have expanded the use of turbochargers in

general-aviation aircraft for moderately high-altitude operation.

2.1 Background

The desire to provide a lightweight aircraft capable of highaltitude operation, coupled with the anticipated problems associated with availability, cost, and quality of aviation fuels led to

the examination of alternate powerplant configurations for this application. On-going progress in this technology that can be directly applied to turbochargers led NASA-LeRC to undertake a reexamination of the issues of turbocharger design, technology, and

application.

2.2 ATDS Program Scope

The program effort comprised the following tasks:

o Task I - To evaluate advanced engine performance

parameters to determine turbocharger requirements, and to evaluate technologies potentially applicable to advanced turbocharger designs .

o Task II - To perform trade-off studies to define technologies and features for a conceptual design, and to select the engine with the most

representative turbocharger requirements.

o TASK III - To prepare a conceptual design of an advanced

turbocharger system for the reference engine

selected in Task II.
o Task IV - To assess the overall market impact of an advanced turbocharger. To prepare a technology

demonstration plan for advanced general-aviation turbochargers.

The turbocharging requirements of three advanced reciprocating engines in the 250-bhp cruise, 25,000-foot altitude class were

used to assess the needed technologies for this application. These

were then compared to the requirements of a current production turbocharged piston engine. These four engines (selected from NASA

programs currently underway) were:

o TIO-540 (Baseline engine) - Current Technology SparkIgnition Turbocharged Reciprocating Engine

o GTSIO-420/SC - Highly Advanced Spark-Ignition, Stratified Charge, Turbocharged Reciprocating Engine

o RC2-32 - Highly Advanced Spark Ignition Stratified

Charge Turbocharged Rotary Combustion Engine

o GTDR - Highly Advanced Two-Stroke Diesel Turbocharged

Radial Engine.

2.3 Related Work Effort

Portions of the ATDS Program activity were subcontracted to

the Cessna Aircraft Co., Pawnee Division and to the AVCO-Lycoming,

Willamsport Division. Cessna support provided aircraft performance

and cost sensitivities to engine parameters which were used to

evaluate the ATDS installation. These sensitivities are listed in

Appendix I. Lycoming support provided data for the baseline

engine. Results of engine design studies prepared for NASA by

Teledyne Continental Motors and by the Curtiss-Wright Corporation

for the highly advanced spark ignition, diesel, and rotary engines

were also taken into consideration.

2.4 Significance of the ATDS Program

The ATDS Program established that advanced technology turbochargers are viable candidates for meeting the general-aviation

needs of the future. The study identified that a family of turbochargers is needed with the range of flow capacities and pressure

ratios required for the diverse requirements of both aviation and

nonaviation engines.

Technology demonstration programs were also identified that,

if initiated immediately, would lend impetus to development of advanced turbocharger designs.
The studies showed that the RC2-32 best represented the turbocharger requirements, since its pressure ratio, airflow, and system

configuration needs had many points in common with the other two

advanced engines. The trade-off studies identified a single-spool,

single-stage free-floating configuration as the simplest and most

cost-effective system. The concept features gas-lubricated bearings, a ceramic turbine rotor, a high-pressure-ratio compressor,

and lightweight structures.

In addition, market impact studies conducted as part of the

program identified the ATDS turbocharger technologies as having

many applications, besides that of general aviation; i.e., automotive, marine, and ground-power usage.

2.5 Purpose of the ATDS Program

The primary objective of the ATDS Program was to develop a new

turbocharger concept for potential use in future light aircraft

capable of high-altitude operation. A secondary objective was to

evaluate and assess the market potential for this newly evolved

turbocharger not only in the general aviation marketplace, but in

related markets, (i.e., automotive, locomotive, marine, etc.).
Page intentionally left blank

Page intentionally left blank
3.0. PROGRAM APPROACH

3.1 Requirement Definition

A review of the several NASA advanced engine and airframe

studies was conducted to select the operating parameters to be used

in the ATDS Program. Using this information, three advanced

engines and one baseline engine were selected for ATDS evaluation.

These were:

Model No. Contractor

GTSIO-420/SC

RC2-32

GTDR-246

TIO-540

Teledyne Continental Motors/

Aircraft Products Div.

(TCM/APD)

Curtiss-Wright (C-W)

Teledyne Continental Motors/

General Products Div.

(TCM/GPD)

AVCO-Lycoming (AL)

Study/Conf iguration

Highly Advanced SparkIgnition Stratified

Charge Turbocharged

Recipricating Engine

Highly Advanced Spark

Ignition StratifiedCharge Turbocharged

Rotary Engine

Highly Advanced TwoStroke Diesel

Turbocharged Radial

Engine

Current Technology

Spark-Ignition

Turbocharged

Reciprocating

(Baseline) Engine

The selected advanced engine configurations are described in

Figure 1 and Table 1. Near the end of the program, NASA requested

that final comparisons allow for a reduction of displacement of the

diesel engine from 246 cubic inches, based on expected technology

advancements to improve comparability among engine technology

levels. In comparisons made on this basis, the diesel engine is

identified as "GTDR" rather than as "GTDR-246". The turbocharger

system requirements for each engine (as identified by each engine

company) are also listed in Table 2. The wide variations in these

requirements are more apparent in Figure 2. These comparisons show

the diverse requirements of the advanced engines, and also show

that a significantly different turbocharger system design would result for each engine. For the turbocharger matching studies, each

engine was aerothermodynamically defined in terms of airflow, temperature rise, indicated specific fuel consumption (ISFC), friction

horsepower, and fuel/air ratio as functions of operating conditions. These parameters were derived from data obtained from both
ADVANCED

ROTARY

STRATIFIED

CHARGE

RC2-32

ADVANCED

SPARK

IGNITION

GTSIO-420/SC

ADVANCED

DIESEL

GTDR-246

Figure 1. Advanced Technology Engines.

NASA and the engine contractors and are shown in Figures 3 through

6, and Table 3. Single- and twin-engine aircraft data for the four

engines was supplied by NASA. Sensitivities ot aircraft performance, takeoff gross weight (TOGW), and direct operating cost (DOC)

to engine parameter changes were supplied by Cessna, and appear in

Appendix I. The overall aircraft parameters used for the study are

identified in Table 4, while the baseline engine parameters(If2,5)

are listed in Table 5. The sensitivity factors of each are

included in Appendices I and II.

3.2 Technology Survey

For the technology survey, the turbocharger system technologies were divided into three classifications: component technologies, cycle technologies, and design opportunities. A large number

of concepts were screened prior to this classification. Each survey candidate (Table 6) was then subjected to a technical risk and

manufacturing cost evaluation. This provided estimates of the

potential availability and cost of incorporation of each candidate

in any proposed turbocharger system. Relative technical risk of

each technology was defined as a function of four factors:

o Concept originality — If the concept were new, or if it

were related to an existing concept.

Functional design — If the design fulfilled the need of

the system concept.
TABLE 1. ENGINE MASTER DATA CHART.

ENGINE TYPE SPARK IGNITION DIESEL

CURRENT/BASELINE MODERATE RISK HIGH RISK HIGH-BISK

TS10-550 GTSUO-420 GTS10-420/SC • • 1993 TECH

ROTARY

ADVANCED

RC2-47

Engine Ht. (Ib)l 585 * 121 • 706 4B5 *121 • 606 405 » 121 • 526 367 * 121 • 488 348 * 139 - 487

Basic »

Additional

Scaling ..- .... a,fl

(Basic WT) «. * HP-81B M0 « HP'818 «„ » HP'810 Specific Ht • scaled H • 92»0.8

B B as shown In Enc. 2 B(TOHP)

Dimensions (In.)

L 50.62 59.25 59 25 41.30 52.0

H 34.06 33.375 33.375 24.00 18.0

B 20.02 19.25 19.25 25.00 17.0

Scaling (Inches)

L None None None None 44 8+0.0225

(TOHP)

It None None None None None None

B None None None None None

CC locations 2

Aft 23.6 32.6 33.6 15.7 22.8

Below 0.9 0.56 O.S2 ON ON

Right 0.65 0.80 0.90 ON ON

ECON.

Power (HP) 4 ALT HCP CRUISE ALT HCP CRUISE ALT MCP CRUISE ALT MCP/CRUISE CRUISE ALT KCP CRUISE

at ISA 0 350 247 0 3SO 229 0 350 228 0 to 0 to

condition 5K 359 256 5K 367 242 5K 364 243 17K 360 297 21K 320 250

10K 358 262 10K 374 251 10K 368 252

15K 348 263 15K 366 256 1SK 360 257

20K 327 259 20K 351 256 2 OK 343 255 201 120 264

25K 298 250 25K 323 250 2SK 316 250 25X 250 206 25K 275 250

301 26S 240 3 OS 290 240 30K 280 240 30K 18$ 153 30K 227 208

351 231 228 35K 252 228 35K 245 228 35K 120 99 35K 290 176

Ri>M 2800 2300 RPM 2400 2150 RPM 2400 2150 RPH 2300 2304 RPM 2400 2000

SFC 0 to 0 to 0 to 0 to I

(Lfl/HP-HR) 35K 0.600 0.446 35K 0.486 0.358 35« 0.334 0.331 17S |°-31J 0.290

2 OK 0.317 0.293

2SK 0.12) 0.299

301 0.330 0.305

351 0.33* 0.311

Scaling

Pow*r Linear Linear Linear Linear

src None None None None

Requirements

(% SUP) 100% 77.61 79 31 70.3%

1 Basic wt (Ib) scaled with power output. Additional ut not constant (unsealed)

All wts - Ib.

2 Center of gravity (CG) locations referenced to prop flange and crankshaft ( ), eicept Cor

date lateral and vertical positions referenced to prop shaft

3 Current and moderate risk spark ignition engines. Eiperience 6 povar loss of 1% per 6*P r

ambient temperature. Other 1C engines not sensitive to temp, variation.

Sate data based on density altitude.

0 to -

25K

30K 0.380 0.367

35K 0.382 0 369

Linear

None

58 5%

gate,

ise in

HIGHLY ADVANCED

RC2-32 .

255 * 138 ' 393

U. • 854* 53

"(TOHP)

48.6

16.0

16.0

42.3*0.019

(TOHP)

None

22.7

01

ON

ALT MCP CRUISE

0 to

21K 320 250

25K 275 250

30K 227 208

35K 290 176

RPM 2400 2000

0 to -

25K

30X 0.370 0.354

35K 0.378 0.351

Linear

None

61.11
TABLE 2. ATDS CHARACTERIZATION OF ENGINE AIRFLOW.

PARAMETER

Compressor

Airflow, Ib/nun

Pressure ratio

Efficiency, %

Intercooler

Effectiveness

Turbine

Gas flow, Ib/min

Inlet temperature,

°F

Inlet pressure, psia

Pressure ratio

Efficiency, %

CURRENT

TIO-540

SI RECIP

36.7

3.6

0.75

0.5

28.0

1360

15.0

2.7

0.70

ADVANCED

GTSIO-420/SC

SI RECIP

36.2

3.3

0.57

0.4

27.6

1518

18.3

3.35

0.62

ADVANCED

RC2-32

SI ROTARY

47

5.5

0.70

0.5

38.5

1029

29.2

5.4

0.74

ADVANCED

GTDR-246

DIESEL

52.0

7.3

0.82

1

0.5

53.7

1117

30.3

5.6

0.72

eo

eo 4 111 ~

3

2

1

GTDR

RC2-32

GTSIO-420/SC

20 40 60 80 100 120

CORRECTED FLOW. LB/MIN

COMPRESSOR

140 -•

120-

ec

=e

£ 100-

1U

eo

cc

£ 80-

h—

LL.

1 60-

LU

| 40-

t—

20-

0-

GTDR

* CURRENT ___

PRODUCTION

LIMIT

•

RC2-32

GTSIO-420/SC

0

•

TIO-540

1 1 1 I

1000 1200 1400 1600 1800

MAX TURBINE INLET TEMPERATURE. °F

TURBINE

Figure 2. Engine Contractors Turbosystem Requirements.
TABLE 3. ENGINE AIRFLOW CHARACTERIZATION.

ENGINE

TIO-540

GTSIO-420/SC

RC2-32

GTDR-246

CHARACTERIZATION

PTP

0.234 fif + 0.648 540 (ERpM) pJE

E

 /TIE

 26°6 TIE

V5T9

„ ,0 . PIE . 420 (ERPM) PIE

E ~ '"• PIT U.' ) 2606 TIE

W = [0.3455 + 8.257 x 10~6 (ERPM)]

 P-^E

-

** /TIE

Y M PIT . ERPM A U " PIE

 + 20871'

WE = 15.63(CA) PIE N

(CA) = 1.833 - 1.267 x 10~4 (ERPM)

N = 3.8875 , //PIEN' 1'

4 3 4

 /PIEX" 1'

7 1 7

V \PIT/ IPIT^

ERPM = Engine speed, rpm

PIE = Engine inlet pressure, in Hg A

PIT = Turbine inlet pressure, in Hg A

TIE = Engine inlet temperature, °R

CA = Effective area, sq. inch

WE = Engine Airflow, Ib/min

1BOO -

^ 1600 -

& 1400 -

= 1200 -

«r

^ 1000 -

" 800

UJ

I 600 H

LU

400

0.02

o GTSIO-420/SC

o RC2-32

A GTDR-246

D TIO-540

D— CRUISE— D

D—TAKEOFF—D

AND DESCENT

0.03 0.04 0.05 0.06

FUEL/AIR RATIO

Figure 3. Engine Temperature Rise Characterization,

11
o Operating requirements — If the levels of component

loading, efficiency, etc., were technologically realistic for the anticipated time period in operational introduction.

o Mechanical design — If the mechanical design met functional and operating requirements.

Figure 7 shows the probability of failure (risk) assigned to the

five phases of product development identified by the study.

Relative manufacturing costs were defined as a function of

three factors:

o Level of operating parameters (i.e., pressure, temperature, life, reliability, and maintainability)

o Material cost — Exotic to conventional, and strategic

versus nonstrategic.

o Fabrication difficulties — Material machinability,

castability, weldability, and adaptability to automated

assembly.

0.71

0.6

0.5 H

0.4

0.451

0.4-

06TSIO-420/SC

ORC2-32

A6TDR-246

QTIO-540

0.035 0.04 0.045 0.05 (OTHERS]

FUEL/AIR RATIO ' ' ' ' TIO-540

0-06 0.07 0.08 00g010 OIJ 0NLY

Figure 4. Characterization of Engine Indicated SFC.

12
50

o GTSIO-420/SC

ORC2-32

AGTDR-246

oTIO-540

2000 2400 2800 3200 3600 (OTHERS)

6000 7200 8400 9600 10800 (RC2-32)

ERPM

Figure 5. Characterization of Engine Friction Horsepower.

o GTSIO-420/SC

o RC2-32

AGTDR-246

DTIO-540

DESCENT CRUISE

POWER LEVEL

TAKEOFF

Figure 6. Characterization of Engine Fuel/Air Ratio.

13
TABLE 4. NASA AIRCRAFT PARAMETERS.

FIXED-ENGINE SIZE, VARIABLE AIRFRAME, FIXED PAYLOAD-RANGE*

CURRENT ADVANCED ADVANCED ADVANCED

ENGINE SI RECIP** SI RECIP ROTARY DIESEL

AIRCRAFT TYPE SINGLE TWIN SINGLE TWIN SINGLE TWIN SINGLE TWIN

Takeoff gross

weight, Ib

Empty weight,

Ib

Takeoff Power,

HP

it- Cruise speedM

kt

Mission fuel,

Ib

Acquisition

cost, ks

DOC, $/hr

4460

2736

340

206

440

202

122

6850

4428

340

228.5

855

381.5

230

3888

2340

350

220

287

186

106

5907

3802

350

240.5

581

347

198

3691

2127

320

229

296

175

103

5454

3327

320

252

592

320.5

190

3849

2310

340

218

278.5

188

107

5753

3680

340

244

555

338.5

195

*TCM/APD TSIO-550

**S ingle: Payload = 1200 Ib: Range = 700 NMI

Twin: Payload = 1200 Ib: Range = 800 NMI

***Cruise at 25,000 feet, ISA

TABLE 5. ENGINE PARAMETERS AT 25,000-FOOT CRUISE

CONDITION (ISA, 250 BHP)

PARAMETER

BSFC, lb/hr-bhp

Fuel/air ratio,

Airflow, Ib/min

Temperature rise

SI RECIP.

TIO-540

(BASELINE)

0.500

Ib/lb 0.077

26.6

, °F 1255

GTS 10-

420/SC

ADVANCED

SI RECIP.

0.331

0.0500

27.6

1600

RC2-32

ADVANCED

SI ROTARY

0.355

0.04

37.9

870

GTDR

ADVANCED

DIESEL

0.323

0.0414

52

880

14
C

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15
LU

100-

90-

70-

£ 60-

u

<

LU

50-

40-

30-

20-

10-

0-

COMPLETELY

THEORETICAL

EXPLORATORY

PHASE

ON-GOING

PROGRAMS

CURRENT

LABORATORY

CURRENT

COMMERCIAL

Figure 7. Standardized Basis for Relative Technical

Risk Assessments.

3.2.1 Component Technology

The following technologies were applicable to potential turbocharger component candidates and were subsequently evaluated in the

ATDS Program.

3.2.1.1 Compressor/Turbine Efficiency Improvements - It is obvious that turbocharger aerodynamic components can benefit from advancements made in gas-turbine engine technology. This technology

can be easily transferred to large turbochargers, since the same

manufacturing processes are used for both and can be justified by

similar cost-quantity relationships. However, to date, the pricequantity relationships in the small-turbocharger market have made

it unfeasible to incorporate the latest gas turbine aerodynamic

technology. This trend is expected to continue until the turbocharger thermodynamic work becomes a sufficiently large proportion

of the total powerplant work, which will mean that the improved

efficiency attained will offset the high cost of obtaining it.

16
Current expectations of industrial compressor and turbine capabilities (allowing for 3 years of component development) are shown in

Figures 8 and 9 and Tables 7 and 8, while the ATDS technology evaluation for these components are listed in Tables 9 and 10, respectively. The turbine efficiencies shown assume good diffusion of

the turbine exit flow to minimize dumping losses.

0.86 -i

0.84-

0.82-

E 0.80 -

£5 0.78 -

»—

GO

^ 0.76

0.74-

%£- K%X ,

V'

\&,

**»

ATDS DIESEL T-C

• GOAL

.

ATDS ROTARY T-C^ >.

 n*

GOAL ^

\

2.0 13" 4.0 5.0 6.0 7.0 8.0

PRESSURE RATIO. TOTAL-TO-TOTAL

9.0

Figure 8e Centrifugal Compressor Industry Capability

Available in 1986.

0.951

0.90-

0.85

ATDS

2 3 4

PRESSURE RATIO. TOTAL-TO-TOTAL

Figure 9. Turbine Industry Capability.

= 0.7 LB/SEC

17
TABLE 7. COMPRESSOR TECHNOLOGY INDUSTRIAL CAPABILITY.

STATUS CAPABILITY

Current commercial practice

Current laboratory state-ofthe-art

Firmly planned on-going

program

Exploratory experimental phase

Theoretical

TH08A turbocharger,

P/P =2.5 ?7= 0.74

GT601 Truck Engine,

P/P = 5.0 T?= 0.78

(machined titanium)

AGT101 Automobile Engine,

P/P = 4.3 - 5.0 »?= 0.80

(PM aluminum)

"Arbitrary" blades:

Twisted /S diffuser,

3-D shape duplicator

machined

Advanced design tools

not developed

TABLE 8. TURBINE TECHNOLOGY INDUSTRIAL CAPABILITY.

STATUS CAPABILITY

Current commercial practice

Current laboratory state-ofthe-art

Firmly planned on-going

program

Exploratory experimental phase

Theoretical

TH08A turbocharger,

P/P = 2.5 17= 0.73

Include vaned scroll,

exit diffuser H = 0.80

to 0.83

AGT101 automobile engine

using ceramic rotor

(2500°F) H = 0.85

P/P =4.0

Var i able-geometry

nozzle

3-D nozzle, clearance

control

18
TABLE 9. ATDS COMPRESSOR EFFICIENCY TECHNOLOGY EVALUATION.

CURRENT TECHNOLOGY: 70-PERCENT AT 2.6 PR AT 50 PPM

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

Experience (AID production)

Current laboratory state-of-the-art

(GTEC GT-601)

Expected results of significant

on-going or firmly planned programs (AGT-101, powdered metal

Aluminide)

In exploratory experimental phase

no major program commitment

(3-D Blades)

Theoretical (3-D viscous, LDV)

Low

Low

Moderate

High

Very high

Low

High

Low

Low

Very high

TABLE 10. ATDS TURBINE EFFICIENCY TECHNOLOGY EVALUATION.

CURRENT TECHNOLOGY: 70-PERCENT AT 2.6 PR,

50 PPM) .

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience (AID production)

Current laboratory state-of-the-art

(GTEC production)

Expected results of significant

on-going or firmly planned programs (3-D nozzles)

In exploratory experimental phase

with no major program commitment

(variable geometry, clearance

control)

Theoretical

Low

Low

Moderate

High

High

Low

Moderate

Moderate

Moderate

Unknown

19
As shown in Figure 8, state-of-the-art impeller fabrication

imposes a significant penalty in attainable efficiency if the

impeller is a cast design. In the case of a single-stage 6:1 compressor, this penalty exceeds 3 points in efficiency. The turbocharger marketplace has accepted this performance decrement in

order to maintain the lowest possible cost.

Programs are being conducted to develop a HIPped powder-metal

fabrication process for titanium impellers. Results have been encouraging, in that the material properties of sample impellers produced through this process have approached those of forged designs.

Achieving blade profile tolerances equivalent to those typical of

the machining process is a major problem. However, improvement in

this area indicates that with sufficient process development efforts, the required tolerances can be met. When that occurs, the

performance of machined blading can be achieved at significantly

reduced costs. A parallel approach to improved compressor performance can be achieved through advanced aerodynamic development techniques that would meet turbocharger performance requirements for

aircraft diesel engines. These aerodynamic improvements would be

achieved through both better flow visualization and in measurement

techniques (such as 3-D viscous modeling to improve visualization

and laser velocimetry to improve measurement).

3.2.1.2 Advanced Lightweight Materials - A large number of advanced material systems are under development for the aerospace

industry that have high strength-to-weight ratios. However, these

materials would benefit turbocharger technologies only if they can

demonstrate significant primary and secondary effects. For example, a ceramic turbine rotor would not only reduce rotating group

weight but would also:

o Allow the use of lightweight housings, since the mass and

size of the largest particle to be contained in the event

of a rotor burst is much less

o Enhance consideration of foil bearings because of low

rotating group weight.

Present industrial capabilities for advanced rotor materials are

listed in Table 11, while Table 12 lists the status of current

lightweight housing materials. Other advanced material systems

(i.e., titanium aluminides, carbon-carbon, and wire-reinforced

composites) are not as attractive as those listed. The ATDS technology evaluations for the advanced rotor materials and lightweight

housing materials are listed in Tables 13 and 14, respectively.

3.2.1.3 Variable Geometry - Variable-geometry techniques have

proven to be very effective in turbomachinery applications. However, this effectiveness is dependent on the physical size of the

flow channel. Leakage through the clearances required to facilitate rotation of the airfoil has proportionately greater adverse

20
TABLE 11. INDUSTRIAL CAPABILITIES FOR ADVANCED

ROTOR MATERIAL.

STATUS PART

Current commercial practice

Current laboratory state-ofthe-art

Firmly planned on-going

programs

Exploratory experimental

Theoretical

Cast aluminum compressor:

GMR235 turbine

Machined titanium compressor,

sintered SijN. or SiC

turbine

PM aluminum compressor

Si3N

4

 or SiC turbine

 (ACT)

Sintered reaction bonded

silicon nitride, carbon-carbon, Ti_ Al, wire-reinforced

composite

None considered

TABLE 12. LIGHTWEIGHT HOUSING MATERIALS INDUSTRY

CAPABILITIES.

STATUS PART

Current commercial practice

Current laboratory state-ofthe-art

Firmly planned on-going

program

Exploratory experimental

phase

rheoretical

Cast A356 aluminum Ni resist

Graphite/epoxy, stainless-steel

sheet

Hybrid fiber/epoxy

High modulus injectionmolded systems

High modulus polymers not

requiring fiber reinforcement

21
TABLE 13. ADVANCED ROTOR MATERIALS TECHNOLOGY EVALUATION.

ADVANCED TECHNOLOGY STATUS

Current commercial practice/

experience (AID production) ,

GMR235 And INCO 713 LC.

Current laboratory state-of-the-art

(GTEC production) , Mar-M 247, IN 100

Expected results of significant

on-going or firmly planned programs (AMMRC sintered Si-,N.,

ACT- 101)

 4

In experimental phase

with no major program commitment

(sintered RBSN)

Theoretical

RELATIVE

TECHNICAL

RISK

Low

High

Moderate

Low

-

RELATIVE

MANUFACTURING

COST

LOW

High

Moderate

Low

-

TABLE 14. LIGHTWEIGHT HOUSING MATERIALS EVALUATION.

ADVANCED TECHNOLOGY STATUS

Current commercial practice/

experience (AID production)

cast iron, cast aluminum

Current laboratory state-of-theart (GTEC production Hastelloy S, 321

Stainless Steel)

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase

with no major program commitment

Theoretical

RELATIVE

TECHNICAL

RISK

LOW

LOW

None

Planned

High

High

RELATIVE

MANUFACTURING

COST

Low

Moderate

None

Planned

Unknown

Unknown

22
effects on component efficiency as the height of the flow passage

decreases. For the study, a suitable efficiency decrement percent

was applied to the variable compressor and the turbine stators.

The present industrial capability for variable compressor and turbine geometry is shown in Table 15, while the ATDS technology evaluation of variable geometry for the compressor and turbine are presented in Tables 16 and 17, respectively.

3.2.1.4 Microprocessor Control - For complex powerplant systems

that combine turbocharging and auxiliary power unit functions, the

control of all system components would best be accomplished using a

microprocessor that integrated all functions, as shown in Figure

10. However, as the powerplant becomes less complex, the need for

integrated microprocessor controls becomes less pronounced. The

control needs for the four selected engines are shown in Table 18,

while the ATDS technology evaluation of microprocessor controls is

presented in Table 19.

TABLE 15. INDUSTRIAL CAPABILITIES FOR VARIABLE-GEOMETRY

TECHNOLOGY.

STATUS CAPABILITY

Compressor

Current commercial practices

Current laboratory state-ofthe-art

Variable IGVs control flow

and pressure ratios.

Possible applicability to

GTDR engine.

Variable diffuser (radial or

axial) controls flow with

relatively constant pressure

ratio

Turbine

Current commercial practice

Current laboratory state-ofthe-art

Theoretical

ATM80 (radial) : TSCP700

(axial)

TFE731-2 VCD: GT601 truck

Clearance control for zero

losses

23
TABLE 16. VARIABLE-GEOMETRY COMPRESSOR TECHNOLOGY EVALUATION

CURRENT TECHNOLOGY: GTEC PRODUCTION ATM'S AND ENGINE.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience (AID production)

Current laboratory state-ofthe-art (GTEC production)

Expected results of significant

on-going or firmly planned programs (USAF variable cycle)

In exploratory experimental phase,

with no major program commitment

Theoretical

None

Existing

Low

Low

None

planned

High

None

Existing

Moderate

Moderate

None

planned

Unknown

TABLE 17. VARIABLE-GEOMETRY TURBINE TECHNOLOGY EVALUATION

CURRENT TECHNOLOGY: GTEC PRODUCTION-AXIAL FLOW.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience (AID production)

Current laboratory state-ofthe-art (GTEC production)

Expected results of significant

on-going or firmly planned programs (high-turning blade)

In exploratory experimental phase,

with no major program commitment

Theoretical

None

existing

Low

Moderate

High

High

None

existing

Moderate

Moderate

Unknown

Unknown

24
0

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25
INPUTS

APU START COMMAND

TURBINE INLET TEMP

TUBBOSPEED COMPRESSOR

DISCHARGE PRESSURE

AIR BYPASS POSITION

POWER COMMAND

MASTER ENGINE SELECT OTHER ENGINE SPEED

ENGINE START COMMAND

ENGINE DISCHARGE TEMP

ENCINE SPEED

COMPR DISCHARGE TEMP

CYLS HEAT TEMP

ENGINE FUEL FLOW

FUNCTIONS

COMPRESSOR GEOMETRY ACTUATOR

TURBINE GEOMETRY ACTUATOR

AIR BYPASS ACTUATOR

FUEL CONTROL ACTUATOR

BURNER IGNITION SWITCH

TURBO STARTER SWITCH

FUNCTION

COMPRESSOR GEOMETRY ACTUATOR

TURBINE GEOMETRY ACTUATOR

AIR BYPASS ACTUATOR

ENGINE STARTER VALVE SWITCH

BURNER FUEL CONTROL ACTUATOR

ENGINE FUEL CONTROL ACTUATOR

ENGINE IGNITION SWITCH

PROPELLER PITCH CONTROL ACTUATOR

Figure 10. Elements of the Maximum Complexity

Microprocessor System.

3.2.1.5 Charge Air Coolers- Cooling the engine charge air will

allow lower inlet pressures and can reduce thermal and mechanical

stresses in the engine design. Two baseline heat-exchanger designs

were selected for comparison: a conventional approach that minimizes the length of the cooling air passage to minimize the coldside pressure drop and maximizes cooling airflow, and an alternate

approach that requires a longer charge-air pass to obtain decreased

cooling air inlet face area (Figure 11). Both configurations had

the same effectiveness and pressure-drop characteristics (Figure 12

and 13). For scaling the nominal design up or down to match each of

the four engines, a family of sizing curves was developed (Figures 14 and 15). Effectiveness versus weight characteristics were

examined for two ratios of cold-flow to hot-flow (Figure 16). As

indicated, the conventional design (with a 2:1 ratio of cold to hot

flow) was the lightest approach for any desired effectiveness or

size (airflow) and was ultimately selected for further evaluation.

The ATDS technology evaluation for charge air coolers is presented

in Table 20.

3.2.2 Cycle Technology

3.2.2.1 Two-Stage Series Turbocharger - The most cost-effective

way to achieve high pressure ratio (where demanded by the application) appeared to be by coupling two high-production, moderatepressure-ratio turbochargers in series. This approach would also

allow a broader operating range of high part-load efficiency than

would be attainable with a single stage. Table 21 is the ATDS

26
TABLE 19. ADTS MICROPROCESSOR CONTROL TECHNOLOGY EVALUATION.

ADVANCED TECHNOLOGY STATUS

Current commercial practice/

experience (AID production)

Current laboratory state-ofthe-art (GTEC production)

Expected results of significant

on-going or firmly planned programs (Garrett development)

In exploratory experimental phase,

with no major program commitment

Theoretical

8.8 IN.

S\

>*

CHARGE

AIR

OUT

f

1

COOLING AIR

IN

^

A /*

10.7 IN.

5 IN

RELATIVE

TECHNICAL

RISK

None

existing

Low

Moderate

High

High

CHARGE

AIR

OUT

j ,

/' ^

^

COO

A

\

L

l

\

 ^

<^S

ING

^

3

RELATIVE

MANUFACTURING

COST

None

existing

High

High

High

High

WH = 55 LB/MIN

EH = 0.5

wc/wH = 1.0

<^ ^/1Z IN.

CHARGE AIR

IN

MINIMUM WEIGHT CASE

CORE WT = 7.0 LB

7,7 IN. i

CHARGE AIR

IN

MINIMUM FACE AREA

CORE WT = 10.7 LB

Figure 11. Heat-Exchanger Core Configurations.

27
4J.U

inn

an

o.u

enO.U

dn

?n

I

n

.U

n a U.O

nfi

fiA

on

/

/

/

/

HO

SIC

/

/

/

T /

/

/

/

/

f

/

/

A

K

/'

/

/

f

ILO

IDE

10 20 40 60 80 100

Wx/TTBlQ

W = AIRFLOW RATE. LB/MIN

T = AVERAGE TEMPERATURE. °R

P = AVERAGE PRESSURE. PSIA

WDESIGN = ss LB/MIN

'DESIGN = 50 PERCENT

WCOLD = WHOT

FOR BOTH MINIMUM WEIGHT

AND MINIMUM FACE AREA DESIGNS

200 400

P/14.696

Figure 12. Heat-Exchanger Pressure Drop Characteristics.

ca

eo

1.2 -

1.1 -

1.0 -

0.9 -

0.8 -

Figure 13.

WC DESIGN = 82.5 LB/MIN

WH DESIGN = 55 LB/MIN

^ 'DESIGN = °-4 T0 °-6

\.

Heat-Exchanger Airflow/Effectiveness

Characteristics.

28
sUJ

16 -

15 -

14 -

13 -

12-

11 -

10 -

9 -

a -

7 -

6 -

5 -

4 -

3 -

CROSSFLOW,

ALUMINUM

WH = 45 LB/MIN

65

Figure 14. Conventional Design Intercooler Core Weight

24

23

22

21

20

19

18

17

16

15

14

13

12

11

10

9

8

7

6

5

4 -

CROSSFLOW. ALUMINUM

75

Figure 15. Intercooler Core Weight for Minimal Cooling

Air Face Area Design.

29
GO

CO

ca

CJ

x

0.60 -i

0.55 H

0.50 - WCOLD

WHOT

0 1.0

A 2.0

a 1.0

V 2.0

CONFIGURATION

WIN. COOLING FACE AREA

MIN. COOLING FACE AREA

CONVENTIONAL

CONVENTIONAL

0.10 0.15 0.20 0.25 0.30

HEAT EXCHANGER CORE WEIGHT/HOT SIDE AIRFLOW

0.35

Figure 16. Heat-Exchanger Sensitivities.

TABLE 20. CHARGE AIR COOLER TECHNOLOGY EVALUATION

CURRENT TECHNOLOGY: AID PRODUCTION

NONFLYING AND GARRETT PRODUCTION FLYING.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience

Current laboratory state-ofthe-art (AID production)

Expected results of significant

on-going or firmly planned programs (Garrett production)

In exploratory experimental phase,

with no major program commitment

Theoretical

Low

Low

None

planned

None

planned

Low

Moderate

None

planned

None

planned

30
TABLE 21. TWO-STAGE TURBOCHARGER TECHNOLOGY EVALUATION

CURRENT TECHNOLOGY: SERIES CONNECTION OF TWO

TURBOS.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience

Current laboratory state-ofthe-art

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

Theoretical

Low

Low

None

planned

None

planned

Very high

Moderate

High

None

planned

None

planned

Moderate

technology evaluation of the two-stage series configuration. The

cost, volume, and weight of the installation were strong negative

factors. In retrospect, a concentric two-spool arrangement is now

recommended as a candidate for applications requiring high pressure

ratio (i.e., diesel aircraft engines, and nonflying diesels in

mountainous terrain). This arrangement would minimize the weight

and volume penalty of the two-spool turbocharger.

3.2.2.2 Mechanical Turbocompounding - The addition of a second

turbine with the shaft mechanically connected to the engine was

considered as a means of maximizing the extraction of energy from

the engine exhaust. This approach was demonstrated on two aircraft

engines—the Curtiss-Wright R3350 production engine and an experimental version of the Allison V1710 engine,and also experimentally

demonstrated by Cummins on a truck diesel engine. Table 22 shows

the ATDS technology evaluation for the mechanical compounding approach .

3.2.2.3 Hyperbar - French manufacturers of large diesel engines

have been using the hyperbar system (Figure 17) for over 10 years.

Until this study, the scalability of the hyberbar concept to small

aircraft engines (where weight and volume are more critical) was

not considered. The ATDS evaluation of the hyberbar technology for

use in four general-aviation engines is shown in Table 23.

31
TABLE 22. MECHANICAL COMPOUNDING TECHNOLOGY EVALUATION

CURRENT TECHNOLOGY: PRODUCTION IN R-3350.

ADVANCED TECHNOLOGY STATUS

Current commercial practice/

experience

Current laboratory state-ofthe-art (Cummins demonstrator)*

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

(Adds variable geometry)

Theoretical

RELATIVE

TECHNICAL

RISK

None

existing

Low

None

planned

High

-

RELATIVE

MANUFACTURING

COST

None

existing

Moderate

None

planned

Moderate

-

*Demonstrated by Cummins for highway diesel use.

E EMERGENCY SHUTOFF

C COMPRESSOR

T TURBINE

B BURNER

BP BYPASS VALVE

CAC CHARGE AIR COOLER

Figure 17. The Hyperbar System.

32
TABLE 23. HYPERBAR TECHNOLOGY EVALUATION

CURRENT TECHNOLOGY: LIMITED FRENCH

PRODUCTION.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience

Current laboratory state-ofthe-art

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

Theoretical

None

Existing

(in U.S.)

Low

None

Planned

None

Existing

(in U.S.)

High

None

Planned

3.2.2.4 Intake Tuning, Exhaust Tuning, and Exhaust Pulse Utilization - Benefits obtained from intake and exhaust effects depend

on the specific engine and turbocharger combination used. Intake

tuning improves volumetric efficiency of four-stroke-cycle engines, while exhaust tuning improves scavenging of two-stroke-cycle

engines. Effects of exhaust pulse utilization by the turbine must

be estimated for the particular speed, torque, exhaust configuration, and scavenging requirements of each engine. Preliminary

technology evaluations are shown in Tables 24, 25, and 26. More

detailed studies of the exhaust pulse utilization in the advanced

technology spark ignition engines were performed later in the program and are discussed in Paragraph 3.3.3.

3.2.2.5 Bottoming Cycles - Vapor-generating bottoming cycle techniques have been evaluated for large stationary and marine powerplants and highway truck engines, but not for aircraft applications. For this study, a system was defined as shown in Figure 18

and Table 27, with the evaluation for two conditions of boiler saturation pressure and condensing temperatures (Figures 19 and 20) .

This data was then used to make the technology evaluation shown in

Table 28.

33
TABLE 24. EXHAUST PULSE UTILIZATION TECHNOLOGY EVALUATION.

CURRENT TECHNOLOGY: WELL-DEVELOPED.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience

Current laboratory state-ofthe-art

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

Theoretical

Low

Low

Low

Moderate

TABLE 25. EXHAUST PULSE TUNING TECHNOLOGY EVALUATION.

CURRENT TECHNOLOGY: REQUIRES ANALYSIS FOR

EACH ENGINE.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience

Current laboratory state-ofthe-art

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

Theoretical

Low

Moderate

Low

Low

34
TABLE 26. INTAKE PULSE TUNING TECHNOLOGY EVALUATION.

CURRENT TECHNOLOGY: WELL-DEVELOPED. REQUIRES

ANALYSIS OF EACH ENGINE.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience

Current laboratory state-ofthe-art

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

Theoretical

Low

Moderate

Low

Low

o j

JO

TUBBOCHARGER

C COMPRESSOR

1C INTERCOOLER

TC TURBINE

WG WASTEGATE

BOTTOMING CYCLE

VG VAPOR GENERATOR

TB TURBINE

6B GEARBOX

FP FEED PUMP

RG REGENERATOR

CO CONDENSER

Figure 18. Bottoming Cycle Schematic.

35
TABLE 27. BOTTOMING CYCLE SYSTEM OPERATING CONDITIONS.

STATION

0 Ambient

1 Compressor

inlet

2 Compressor

exit

3 Intercooler

exit

4 Engine exhaust

4.1 Tc turbine

inlet

bypass

5 Tc turbine

exit

bypass

6 Vapor generator

inlet

FLOW RATE

(LB/MIN)

51.98

51.98

51.98

53.70

48.27

5.43

48.27

5.43

53.7

PRESSURE

(PSIA)

5.461

5.461

39.58

38.79

30.48

30.33

6.076

6.064

TEMPERATURE

(°R)

429.6

429.6

826.83

555.8

1577.4

1577.4

1577.4

1191.33

1577.4

1230

07)

0.823

0.718

s

- 11.5

Ej5' 11

eo co

_,= 10.5

«*: u

t— LU ,-

00- 10

1—09

0.

* 40

35-

30-

BOILER GAS INLET TEMPERATURE (T6) = 770° F

BOILER SATURATION PRESS.. PSIA = 120

VAPOR TURBINE INLET TEMP (T8) = 650° F

CONDENSING TEMP. (T10) = 60° F

VAPOR TURBINE EFFICIENCY = 0.70

250 300 350 400 450

BOILER EXHAUST TEMPERATURE (T7). °F

Figure 19. Steam Bottoming Cycle Performance,

36
U

5LU ^£

O. OD c

co _i DJ

BOILER SATURATION PRESSURE = 250 PSIA

VAPOR TURBINE INLET TEMP. = 650° F

CONDENSING TEMPERATURE = 100° F

VAPOR TURBINE EFFICIENCY = 0.70

u 100-

LU

CO

tj

LU

U

80

60

40

20 40

 600 800 1000 1200 1400 1600

BOILER INLET AIR TEMP. °F

Figure 20. Steam Bottoming Cycle Performance

TABLE 28. BOTTOMING CYCLE TECHNOLOGY EVALUATION.

CURRENT TECHNOLOGY: LABORATORY

DEMONSTRATION ONLY.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience

Current laboratory state-ofthe-art

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

Theoretical

None

existing

High

Substantial

None

existing

High

High

37
3.2.3 Design Opportunities

3.2.3.1 Improved Bearings - Experience has shown that currently

used journal bearings require a consistent supply of oil that must

be cleaner than that required by the engine. High-DN rollingelement bearings—although well-developed for gas turbine engines—

are too expensive for application in a small high-speed turbocharger and are also dependent on the integrity of the oil supply.

Since installation considerations favor a turbocharger that is independent of the engine oil supply, study emphasis was placed on

bearing systems not requiring engine oil. Gas-lubricated bearings

were an obvious candidate for this application. The results of the

ATDS gas-lubricated foil bearing technology evaluation are shown in

Table 29, and were based on the considerations listed in Table 30.

Gas-lubricated foil bearings were believed to offer significant

benefits; however, not all of the penalties were apparent at this

phase of the study. Therefore, a detailed design analysis was specified as part of Task III, and is discussed in Paragraph 3.4.5 of

this report.

TABLE 29. GAS-LUBRICATED FOIL BEARINGS TECHNOLOGY

EVALUATION. CURRENT TECHNOLOGY:

COLD: GTEC PRODUCTION

HOT: LABORATORY DEMONSTRATION ONLY

ADVANCED TECHNOLOGY STATUS

Current commercial practice/

experience

Current laboratory state-ofthe-art

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

Theoretical

RELATIVE

TECHNICAL

RISK

High

(for turbo)

High

Moderate

Moderate

-

RELATIVE

MANUFACTURING

COST

Low

Low

Low

Low

-

Note: High-temperature and dynamic stability capability

needed.

38
TABLE 30. BEARING INDUSTRY CAPABILITY.

Current commercial practice

Current laboratory state-ofthe-art

Coated foil bearings (1200°F)

Foil

Benefit

Penalty

Bronze or aluminum journal

Bearings

High DN antifriction

Bearings (not considered

feasible)

Journal - Raman

silica-chromia-alumina (SCA)

Low cost

Sputtered titanium carbide

Eliminates oil system:

Improves installation

High start-up friction

3.2.3.2 Turbocharger Auxiliary Power Unit (APU) Application - The

semi-independent turbocharger configuration proposed for the GTDR246 engine led to consideration of a system combining the features

of the hyperbar approach and an accessory drive gearbox, as shown

in Figure 21. Performance of this APU system was estimated and

OIL

C COMPRESSOR

VD VARIABLE OIFFUSER

BV BLEED VALVE

BP BYPASS VALVE

SIV STARTING INTAKE VALVE

LP LUBRICATION PUMP

A6B ACCESSORY GEARBOX

FC FUEL CONTROL

SG STARTER-GENERATOR

HP HYDRAULIC PUMP

AC AIR CONO. COMPRESSOR

T TURBINE

VN VARIABLE NOZZLE

B BURNER

SO SHUTOFF VALVE

WG WASTEGATE

Figure 21. Schematic of the Turbocharger APU Mode.

39
benefits assessed. The preliminary estimates of the performance of

a current production turbocharger adapted for use as an APU are

shown in Figure 22. The overall technology evaluation is listed in

Table 31.

£

UJ

eo

CO

CO

27 -

1

AV6 BLEED AIR TEMPERATURE = 6I6"R

25 -i

20

15 -

10 -

5

0 ... .

0 5 10 15 20

BLEED AIRFLOW (WB|. LB/MIN

TH08A TURBOCHARGER/APU

5000 FEET. COLD DAY

PAMB = 1223 PSIA

TAMB = IS°F

N = 60160 RPM . OR

Figure 22. Performance of the Turbocharger as an APU.

TABLE 31. APU-MODE TURBOCHARGER TECHNOLOGY EVALUATION.

CURRENT TECHNOLOGY: RELATED TO WELL-DEVELOPED

HYPERBAR AND APU TECHNOLOGIES.

ADVANCED TECHNOLOGY STATUS

RELATIVE

TECHNICAL

RISK

RELATIVE

MANUFACTURING

COST

Current commercial practice/

experience

Current laboratory state-ofthe-art

Expected results of significant

on-going or firmly planned programs

In exploratory experimental phase,

with no major program commitment

Theoretical

None

Existing

None

Existing

High High

40
3.2.4 ATDS Technology Survey Results - Figure 23 summarizes the

overall results of the ATDS technology survey. These results were

then applied to the trade-off analysis by using these technologies

to optimize turbocharger systems for each of the four engines.

3.3 Turbocharqer Trade-off Studies

3.3.1 Evaluation Criteria

For the ATDS trade-off studies, the contract directed the use

of a benefit/penalty assessment scheme setup as performed by Cessna

in a prior study(6) (Table 32), but adapted to the needs of the advanced turbocharger design study, as shown in Table 33. Only those

factors which would be significant to the pilot or operator were

considered. Fuel usage was considered an evaluation factor, in addition to DOC even though it contributes to DOC, since fuel conservation is a separate, important element of national policy. Trip

time was also an evaluation factor, but was determined not to be

significantly affected by variations in turbocharger system design.

The evaluation points assigned to fuel consumption and operating

cost maintain the same relative weighting of cost and fuel as the

evaluation system used in Table 32.

RELATIVE TECHNICAL RISK. PERCENT

1. CURRENT COMMERCIAL

PRACTICE/EXPERIENCE

b. CURRENT LABORATORY

8TME-CF THE4RT

C. ONGOING OR FIRMLY

NAMED PROGRAMS

d EXPLORATORY

EXPERIMENTAL PHASE

«. COMPLETELY THEORETICAL

0

0

30

BO

go

0

ID

20

-

SO

0

«

IS

30

-

0

»

n

K

»

-

20

-

-

SO

0

0

—

-

BO

-

0

n

-

—

0

0

-

-

10

-

0

-

-

—

in

/

/

/

/

-

—

-

IS

IS

—

—

-

-

40

IELATIVE MANUFACTURING COST CURRENT EXPERIENCE • 14

1 CURRENT COMMERCIAL

PRACTICE/EXPERIENCE

b. CURRENT LAIOMTOIY

8TATE-OF-THE4RT

C. ONGOING OR FIRMLY

PLAINED PROGRAMS

t. EXPLORATORY

EXPERIMENTAL PHASE

1. COMPLETELY THEORETICAL

to

204

44

to

•

14

12

\1

—

to

ID

it

24

14

-

1J>

34

14

U

<U

-

14

-

-

>4

14

II

-

—

44

—

14

4

—

—

14

14

-

-

14

-

14

-

-

-

Ml

/

/

/

/

-

-

-

14

U

-

-

-

-

14

(21

/

/

/

/

0

0

-

-

-

0

1

-

-

-

0

3

-

-

-

—

5

-

-

-

131

/

/

/

/

0

SO

-

-

-

Ml

/

/

/

/

-

20

-

-

—

m

i

i

i

i

14

tl

-

-

-

14

II

-

-

-

14

II

-

-

-

—

14

-

-

-

PI

/

/

/

/

14

a

-

-

-

Ml

/

/

/

/

-

14

-

-

—

(1) INHERENT IN CONSIDERATION OP HATCHING TURBOCHAROER TO ENGINE

(2) HATCHING IS THE CONSIDERATION OF THE NONOPTIMUH HATCH OF THE TURBOCHARQER

TO THE ENGINE IN RETURN FOR GREATER COMMONALITY OF INVENTORY PARTS

(3) HUST BE CONSIDERED AS A TRADE-OFF VERSUS PERFORMANCE

(4) NO PERCEIVED ADVANTAGE TO DRIVE ACCESSORIES WITH TURBOCHARQER

Figure 23. Technical Survey Results,

41
TABLE 32. NASA CR-165564 EVALUATION SCHEME.

AIRCRAFT

CHARACTERISTIC EVALUATION CRITERIA

WEIGHTING

FACTOR

Fuel usage

DOC

Acquisition

cost

Multifuel

capability

Flyover

noise

Installation

factor

10 points for 25-percent less fuel

than baseline

10 points for 25-percent lower

10 points for 25-percent lower

purchase price

0 points AVGAS only

1 point jet fuel only

2 points (both)

+1 quieter than baseline

0 same as baseline (±2 dBA)

0 equivalent to baseline

1 somewhat better than baseline

2 much better than baseline

10

8

6

10

10

TABLE 33. ATDS BENEFIT RANKING

AIRCRAFT

CHARACTERISTIC CRITERIA

EVALUATION

POINTS

Total

operating cost

Trip time

Fuel use

Each 1 percent less than baseline

1 percent less than baseline

1 percent less than baseline

+10

+ 3

+ 7

42
As stated, Cessna supplied GTEC with aircraft performance sensitivities to changes in engine characteristics. Selected sensitivities are shown in Figures 24 through 28. The complete Cessna

package is included Appendices I, II, and III of this report. The

sensitivity of engine BSFC to turbocharger component efficiencies

was estimated based on the effects of engine pressure difference on

mean effective pressure for 4-stroke-cycle engines, and on engine

flow for the 2-stroke-cycle engine. Resulting sensitivities are

shown in Figure 29.

3.3.2 Turbocharger System Configurations - The overall trade-off

matrix involved all combinations of engines, turbocharger configurations, and technologies (Figure 30). Preliminary screening of

applicable technologies eliminated some component and cycle technologies from further consideration, as indicated in Table 34. The

remaining technologies and design opportunities were considered

elements of alternate turbosystems for each advanced technology

engine. These alternate systems were then compared on the basis of

fuel consumption, purchase, and maintenance cost for each engine.

These benefits were evaluated according to the benefit ranking system. Results of each evaluation are summarized in Tables 35

through 38. The system on each engine with the highest score for

that engine was selected for comparison of turbocharger system requirements among engines. Features not inherent to the alternate

300

CONSTANT PAYLOAD RANGE =

1ZOO LB. 700 NM 1000 1

900 -

800 -

CONSTANT PAYLOAO RANGE

1400 LB, 800 NM

u

s

0.38 0.40 044

SPECIFIC FUEL CONSUMPTION

SINGLE-ENGINE

048 0.36 040 044

SPECIFIC FUEL CONSUMPTION

TWIN-ENGINE

048

Figure 24. Cessna Airplane Sensitivities to BSFC.

43
60 70 80 90 100 110 120 130

DIRECT OPERATING COST. S/HR

SINGLE-ENGINE

0 100 200 300

DIRECT OPERATING COST. S/HR

TWIN-ENGINE

Figure 25. Cessna Airplane Sensitivities to BSFC.

:*:

CO

CONSTANT PAYLOAD RANGE

1200 LB (700 NM)

CONSTANT PAYLOAD RANGE

1400 LB (BOO NM)

-100-80-60-40-2 0 0 20 40 60 80 100

CHANGE IN ENGINE WEIGHT, LB

 BS

SINGLE-ENGINE

BASELINE ENGINE WEIGHT = 700 LB

100 -80 -60 -40 -20 0 20 40 60 BO 100

CHANGE IN ENGINE WEIGHT. LB

TWIN-ENGINE

Figure 26. Cessna Airplane Sensitivities to Engine Weight

44
CONSTANT PAYLOAD RANGE

1200 LB (700 NM)

CONSTANT PAYLOAD RANGE

1400 LB (BOO NM)

ta

o

CJ

CD

•100 -80 -60 -40 -20 0 20 40

CHANGE IN ENGINE WEIGHT. LB

SINGLE-ENGINE

80 100

BASELINE ENGINE WEIGHT = 700 LB

-100 -80 -60 -40 -20 0 20 40 60 80 100

CHANGE IN ENGINE WEIGHT. LB

TWIN-ENGINE

Figure 27. Cessna Airplane Sensitivities to Engine Weight.

500

480

460

440

420

400

380

380

340

320

300

u

s

1 70 80 90 100 110 120 130

DIRECT OPERATING COST. S/HR

SINGLE-ENGINE

1000 -,

900-

800-

700-

600-

500

0 100 200 300

DIRECT OPERATING COST. S/HR

TWIN-ENGINE

Figure 28. Cessna Airplane Sensitivities to Block Fuel,

45
•2.0-

•1.5-

•1.0-

•05-

a" -2.0 -i

CO

oa

< -1.5-

•1J)-

•0.5-

0

TI0440

6TSIO-420/SC

DM OD6

Ar,

6TOR-Z46

0 *OJ12 OM 0.08

Figure 29. Turbocharger Component Sensitivities

Figure 30. ATDS Tradeoff Study Matrix.

46
TABLE 34. INDIVIDUAL COMPONENT TECHNOLOGY TRADE-OFFS FOR ATDS.

SCREENED

ir'OK

INDIVIDUAL

ENGINES

SCKL'ENED

AS A

TECHNOLOGY KEftSON

COMPONENT TECHNOLOGY

Compressor efficiency

Turbine efficiency

Advanced rotor materials

Variable compressor geometry

Variable turbine geometry

Microprocessor control

Inter cooler/after cooler

X

X

X

X

X

X

X

(2)

(2)

(2)

CYCLE TECHNOLOGY

Two-stage turbocharging

Pressure turbocompounding

Mechanical turbocompounding

Hyper bar system

Turbocharger matching

Exhaust pulse utilization

Exhaust pulse tuning

Intake pulse tuning

Bottoming cycles

Compact designs

Improved bearings

Turbocharger-driven

accessories

Turbocharger in APU mode

Lightweight housing materials

X

X

X

X

X

X

X

X

X

X

(1)

Evaluated for elimination,

but not for inclusion.

(1)

Beneficial if no pulse

utilization.

(2)

Excessive weight penalty.

DESIGN OPPORTUNITIES

—

X

X

X

X

X

(2)

(2)

^'Consideration is implicit in the turbocharger match.

^ Beneficial to all engine systems; only consideration is cost of improvement.

TABLE 35. OPTIMIZED TIO-540 SI RECIP ENGINE, CURRENT TECHNOLOGY.

TURBOCHARGER

REQUIREMENTS

Compressor P/P

Turbine P/P

Benefit relative

to baseline

Penalty relative

to baseline

BASELINE

SYSTEM

Wastegate

3.5

3.0

TRADEOFF

SYSTEM

Var i able-geomet r y

turbine

3.2

2.5

-1% BSFC

+1% cost

EVALUATION

POINTS

+7

-10

^Evaluation -3

47
TABLE 36. OPTIMIZED GTSIO-420/SC HIGHLY ADVANCED SCSI RECIP.

TORBOCHARGER

REQUIREMENTS

Components

Compressor P/P

Turbine P/P

Power turbine P/P

Benefit relative

to baseline

Penalty relative

to baseline

BASELINE

SYSTEM

Turbocompound ,

wastegate

Power turbine

and wastegate

3.0

2.0

2.5

TRADEOFF

SYSTEM

Noncom pounded

var lable-geometry

turbine

Variable turbine

nozzle

2.4

1.7

-

No separate power

turbine and

transmission

(weight and size)

Var lable-geometry

turbine housing

7.4%-cruise BSFC

2 Evaluation

EVAL

POINTS

+26

-12

-52

-38

TRADE-OFF

SYSTEM

Noncom pounded >

wastegated

turbine

Wastegate

2.7

2.3

-

No separate

power turbine

and transmission

8.4%-cruise BSFC

EVAL

POINTS

+26

-59

-33

TABLE 37. OPTIMIZED GTDR-246 HIGHLY ADVANCED DIESEL.

TURBOCHARGER

REQUIREMENT

COMPONENTS

Compressor P/I

Turbine P/P

Benefit

relative to

baseline

Penalty

relative to

baseline

BASELINE

SYSTEM

Single-spool

turbo,

af tercooled;

variable

geometry

One turbo

with turbodriven

accessories

7.2

5.6

TRADEOFF

SYSTEM

Two- spool

Turbo, af tercooled

Two turbos in

series

2.8 ea

2.5 ea

Aerodynamic commonality commercial

market

Height, size, and

extra ducts (cost)

Extra aftercooler

Or high- tern p. /TwoStage Compressor

2 Evaluation

EVALUATION

POINTS

+ 27

-43

-16

TRADEOFF

SYSTEM

Single- spool

turbo,

Not aftercooled

One turbo

f\

-3.0% BSFC,

no aftercooler

EVALUATION

POINTS

+21

+5

+26

Note: All systems may require intercooler to limit engine intake temperature.

48
turbosystems were evaluated for their contribution to system performance. Table 39 shows evaluation point assignments for each of

these features. Results of these comparisons are shown in Table

40, in which similarities to different engine requirements are outlined.

TABLE 38. OPTIMIZED RC2-32 HIGHLY ADVANCED SCSI ROTARY.

TURBOCHARGER

REQUIREMENTS

Components

Compressor P/P

Turbine P/P

Benefit relative

to baseline

Penalty relative

to baseline

BASELINE

SYSTEM

Free-floating

af tercooled

One turbo

5.5

5.4

TRADEOFF

SYSTEM

Two- spool

af tercooled

Two turbos

in series

2.5 ea

2.5 ea

Commonality with

commercial market

Weight, size and

extra ducts (cost)

extra aftercooler

high-temp. , two

stage compressor

Z Evaluation

EVAL

POINTS

+ 14

>r

-16

-2

TRADE-OFF

SYSTEM

Free-floating

nonaf tercooled

One turbo

no aftercooler

5.7

5. 1

-3.8% BSFC and

no aftercooler

EVAL

POINTS

1

+27

+ 5

+32

NOTE: All systems may require wastegate or variable-geometry turbine to limit

firing pressure. Intercooler may be required to limit engine intake temperature.

TABLE 39. TECHNOLOGY TRADE-OFF EVALUATION POINTS.

TURBOCHARGER

REQUIREMENT

Inter cooler/after cooler

Wastegate

Variable-geometry turbine

Fixed-vane turbine inlet

Vaned compressor diffuse r

Turbine exhaust diffuse r

Turbine pulse utilization*

Intake tuning

Intercooler improvement

Microprocessor control

Lightweight rotor/housing

Lightweight compressor housing

'Questionable with wastegate

CURRENT

SI RECIP

TI10-540

BSFC COST

Required

Required

--

3 -3

2 -3

2 -1

3 0

-1 -1

18 -6

0 -7

0 -4

0 -2

SI RECIP

GTSIO-420/SC

BSFC COST

Required

Required

—

5 -4

5 -3

3 -1

5 0

0 -1

0 0

0 -8

0 -5

0 -2

ADVANCED

SI ROTARY

RC2-32

BSFC COST

Not Used

Not Used

—

Required

Requ i red

Required

13 0

2 -2

Not Used

0 -8

0 -5

0 -2

DIESEL

GTDR-246

BSFC COST

Not Used

Not Used

-24

Not Used

Required

Required

9 0

-1

Not Used

Required

0 -5

0 -2

49
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50
The alternate system evaluated for the TIO-540 spark ignition

engine was the same as the baseline engine (i.e., a current production turbocharger system, except for the substitution of a

variable-geometry turbine nozzle for the wastegate). A charge air

cooler is desirable to maintain power during climb to altitude, but

is not essential in obtaining cruise power at altitude. For current systems, the turbocharger is generally matched so that the

wastegate is nearly closed at both high-altitude and economy-cruise

power. Using a variable-geometry turbine correctly matched at this

point should produce little improvement in cruise fuel economy.

The cost of incorporating a variable-geometry design will depend on

the degree of its incorporation in other high-production applications. In turn, this acceptance depends on achievement of a lowcost, effective means of clearance control around the movable

vanes.

Two alternate systems were evaluated for the GTSIO-420/SC

engine. The baseline system selected by the engine designers

included a wastegate and mechanical turbocompounding. To quantify

the benefits of these components, one of the systems incorporated a

conventional, wastegated, noncompounded system and the other was

defined with a noncompounded, variable-geometry turbine nozzle.

Benefits and penalties of the variable-geometry turbine parallel

those established for the TIO-540 engine configuration. The compound turbine significantly improved cruise BSFC; moreover, the

advantages in fuel economy and operating cost were not outweighed

by weight and system cost increases.

The RC2-32 engine was designed without an exhaust bypass or

variable-geometry turbine, but included an aftercooler. The alternate systems evaluated for the RC2-32 engine were a series turbocharger system (meeting the high-pressure requirements associated

with commonly available turbochargers) and a nonaftercooled,

single-stage turbocharger system. The lower cost realized with the

series system were outweighed by the added complexity of the installation, including either an interstage cooler or a special

high-temperature compressor wheel for the high-pressure stage.

Though eliminating the aftercooler decreases both weight and cost

and increases thermal efficiency, it could prove desirable in reducing thermal and mechanical stresses.

The system evaluation for the GTDR-246 engine followed that of

the the RC2-32 engine, due to the similar pressure ratios and

fuel/air ratios of each. Inclusion of variable-geometry turbine

nozzle or a wastegate is necessary to limit the back-pressure that

would impair scavenging efficiency.

The RC2-32 engine was selected as the reference engine for

which the conceptual turbocharger design of Task III was to be performed, limiting the effort required to a single turbocharger

design, since the requisite turbocharger parameters were most

representative of those for all of the advanced engines.

51
Estimated manufacturing costs were based on the relatively

small manufacturing quantities of a turbocharger specially designed

for an aircraft. This initially led to the selection of conventional turbine and housing materials and the use of oil-lubricated

journal bearings. However, subsequent direction by NASA reduced

the relative importance of cost in the selection of technologies

for the conceptual design. This allowed the incorporation of airlubricated bearings, lightweight turbine materials and housings in

the design. The resulting recommended features incorporated in the

conceptual design are shown in Table 41.

Variable turbine nozzle vanes were eliminated before conceptual design features were established because of relatively high

manufacturing costs. This technology was considered again during a

later program review. Tables 35 through 38 showed that variable

geometry offers a significant performance improvement when compared

to wastegate control and is an important part of the baseline system for the diesel engine. Reduction of the relative importance of

cost makes the performance improvements the dominant consideration.

While variable nozzle vanes have been demonstrated in laboratory

turbochargers, durability and leakage through clearances around the

vanes have hindered incorporation in high-performance production

turbochargers. These problems are expected to yield to determined

development efforts. However, determining the optimum nozzle control scheme for the best engine performance and fuel economy will

constitute a significant advancement of technology.

TABLE 41. RECOMMENDED FEATURES FOR RC2-32 TURBOCHARGER

CONCEPTUAL DESIGN.

o Single-spool

o Vaned compressor diftuser

o Fixed vane turbine nozzle

o Turbine exhaust diffuser

o No intercooler

o No wastegate

o Air bearings

o Ceramic turbine rotor

52
3.3.3 Exhaust Pulse Utilization

3.3.3.1 RC2-32 Engine

NASA requested further study of exhaust pulse utilization for

the RC2-32 engine. To accomplish this, a University of Manchester

Institute of Technology (UMIST) Mark 12 Engine Simulation Computer

Program was used. This computer program simulates pressure waves

in manifolds, without requiring details of the combustion process.

The RC2-32 rotary engine was modeled as a 6-cylinder reciprocating

engine with short ducts between gas exchange ports and the manifolds. This model was the best available simulation of exhaust

ports that serve three combustion chambers. Crankshaft angles were

reconciled to the UMIST Program by considering an imaginary output

shaft turning at two-thirds actual crankshaft speed. Port areas

were estimated from scaled plots of rotor position in the trochoid,

including the effects of assumed pocket geometries in the rotor

faces (Figures 31 and 32). The four engine configurations were:

o Two turbochargers mounted close to the engine, each receiving exhaust gasses from one rotor

o One turbocharger mounted close beside the engine, receiving exhaust gasses from both rotors

INJECTOR1 , i SPARK

INLET

HAUST

CRANK ANGLE 470 CRANK ANGLE 510

CRANK ANGLE 575 CRANK ANGLE 590

Figure 31. Port Opening/Closing Geometries

53
0.008

0.008

0.004-

0002

GEOMETRIC

PORT AREA

' i ESTIMATED AREA

y= AS RESTRICTED BY

| | ROTOR/HOUSING

' I

U— OVERLAP

'.I

°1

200

A

300 400 500

EXHAUST PORT

600

 V 400 500 600 700

INTAKE PORT

800 901

Figure 32. Port Effective Areas.

o Two turbochargers mounted to the rear of the engine, each

receiving the exhaust gasses from one rotor

o One turbocharger mounted to the rear of the engine and

receiving the exhaust gasses from both rotors.

The effects of using one turbocharger per rotor or of placing

the turbines close to the exhaust ports minimized exhaust manifold

volume and increased pulse amplitudes in the exhaust stream. Most

of the exhaust flow occurred at turbine pressure ratios greater

than the time average (Figure 33) . This, caused an increase in the

power that could be extracted by a theoretical adiabatic expansion

of flow from exhaust manifold conditions. However, the wide variation in turbine inlet pressure during one exhaust pulse caused a

poor match between the gas velocity entering the turbine wheel and

the wheel speed, that reduced instantaneous turbine efficiency at

extremes of inlet pressure (Figure 34). Therefore, power extracted

by a real turbine with larger pulse amplitudes may be increased by

a smaller proportion (relative to the performance of a constantpressure system) than the power extracted by an ideal turbine.

Energy recovery integrated over the engine cycle is compared for

the four possible turbocharger mounting arrangements in Figure 35.

To properly evaluate tne benefits of increased utilization of

pulse energy in exhaust, various offsetting effects were considered. Each difference in energy extraction by the turbine was used

to estimate a difference in average back pressure on the engine in

54
ONE TURBOCHARGER

REAR MOUNTING

o

UJ

60

1.2 -i

0.8 -

0.4

0

CO

8 -

4 -

0

EXHAUST

VALVES OPEN

• i • i

t t

CRANK ANGLE -

TWO TURBOCHARGERS

CLOSE MOUNTING

EXHAUST

VALVE OPENS

CRANK ANGLE -

Figure 33. Turbocharger Location Tradeoff for RC2-32 Engine

(25,000 Ft, 250 HP).

ONE TURBOCHARGER TWO TURBOCHARGERS

REAR MOUNTING CLOSE MOUNTING

I M 80 n

|ld

 70"

1

2

 i 60 -

= E 50

EXHAUST

VALVES OPEN

EXHAUST

VALVE OPENS

3001

 f

y|| 200-

ill 100- AA

CRANK ANGLE - CRANK ANGLE -

Figure 34. Exhaust Energy Recovery for RC2-32 Engine

(25,000 Ft, 250 HP).

55
200-

180-

160-

2 140-

iu

1 120-

iU

I 100-

LU

| 80-

2 60-

40-

20-

0-

(WfWfW*

' 1

'

, -1

•

>v

s

1 ADIABATIC

* EXPANSION

•^

*r

» PREDICTED

RECOVERY

SIDE REAR SIDE REAR

112 2

TURBOCHARGER LOCATION

NO. OF TURBOCHARBERS

Figure 35. Exhaust Configuration Influences Energy

Recovery by Turbocharger for RC2-32

Engine (25,000 Ft, 250 HP).

cases where the turbine was resized to extract just enough energy

to drive the compressor. This estimate was accomplished parametrically to avoid excessive use of computer time in iterating conditions and turbine parameters to obtain an exact match. The torque

exerted on the rotor was also integrated to compare differences due

to pressure fluctuations during the gas exchange process. Finally,

the effects of differences in power required to overcome drag of an

appropriate nacelle were estimated for each configuration. These

effects are shown in Figure 36, where the effect for each engine is

measured as the amount of improvement from an arbitrary baseline

for that effect.

Cylinder pressure simulations for different configurations of

exhaust system and turbocharger showed noticeable engine performance differences during the gas-exchange period. The pressure plot

for the single, rear-mounted turbocharger system showed a pressure

wave in the engine inlet port and combustion chamber that was timed

to produce significant torque on the rotor. This intake tuning effect overcame the exhaust system performance advantage of individual, side-mounted turbochargers. Since pulse timing is a critical

factor, further studies should be conducted to provide insight into

the best use of this effect on the engine.

56
^

Urf

UJ

?

UJ

U

E

i^

UJ

&.

i

2 HP

1

TUG

NO

"

L

•

BO I

OF

.OCATION

TURBOS

SIDE

1

X

/

/

/ X

X

X

_ .. — ""

X

X

X

\

BEST

•^^^^•M

REAR

1

"^

^»

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/

/

/

/

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•M^HBM

SIDE

2

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\

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k

 \ \ \ \ \

^ \ \ \ \ N

\ \\ \\ \ \ \ \

\

>

^ —

^ —

m —

X ^^TM

REAR

2

CHARGING

'EFFICIENCY

DRAG

'REDUCTION

INTAKE

•PRESSURE

TUNING

EXHAUST

^ENERGY

RECOVERY

Figure 36. Turbocharger Location Affects Available

Engine Power for RC2-32 Engine (25,000

FT, 250 HP).

3.3.3.2 GTSIO-420/SC Engine

Results similar in form to those for the RC2-32 were obtained

in a less detailed evaluation conducted on the GTSIO-420/SC exhaust

system, where the performance of a partially divided exhaust configuration (2) was compared to a fully divided exhaust system. The

divided system included separate turbines, or a single divided turbine, with a smaller total manifold volume. To obtain substantially better exhaust energy recovery, a 10-fold volume reduction

was necessary. Results of the comparison are shown in Figures 37,

38, and 39. It is recognized that such a drastic volume reduction

would be extremely difficult to obtain without a substantial redesign of the GTSIO-420/SC exhaust system and turbocharger installation.

3.3.3.3 GTDR-246 Engine

The 2-stroke diesel engine was not susceptible to improvements

made by exhaust pulse energy recovery. Techniques that increase

pulse recovery also momentarily increase exhaust manifold pressure.

This increase in exhaust pressure reduces engine scavenging flow,

subsequently reducing combustion efficiency.

57
EXHAUST MANIFOLD

VOLUME

o

UJ

CO

0.6

0.4-

0.2-

CO

9 J

'"

AS PROPOSED

EXHAUST VALVES OPEN

CRANK ANGLE —

MINIMUM.

DIVIDED EXHAUST

CRANK ANGLE •»

Figure 37. Effect of Exhaust Manifold Volume for

GTSIO-420/SC (25,000 Ft, 250 HP).

EXHAUST

MANIFOLD VOLUME AS PROPOSED

ii 6 ?g

S* 4

MINIMUM.

DIVIDED EXHAUST

« u.

^ u.

00 LU

EXHAUST VALVES

CRANK ANGLE — CRANK ANGLE•*

Figure 38. Exhaust Energy Recovery for GTSIO-420/SC

(25,000 Ft, 250 HP).

58
Q.

UJ

co

70-

60-

50-

20-

10-

n-

•^/f^£.

<' -f

%#"'*< '"

s

 Jt •*

^ ^

t :

X.

ADIABATIC

EXPANSION

> PREDICTED

RECOVERY

AS MINIMUM.

PROPOSED

 DIVIDEO

EXHAUST

EXHAUST MANIFOLD SIZE

Figure 39. Exhaust Energy Recovery (GTS 10-420/SC 25,000 Ft

250 HP Divided Exhaust System).

3.3.4 Engine Vibration

Because of their lower load capacity, air bearings are more

sensitive than oil-film bearings to engine and turbocharger rotor

vibration. The vibration estimates for the three advanced technology engine candidates are presented in Table 42. The RC2-32 engine

vibration study was conducted by Curtiss-Wright. The GTDR-246 engine vibration estimates were made from an analysis of reaction

torque. This was based on a NASA bearing load profile(2) which was

then adjusted for engine speed and power rating. The vibration estimates for the GTSIO-420/SC were made through comparisons with

vibration characteristics of existing TSIO engines.

3.4 Turbocharger Conceptual Design

Task III included layout design and performance estimates for

the turbocharger. An estimate of production cost and market penetration for the designed unit had been originally planned as part

of this task; however, the program was redirected to de-emphasize

economic considerations. Instead, efforts were extended in the

area of exhaust pulse utilization analyses and in foil bearing

design.

59
TABLE 42. ATDS ENGINE VIBRATION ESTIMATES.

ENGINE

RC2-32

GTDR-246

GTS 10- 4 20 /SC

MAXIMUM

VIBRATION

LEVEL (g)

5g(1>

3.6g<2>

79( 3>

FREQUENCY OF 1

MAXIMUM '.

VIBRATION (Hz)

314 j

233

160

(1) Vibration at periphery of engine

(2) Vibration at turbine bearing as

(Figure 3-6, NASA CR-3261)

(3) Vibration of turbocharger centerline as

(Figure 23, NASA CR-165162) .

The turbocharger requirements of the three advanced technology

engines were sufficiently different that one design could not satisfy the requirements of all three. Many requirements of the

RC2-32 engine were also characteristic of the other engines; i.e.,

the RC2-32 engine turbocharger was more representative than one designed for either of the reciprocating engines.

The airflow requirements were reviewed for the compressor and

turbine and are shown in Table 43. It was determined that 26 percent of the compressor flow does not have to be delivered at peak

pressure ratio.

The conceptual design effort consisted of establishing the

flow path for the compressor and turbine based on the flows and

temperatures already established. Pressure ratios were determined

from estimated aerodynamic efficiencies and engine airflow characteristics (Table 44). Thus, the size and speed of the rotating

group allowed the geometry of the gas-lubricated foil bearings —

including their spacing — to be established. Off-design performance was estimated, and layout and envelope drawings were prepared.

The following paragraphs discuss the details of this effort.

3.4.1 Compressor

A preliminary design optimization study for the compressor

design point indicated that maximum efficiency would be obtained at

60
TABLE 43. RC2-32 TURBOCHARGER FLOW REQUIREMENTS

(25,000 Ft., 250 BHP).

Requirement

Engine combustion

Engine scavenge

Cabin pressurization

Bearing cooling

Fuel

Total

FLOW

Compressor

36.5

3.7

9.0*

4.9*

—

54.1

(0.90 Ib/sec)

*Bleed may be obtained from an intermediate

(LB/MIN)

TURBINE

36.5

3.7

—

—

1.5

41.7

(0.70 Ib/sec)

pressure level

TABLE 44. COMPONENT DESIGN POINTS FOR 25,000-FOOT

CONDITION (200 TO 225 KTS, 250-ENGINE BHP).

PARAMETER

Pressure ratio

Corrected flow, Ib/sec

Efficiency, %

Inlet temperature, °R

Inlet pressure, psia

Corrected speed, rpm

COMPRESSOR

6.15

2.2

(0.90,

physical)

76.5

440

5.46

85,576

(80,000,

physical)

TURBINE

5.04

0.81

(0.70,

physical)

82.0

1620

27.5

45,300

61
approximately 60,000 rpm. However, preliminary matching with the

turbine on the basis of maximum combined efficiency showed the desirability of 80,000-rpm speed. This 80,000-rpm design speed

slightly compromised compressor performance. However, it resulted

in optimum turbocharger system performance, since the turbine optimized at much higher rotating speeds. In addition, this higher

speed resulted in a more compact unit.

The centrifugal impeller for the ATDS turbocharger is a 50-

degree, backward-curved design incorporating 15 main blades and 15

splitter blades. The radial diffuser incorporates 22 vanes and

exits through a 90-degree bend into a collecting scroll which, in

turn, feeds the intake manifold (Figure 40).

This design was scaled from an existing, demonstrated gas turbine engine compressor. While the ATDS compressor required a different mechanical configuration, the aerodynamic features of the

design were consistent with proven design criteria. The expected

compressor performance map is shown in Figure 41, with relative

design parameters listed in Table 45. This design was selected because it has demonstrated performance in the flow regime below

5-lb/sec, and because the configuration could be manufactured using known processes.

CO

LLJ

5.0-

4.0-

I- 3.0-

2.0-

1.0-

SCROLL

COLLECTOR

0 1.0 2.0 3.0 4.0

AXIAL DISTANCE, INCHES

Figure 40. Compressor Flow Path is Compact.

62
0.0 0.5 1.0 1.5 2.0

CORRECTED AIRFLOW. LB/SEC

2.5

Figure 41. High Efficiency and Pressure Ratio

are Major Advancements.

TABLE 45. COMPRESSOR AERODYNAMIC DESIGN.

PART DESCRIPTION

Blade, qty.

Vane, qty.

B Width, inch

Corrected flow, Ib/sec

Corrected speed, rpm

Pressure ratio

Impeller tip diameter, inch

Impeller inlet diameter, inch

Diffuser exit diameter, inch

Diffuser inlet diameter, inch

15 (full) +

15 splitter blades

22

0.218

2.2

85,576

6.15 (total-static)

5.40

3.55

9.64

5.90

63
Mechanical design of impellers in the 1700- to 1800-ft/sec

tip-speed range is a well-developed technology. Therefore, the

impeller mechanical design analysis was omitted in order to more

thoroughly examine areas of greater uncertainty, such as ceramic

turbine rotor design and use of foil bearings to support the rotating group.

3.4.2 Turbine Aerodynamic Design

The ATDS turbine design included the definition of the gas

path from inlet of the 90-degree elbow (upstream of the scroll) to

the diffuser exit plane. The design objective was to maximize turbocharger performance while maintaining the required geometric envelope. For the ATDS Program, the turbine stator/rotor were first

sized and then assessed for the feasibility of acceptable scroll

configurations and exhaust diffusers.

A parametric study was conducted to define the geometric and

velocity diagrams for candidate stator/rotor configurations. To

accomplish this, specific speed and stage work coefficients were

varied over a wide range. Candidate turbines with diameters

greater than the maximum envelope were disregarded. To optimize

overall performance, the products of turbine total-to-total efficiency and the corresponding compressor efficiency were calculated

for the remaining configurations. The turbine providing the maximum product of compressor and turbine efficiencies was then optimized and is shown in Figure 42.

The turbine scroll was defined using the results of the parametric study. Because of the maximum performance requirement, a

carefully configured geometry was chosen. A large inlet-to-exit

area ratio and overall symmetry about the stator were established

to ensure low losses due to secondary flows.

The exhaust pipe from the engine was located at a 90-degree

angle to the scroll inlet. To ensure desirable scroll inlet flow

conditions, a low-loss "arthritic" elbow(3) was incorporated.

Since no diffuser length limit was set, the exhaust diffuser

was designed with a high (0.76) pressure recovery coefficient, as

shown in Figure 43. This maximized turbine total-to-static performance.

The parametric study resulted in a candidate blade shape similar to the GTEC AGT101 ceramic turbine rotor blade. The turbine

velocity diagram and performance curves are presented in Figures 44

and 45. A comparison of this design with current industrial capability (Figure 9) showed that further technology efforts are required to meet this goal.

64
Figure 42. Turbine Stator and Rotor Flow Path,

14.00

12 BLADES

11 VANES

DIMENSIONS IN INCHES

Figure 43. High Turbine Efficiency Dictates Long Flow Path.

65
STATOR

INLET

ROTOR

INLET

ROTOR

EXIT

Figure 44. Advanced Turbocharger Stage Inlet and Exit Diagrams.

Further turbine performance optimization is needed during the

detailed design process since the wheel is not stress-limited, and

additional performance can be gained by increasing tip speed.

Since the physical speed is fixed, tip speed can only be increased

by increasing the rotor-tip radius. To maintain the given envelope, a trade-off of scroll and stator/rotor performance is necessary.

90

|

g 80

i

§ 70

1.00

PERCENT SPEED «>

110- 3 O.WH

100- »•

0.60

1.0 2.0 3.0 4.0

PRESSURE RATIO

5.0

0.40

110

PERCENT SPEED

1.0 2.0 3.0 4.0

PRESSURE RATIO

5.0

Figure 45. Required Turbine Performance is Close to

State-of-the-Art.

66
Increasing performance by increasing tip speed is a routine

turbine design effort. However, assessing scroll performance

decrements will require rigorous analyses and development testing.

The proposed approach is to design a mixed-flow stator; that is,

one having an axial component of velocity in addition to the normal

radial and tangential components. This will allow the scroll to be

offset in the axial direction such that the maximum radius constraint will not be violated. The scroll configurations will be

optimized using GTEC's 3-D finite-element program, while the stator

will be defined using the GTEC 3-D viscous finite-difference procedure. Stator end-wall contouring will be required to minimize the

secondary flow losses resulting from the nonsymmetric scroll configuration. Using axial-radial turning in the stator instead of in

the scroll will result in the advantage of a higher rate of flow

acceleration within the stator.

3.4.3 Turbine Rotor Mechanical Design

Stress calculations were performed on the ceramic turbine

wheel, including stress-rupture, vibration, and maximum stress

levels. Results showed that a radial wheel aerodynamic design is

mechanically feasible as a ceramic component. No stress-rupture

life limitations were encountered. For the stress-rupture analyses, maximum allowable stress for the Si,N4 ceramic part was set

at <20 ksi.

The GTEC 3-D vibration program, ISOVIB, was used to calculate

the resonant frequencies, the normal displacements, and the relative stress levels of each vibratory mode. The normal mode shapes

for the first ten modes are shown in Figure 46. The frequencies

associated with the first eight modes are shown on the Campbell

diagram, Figure 47. The criterion used to enhance the probability

of success for wheels of this type is to keep the lower order frequencies above the expected 4/rev excitation source. As shown in

Figure 47, the lowest mode was above 6/rev at 111-percent turbine

design speed (89,017 rpm), which can be experienced at 20,000 feet

at climb power.

To meet the stress requirement of all airfoil points at <20

ksi, the GTEC life analysis program, 6WIBUL, was used to determine

the required taper ratio at critical locations at the airfoil (Figure 48) . The blade area ratio required to meet the stress requirement was 7.23:1 (i.e., area at the hub line divided by the area at

the tip). The inducer thickness was held at 0.030 inch. The

inducer radius was 2.5 inches, with a hub-line radius at the discharge of 0.50 inch. This hub-line radius, tip thickness, and area

ratio yielded a blockage at the hub line (i.e., worst case) of

56.9 percent. The maximum blockage at the exducer at 20-ksi maximum stress was 12.76 percent.

67
Figure 46. Turbine Blade Vibration Study Shows Basic

Concept Feasibility.

RESONANT FREQUENCIES

MODE 8 REQUIREMENT: LOWER VIBRATORY

MODES TO EXCEED 4/REV

EXCITATION

RESULTS: FIRST MODE ABOVE

6/REV AT 100% SPEED

(80.000 RPM)

3-0 MODEL

25 50 75

SPEED. PERCENT

Figure 47. Blade Natural Frequencies are Greater Than the

Minimum Requirements.

68
7

6

5

4

3 •

2 -

1

0

<7MAX = 25 KSI

AREA

AT B-1

ADVANCED TURBOCHARGER

RATIO (A/ATIP) VERSUS RADIUS

WIDTH

= 20 KSI

0.4 0.8 1.2 1.6 2.0 2.4 2.8

RADIUS. INCHES

Figure 48. Turbine Rotor Blade Taper Chosen for Low Stress.

Since the wheel was not stress-rupture limited, the area ratio

optimized for maximum stress-rupture life could not be achieved.

The <20-ksi stress requirement was used as the criterion for establishing the taper ratio of the airfoil. A ceramic turbine wheel

has already been successfully operated at 20-ksi by GTEC for the

ACT-101 Program.

In the ATDS ceramic turbine rotor feasibility study, no stress

or vibration problems were indicated that would prevent initiation

of a detailed design effort. However, prior to committing a part

to a production development effort, the following should be investigated:

o Detailed 2- and 3-D analyses should be conducted to verify the integrity of parts prior to development and production.

o Although high strengths were obtained in test specimens

and wheels, surface and embedded flaws in some parts

created stress concentrations which could cause higher

than tolerable stresses. Additional materials testing

should be performed to quantify the effects of long hold

times at high speeds, and the effects of speed and temperature cycling on the long-term properties of ceramics.

69
3.4.4 Ceramic Turbine Rotor Material and Process Analysis

It is recognized that the use of ceramic rotor materials

yields less weight and eventually leads to lower costs. Less

weight, in turn, means less inertia thus lower gyroscopic reactions

and faster acceleration/deceleration. Moreover, this decreased

weight allows the use of foil bearings, which a heavier rotor would

not permit. Due to the fracture properties of ceramic materials,

burst containment shielding can be decreased, thus reducing the

overall weight of the turbine housing. Projected lower turbine

costs were based on the projected lower cost and high availability

of ceramic raw materials when compared to metal alloys. Fabrication costs of ceramics are also predicted to be lower than for

metal alloys.

GTEC has extensive experience in analysis and test evaluation

of both Si3N4 and SiC rotating components. Sintered siliconnitride (SisN4) was selected for the ATDS turbine rotor because it

exhibited high strength at the planned operating temperature.

Moreover, elastic modulus of SisN4 is lower than silicon carbide

(SiC), thus reducing component stress. The strength of Si3N4 ceramics is increasing at a faster rate than SiC.

Injection molding techniques were selected for fabrication

because they are a demonstrated production method that produces

close-tolerance ceramics.

Significant accomplishments have been made in the fabrication

of ceramic components for both DOE and DOD advanced heat engine

programs. These advances have provided evidence that operation of

ceramic parts in high-temperature engine environments is feasible.

Current ceramic development programs, particularly those for ceramic rotors, have been aimed at feasibility demonstrations only,

and not at developing sufficient technology to warrant full-scale

production development. These programs have demonstrated that a

sufficient technology base is needed consisting of additional material and process development, design methodology, and a data base

for life prediction prior to commitment to development and commercial production of reliable, cost-effective ceramic engine components.

3.4.5 Turbocharger Design Details

The preceding evaluations resulted in the ATDS turbocharger

layout shown in Figure 49. The spacing between the two rotor housings was determined by the dynamics of the rotating group in a

foil-bearing suspension system. The weight breakdown and bill-ofmaterial are shown in Figure 50 and Table 46, respectively. This

weight is approximately one-half that of a turbocharger with the

same impeller diameter and using current design approaches. Envelope drawings of the turbocharger are shown in Figure 51.

70
DIMENSIONS IN INCHES

Figure 49. ATDS Turbocharger Cross Section.

The only constraint placed on turbocharger size was a maximum

diameter not greater than the 16-inch diameter of a RC2-32 engine.

The high pressure-ratio and efficiency requirements of the RC2-32

engine necessitated a much larger turbocharger size than for current units. However, ATDS technology represents higher specific

COMPRESSOR

HOUSING

(6.0 LB)

CENTER

HOUSING

[9.0 LB)

TURBINE

HOUSING

(20.0 LB)

TOTAL WEIGHT = 42 LB

Figure 50. Advanced Turbocharger Weighs Half of Conventional.

71
TABLE 46. ATDS CONCEPTUAL DESIGN BILL OF MATERIALS.

PART MATERIAL

Compressor housing assembly brazed

Shroud and vane ring

Scroll

Outer ring

Outlet

Outlet flange

Inlet cone

Inlet flange

Turbine housing assembly - brazed

Shroud and vane ring

Scroll

Inlet

Inlet flange

Outlet cone

Outlet flange

Center housing and rotating assembly

Shaft wheel assembly - bonded

Turbine wheel

Shaft assembly - welded

Shaft

Cup

Sleeve

Thrust disk

Compressor wheel

Washer

Nut

Center housing assembly

Center housing

Bearing foils - journal

Bearing springs - journal

Seal ring

Retaining rings - bearing

Retaining ring - seal

Bearing foils - thrust

Bearing springs - thrust

Seal plate assembly - pressed

Seal plate

Seal ring

Cap screws

Heat shield

Turbine shroud

Shims

Bolts

Lock washers

Shims

Al casting

Al sheet

Al stage

Al sheet

Al plate or tube

Al sheet

Al plate or tube

Fe casting

CRES sheet

CRES sheet

CRES plate or tube

CRES sheet

CRES plate or tube

Ceramic

Steel bar

Nickel alloy bar

Steel tubing

Steel bar

Titanium

Steel

Steel

Al casting

CRES sheet

CRES sheet

Al bar or tube

Steel

Steel

CRES sheet

CRES sheet

Al casting

Al bar or tube

CRES

CRES sheet

Fe casting

CRES sheet

CRES

CRES

CRES sheet

72
155.5 [6.12]

.4 [6.00J

31.7 [125]

COMPRESSOR

INLET

TWIN-ENGINE INSTALLATION

TURBINE

INLET.

184.2 [7.25]

152.4 [6.00]

COMPRESSOR VI

OUTLET

SINGLE-ENGINE INSTALLATION

Figure 51. ATDS Conceptual Design Outline.

power (i.e., turbocharger shaft power divided by airflow) than

today's turbocharger. Moreover, overall powerplant size would significantly decrease for an equivalent power output, due to the

higher working fluid density and higher peak cycle pressure.

Overall performance of the ATDS is better indicated by the

power augmentation it provides the RC2-32 engine. Table 47 lists

both the turbocharged and nonturbocharged power outputs of the engine at several operating points. The compressor map for the ATDS

(Figure 52) shows the engine operating linesxat several conditions.

3.4.6 Bearings

A preliminary design study was completed that investigated the

feasibility of using gas-lubricated foil bearings on the ATDS conceptual design. The bearing load requirements at different modes

of aircraft operation are listed in Table 48, and were based on

information received from Cessna. These loads were increased by

design margins totaling 72.5 percent for use in journal bearing

sizing and load capacity calculations. Total loads were obtained

73
TABLE 47. RC2-32 RATINGS FOR BOTH TURBOCHARGED AND

NONTURBOCHARGED ENGINES.

ENGINE

SPEED

(RPM)

9420

7850

6390

ALTITUDE

(FEET)

0

20,000

0

25,000

0

25,000

ENGINE HORSEPOWER

(TURBOCHARGED) (NONTURBOCHARGED)

320

320

250

250

134

134

168

83

136

55

107

44

e/9 . n

co 4.0

0.5 1.0 1.5 2.0

CORRECTED AIRFLOW. LB/SEC

2.5 3.0

Figure 52. Compressor Performance RC2-32 Engine.

for each mode (Table 49) and compared to bearing load capacity

(Figure 53). Detailed calculations were performed on the journal

bearings. For this application, the journal bearings are expected

to experience more severe loads than the thrust bearings. The

loads imposed on the bearings were derived from:

o Gyroscopic forces from aircraft turn, roll, and spin

o Engine and airframe vibration

74
TABLE 48. AIRCRAFT OPERATIONAL MODE AND ASSOCIATED LOADS.

LANDING

Landing shock loads, g

Aircraft vibration at engine

idle, g

Turbocharger speed, rpm

= 4.25

= 1.5

= 25,000

TURNING

Turn rate

Aircraft vibration at

Cruise throttle, g

Turbocharger speed, rpm

= 75 deg/dec in a 4.0g

coordinated turn at

80-degree bank

= 2.15

= 45,000

ROLLING

Roll rate, deg/sec

Aircraft vibration at

throttle, g

Turbocharger speed, rpm

= 100

= 2.15

= 45,000

SPINNING

Spin rate, deg/sec

Aircraft vibration at

engine idle, g

Turbocharger speed, rpm

= 180

= 1.15

= 25,000

MANEUVERING

Maneuvering loads, g

Aircraft vibration at

cruise throttle, g

Turbocharger speed, rpm

= 4.4

= 2.15

= 45,000

75
OzMs

PL)

§

S

1

J«

. M4J S W •« SI

o .

in £ • k4 3 o

r- 8 •• O 4t CD

132S1*Aircraf •t Crui13

41 4J « O II«s8l

I

Vibra

ne Idle Eng

76
§uEHHU<cmoW(0W

S3

• 2B S ^2

77
60-

50-

30-

LU

a.

£ 20 H

10-

DIAMETER = 2.0 IN

LENGTH = 2.0 IN

10 20 30 40 50 60

ROTOR SPEED. RPM x 10"3

70

Figure 53. ATDS Foil Bearing Load Capacity.

o Turbocharger rotor unbalance

o Landing shock 'g' loads

o Aircraft maneuvering 'g' loads.

The compressor and turbine wheels designed for the ATDS are

similar in size to AID T18A turbocharger wheels. Therefore, physical characteristics of the ATDS turbocharger wheels (i.e., mass,

moment of inertia, center of gravity, unbalance forces, etc.) were

estimated in relation to the physical characteristics of T18A

wheels. In addition, a parallel study of the power loss and load

capacity of a T18A sized free-floating "sleeve oil journal bearing

was completed and used as a basis of comparison for the two bearing

designs. Figure 54 presents the power consumption versus load on

one journal bearing at 80,000-rpm rotor speed. This is approximately 50-percent of that of an oil-film bearing sized for a comparable turbocharger. However, the additional power required for

supplying cooling air is substantially greater than the power required to supply lubricating oil. Depending on the means used for

supplying cooling air, the total power demand for an ATDS air bearing would be from 70- to 250-percent of the total power demand of a

similarly sized oil-film bearing. The results of this design study

showed that:

78
CO

GO

1400

1200

1000

800

600

400

200

DIAMETER = 2.0 IN

LENGTH = 2.0 IN

50 100 ISO

BEARING LOAD. LB

200

Figure 54. ATDS Foil Bearing Power Consumption.

o

o

Two 2- x 2-inch air journal bearings [similar to an

AiResearch design used on the DC-10 environmental control

unit (Figure 55)] will provide sufficient load capacity

for the ATDS turbocharger when subjected to aircraft

operation loads outlined in Table 49.

An aircraft spinning at a rate of 180 deg/sec, at a turbocharger speed of 25,000 rpm at engine idle is expected

to be the worst-case operation mode to which the journal

bearings are subjected.

Load capacity of a 2- x 2-inch air journal bearing at

40,000 rpm turbospeed is 4 times less than that of a

T18A-sized oil journal bearing suitable for a similar

sized machine.

Critical speeds are shown in Figure 56.

The power consumption of one air journal bearing at

80,000-rpm turbospeed at aircraft cruise power is 0.60 hp

(compared to 1.22-hp power consumption for one T18A-sized

oil journal bearing).

79
THRUST

BEARING

JOURNAL

BEARING

FOIL

SPRING

ASSEMBLY

FOIL HOUSING

ASSEMBLED

FOIL SEGMENTS

FOIL SEGMENT

REMOVED

FOIL SEGMENT

REMOVED

BACKING

PLATE

FOIL

ASSEMBLY

Figure 55. Details of Typical Gas Lubricated Foil Bearing

z 100-

= n/w

A = 15

CONICAL

SHAFT SPEED O. RPM > 11T3

Figure 56. Shaft Whirl Analysis of the ATDS Rotor on

Foil Bearings.

80
o Bearing design should be optimized after compressor and

turbine wheel geometries and other turbine/rotor components are finalized, and after completion of a detailed

critical speed analysis of the turbine rotor.

o Total system power loss should be calculated for each design to compare air bearing and oil bearing designs on

the basis of their power consumption. For example, a

total power loss for an air bearing design should include

power loss due to bearing friction and cycle air-bleed

for bearing lubrication and cooling. For an oil bearing

design, this should include power losses due to bearing

friction and oil supply losses for bearing lubrication

and cooling.

o Further development of high-temperature coatings for

foils and journals is also required in order to achieve

the reliability currently demonstrated by Teflon and

polyamide coatings at moderate temperatures.

A TiC foil coating currently used in conjunction with the

Kaman SCA journal coating has demonstrated acceptable bearing performance at temperatures exceeding 1200°F. However, this coating

is very brittle and thickness must be maintained at 15,000 A.

While this coating has demonstrated good wear and friction resistance, its thinness raises questions regarding conformability, which

is characteristic of relatively thick (~1 mil) low-temperature

coatings. Since TiC has no wear-in characteristics, load capacity

and wear tolerance is reduced. A high-temperature foil coating

that possesses such wear characteristics should be identified.

81
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4.0 TECHNOLOGY IMPACT

4.1 Key Technology Requirements

The key technologies required to achieve the proposed ATDS design are listed in Table 50 which includes current status and

availability. For turbochargers, the production capability of gas

turbine engines was defined as laboratory demonstration capability

because of differences in manufacturing technology.

It is evident that high-pressure-ratio compressor efficiency

at low flows must be developed. This need is particularly apparent

in the case of a turbocharger for aircraft diesel engine applications, where a 7.3-pressure ratio at 83-percent efficiency may be

required. The overall industry capability, as previously shown in

Figure 8, indicates that this requirement will be the most difficult to satisfy.

TABLE 50. KEY TURBOCHARGER TECHNOLOGY REQUIREMENTS.

REQUIREMENTS

High-pressure-ratio

aerodynamics

o Efficiency

o Flow range

Foil bearings

o Cooling

o Dynamics

Ceramic turbine

o Useful materials

data base

Lightweight housings

o Containment

o Acoustics

Gas seals

Engine exhaust

system optimization

AVAILABLE

X

X

X

X

X

UNDER

INVESTIGATION

X

X

X

X

X

X

X

X

NEEDS

ADDITIONAL

DEMONSTRATION

X

X

X

X

X

X

X

83
Surge margin (flow range) in high-pressure-ratio compressors

is difficult to obtain where higher pressure ratios are demanded.

Techniques to increase surge margin should begin by increasing the

amount of backward curvature on impeller vanes, and then by progressing to the use of variable diffuser vane angles. To date,

development activity in this area has been concentrated at lower

pressure ratios for backward curvature, and at higher flows for

variable diffuser vanes. Further technology demonstrations are required to ensure that broad surge margins are possible at the flows

and pressure ratios required for advanced turbochargers.

Hydrodynamic lubrication processes (including the use of air)

generate sufficient friction that a significant part of the air

supply required by the bearing is used to remove heat generated by

this friction. The turbine-end foil bearing also requires cooling

because of the heat conducted from the turbine. The total cooling

required is a significant proportion of compressor airflow. As a

result of this cooling requirement, considerable incentive exists

to identify foil coatings that can withstand higher temperatures

and to develop design techniques that reduce aerodynamic losses in

the foil suspension system. To do this, further development effort

is required.

Cold foil bearing support of a rotor is an existing production

capability that is used for aircraft cabin air-conditioning units.

Use of hot foil journal bearings (with no thrust loading) has been

demonstrated in three gas turbine engines. However, the ATDS design requires full support (journal and thrust bearings) of a hot

rotor. This capability has not yet been demonstrated. Furthermore, test experience indicates that the dynamics of the foil suspension favors very light rotors, and that development of hightemperature, low-friction foil coatings is necessary to meet this

need. Foil bearings are sensitive to rotor mass, which increases

as the cube of linear dimension increases, while bearing capability

increases only as a square function. To date, this has prevented

the use of foil bearings in larger turbochargers. Moreover, the

ability of the foil bearing suspension system to absorb shock or

impact loads without rotor contact and foil wear is another uncertainty that requires development effort before foil bearings can be

committed to production.

Ceramic turbine rotors have been the focus of much development

activity, yet the durability and reliability required for an aircraft turbocharger have yet to be demonstrated. An urgent need

exists for a material and process that has consistent, dependable

properties. The most ambitious ceramic demonstration programs now

in work are just that—demonstrator programs with no durability

validation planned. The ATDS Program has shown that continued ceramic material and process development is needed to provide reproducible properties in rotors sized and configured for turbocharger

use.

84
Certification of general-aviation engines to date has been

accomplished using turbochargers with thick housings designed to

contain a 3-piece burst of a heavy metal wheel. The use of a ceramic rotor will permit the•use of a lighter weight turbine housing, since a ceramic wheel burst would generate only lightweight,

small-sized particles. This has provided the motivation to develop

lightweight housings and more carefully assess their containment

capability. The acoustic properties of these lightweight housings

must also be studied, since they may transmit significantly increased noise levels. It should be noted that the design and analysis technology for lightweight housings is already well-developed

for gas turbines, and additional technology demonstration is not

expected to be required.

The use of a foil bearing suspension system could allow

greater displacement of the rotating group during transient maneuvers, startups, shutdowns, and impact loading than experienced with

conventional bearing support. This could result in more frequent

contact with labyrinth seals that are not designed for such contact. This problem could be alleviated with a design that allows

easy seal replacement, but a longer-term solution is needed.

Abradable seal technology is well-developed in gas turbines, and

this approach should be studied as an alternate technique for use

in advanced turbocharger designs.

The exhaust optimization studies conducted in Task III indicated that engine/turbocharger installation efficiency will benefit from further work on the intake and exhaust systems. Basic

analytical tools are available (i.e., UMIST Mark 12, etc.) but

adaptations may be needed for rotary engines.

In summary, further technology demonstration/development programs are required prior to the commitment of the ATDS design for

commercial development. These additional efforts are justified by

the desireable turbocharger features listed in Table 51. Areas of

particular emphasis are listed in Table 52, while the specific factors needing further demonstration are identified in Table 53.

These justifications are examined with respect to other market

needs in the following paragraphs.

4.2 Overall Market Impact

The turbocharger market can be divided into at least seven

categories:

o General aviation

o Highway trucks

o Marine

o Off-highway vehicles

85
TABLE 51. TURBOCHARGER FEATURES HAVE BROAD APPEAL.

O

o

o

o

o

o

o

No oil lubrication

High-temperature capability turbine

Lower weight

Higher pressure ratio for higher engine power density

Less rotating mass

Reduced heat storage

High-efficiency aerodynamics

TABLE 52. SUGGESTED AREAS OF EMPHASIS.

COMPONENT POTENTIAL

Foil bearings

Ceramic turbine

Compressor

Lightweight

housing

Greatest benefit in engine system

reliability improvement

Greatest benefit in weight and

transient performance improvement

Greatest benefit in system efficiency improvement

Greatest benefit in weight and

transient performance improvement

when combined with ceramic turbine

TABLE 53. FACTORS NEEDING FURTHER DEMONSTRATION,

COMPONENT/PART ADVANTAGES

Ceramics and

foil bearings

Foil bearings

Compressor

Ceramic turbine

materials

Applicable for large-sized components

Resistance to impact loads

Greater flow range for commonality

Consistent properties

86
o Automobiles

o Locomot i ve

o Ground power (stationary/mobile).

The ground power category was divided into stationary (large) and

mobile (small) classes to further illustrate the applicability of

the key technologies from the ATDS design. The impact of the

entire design, as well as that of the following four key technologies was used to assess the need for ATDS technology:

o High compressor pressure ratio

o Full gas-bearing-supported rotor

o Ceramics for radial turbines

o Lightweight housings.

The assessment of the turbocharger market impact was based on an

ATDS design that was sized for the RC2-32 engine. Therefore,

market penetration of the ATDS turbocharger was closely tied to

that of the RC2-32 engine. High-altitude operation requires high

inlet boost, and, other than general-aviation aircraft, very few

applications require significant operation at altitudes over 5000

feet. It was uniformly desireable that the turbocharger and engine

lubrication system be separate. Though all other engine applications with a fixed-load operating line require a broad flow-range

turbocharger, the diesel does not. Moreover, engines with a wide

variation in load require good turbocharger transient response. A

standard set of needs and how well they were met by the ATDS was

reviewed for each potential market. These were:

o Increased pressure ratio

o Mounting attitude insensitivity

o Small size (volume)

o Large range of flow capacity

o Higher efficiency

o Greater reliability

o Independence from engine lubrication system

o Large surge margin

o Better transient response

o Impact loading tolerance.

87
4.2.1 General-Aviation Market - The ATDS Program has shown that

the proposed design has the potential for major impact in the

general-aviation market. Its four key technologies supply the specific needs listed in Table 54. It is expected that at least three

variations to the design will be needed to match the pressure/flow

characteristics of Otto, rotary, and diesel engines (expected to be

in production in the 1990's). Size and exhaust considerations

dictate rear installation. In the event that very high-pressureratios are needed by diesel aircraft engines, a two-stage compressor may be required. Ultimately, a turbocharger would be optimized

for each engine class, since turbocharger market penetration is

tied to individual engine model installations.

4.2.2 Truck Market - By far the largest turbocharger market at

this time is for highway truck engines. A number of turbocharger

models are available for this market, with approximately 200,000

units-per-year produced. Truck engines require a turbocharger with

good transient response and broad surge margin (Table 55). At high

pressure ratios, both requirements could be met using single-stage

units with variable geometry. Another way to meet this requirement

is to use two stages in series, or possibly two-spool units. Small

size is almost as critical a factor as with aircraft applications.

The high utilization of trucks places a premium on efficiency over

TABLE 54. GENERAL-AVIATION MARKET.

NEEDS ATDS-SUPPLIED

Increased pressure ratio

Mount attitude insensitivity

Small size

Higher efficiency

Greater reliability/Life

Independence from engine lube system

Yes

Yes

No

Yes

Yes

Yes

PENETRATION BY ATDS

Current design - As great as that of each engine application

Selected components - Major (turbo optimized for engine class)

88
TABLE 55. TRUCK MARKET.

NEEDS

- Increased pressure ratio

Wide flow range

Small size

Higher efficiency

Greater reliability/life

Better transient response

Independence from engine

lube system

ATDS-SUPPLIED

Yes

No

No

Yes |

f

Yes

Yes

Yes

PENETRATION BY ATDS

- Current Design - Minor (a range

of sizes is needed)

Selected components - Major: (foil

bearings, low-mass turbine housing, ceramic turbine)

a broad range, which also favors the two-stage turbocharger.

Ceramic rotor and foil-bearing supension would be of particular

value in truck applications, to provide independence from the

engine lubrication system, which is relatively dirty.

4.2.3 Marine Market - The marine market is not as large nor as

price-sensitive as the highway truck market. Marine diesel engines

range from less than 100 to over 20,000-horsepower and require a

large range of turbocharger flow. Because most engines operate

under fixed propeller load characteristics, this turbocharger

application does not require broad surge margin. It is not known

how easily the ceramic turbine and foil bearings will scale up to

accommodate larger flow capacities. Specific speed considerations

indicate that axial flow turbines are preferable with large engines. ATDS suitability for the marine market is summarized in

Table 56.

89
TABLE 56. MARINE MARKET.

NEEDS ATDS-SUPPLIED

Increased pressure ratio

Large range of flow capacity

Higher efficiency

Greater reliability/life

Better transient response

Independence from engine

Lube system

Yes

Partly

Yes

Yes

Yes

Yes

PENETRATION BY ATDS

Current design - moderate (limited

by engine applications)

Selected Components - Major (turbo

optimized for engine class) to moderate will depend on scaled-up capability for ceramic wheel and foil

bearings)

4.2.4 Off-highway Market - The off-highway market (including

military tracked vehicles) is dominated by diesel engines and has a

moderately wide range of flow capacity requirements (Table 57). It

is expected that rough terrain would place high impact loading on a

turbocharger rotor. Thus, air bearings might not have adequate

load capacity for this application. Depending on the drive train

design, high surge margins may or may not be required. Again,

since engine scale-up capabilities are uncertain, ceramic turbine

and foil bearing technologies may not be applicable to larger sized

engines in off-highway applications.

4.2.5 Automotive Market - The automotive market offers high production potential for small turbochargers. Automobile engines require good surge margin and small-sized turbochargers. This can be

accomplished by adapting the ATDS technologies to the specific

needs of the automotive engine (Table 58) . The problem of downscaling these technologies from the ATDS is one of manufacturing at

90
TABLE 57. OFF-HIGHWAY MARKET.

NEEDS ATDS-SUPPLIED

Increased pressure ratio

Varied flow range

(Mechanical Trans - wide range

Hydraulic Trans - narrow range)

Large flow capacity range

Higher efficiency

Greater reliability/life

Better transient response

High-amplitude vibration

tolerance

Independence from engine lube

system

Yes

Partly

Partly

Yes

Yes

Yes

No

Yes

PENETRATION BY ATDS

Current Design - Moderate (mostly on

units with hydraulic transmission)

Selected Components - Moderate (i.e., foil

bearings may not have adequate stiffness)

to Major (depends on scaled-up capability

for ceramic wheel and foil bearings)

a competitive price; i.e., the performance difference between tomorrow's and today's technologies may not justify the cost differences unless advanced manufacturing technology developments are

made that allow new turbochargers to use processes similar to

today's turbocharger.

91
TABLE 58. AUTOMOTIVE MARKET.

NEEDS

Mount attitude insensitivity

Increased pressure ratio

- Wide flow range

- Very small size

Higher efficiency

Greater reliability/life

Better transient response

Independence from engine lube

system

ATDS-SUPPLIED

Yes

Yes

NO

No

Yes

Yes

Yes

Yes

PENETRATION BY ATDS

Current Design - Minor (A range of sizes

is needed)

Selected Components - Major (easily

scaled down)

4.2.6 Locomotive Market - The locomotive market is small, and uses

large diesel engines almost exclusively. The high-pressure-ratio

compressor is the only ATDS technology that can definitely be

scaled up for a locomotive turbocharger (Table 59).

4.2.7 Ground Power Market - The ground power market is divided

into two groups—mobile and stationary engines. Although each

group includes a wide range of airflow capacities, the mobile

market is dominated by small engines, while the stationary market

is dominated by large engines. ATDS technology is applicable to

mobile engines. However, stationary engines will require large

turbochargers, and uncertainty exists concerning the scaling of

ATDS technology. The needs of the ground power market (Table 60)

are not unique. However, a general improvement in turbocharger

parameters would be beneficial in this area.

92
TABLE 59. LOCOMOTIVE MARKET.

NEEDS

Increased pressure ratio

Large flow capacity

Higher efficiency

Greater reliability/life

Independence from engine lube

system

ATDS-SUPPLIED

Yes

No

Yes

Yes

Yes

PENETRATION BY ATDS

Current Design - Minor (Too small)

Selected Components - Major (If

scale-up capability exists)

TABLE 60. GROUND POWER MARKET.

NEEDS

-

-

-

-

-

-

~

Increased pressure ratio

Wide flow range

Large flow capacity range

Higher efficiency

Greater reliability/life

Better transient response

Independence from engine

lube system

ATDS-SUPPLIED

Yes

Yes

Partly

Yes

Yes

Yes

Yes

PENETRATION BY ATDS

Current Design - Moderate (for

engines with airflow match)

Selected Components - Major (Family

of sizes needed)

93
4.2.8 Market Impact Summary - A matrix summary of the impact of

the ATDS technologies is shown in Figure 57. The perspective of

the entire market indicates that emphasis should first be placed on

developing a ceramic turbine — not so much because of its temperature capabilities, but because of its light weight and potentially

low manufacturing cost. The second-most important technology for

further study is the foil bearing, principally because it reduces

the risk of damage and/or shutdown of the engine. The third most

important technology is the high-pressure-ratio compressor, which

will provide engine cycles with higher compressor efficiency and

pressure ratio.

A cursory estimate of the benefits of the ATDS technology

based on an assumption of a one-million-unit annual turbocharger

market in 1990 is shown in Table 61. This table also shows the

utilization assumed for a representative application in each

market. The assumption of only a 1-percent SFC improvement results

in a potential gain of $54-million-per-year for newly produced

engines. Based on an average useful life of at least 10 years, the

lifetime benefit would be $540 million dollars. Moreover, the

weight savings offered by the ATDS technology for weight-sensitive

applications would result in a multimillion-dollar benefit in DOC

savings.

4.3 ATDS - Related Programs

Several technology programs underway or recently completed

that involve scaled components of turbocharger size, or components

that are scalable for turbocharger use (Table 62) are:

ATDS HIGH- LIGHT

DESIGN PRESSURE FOIL CERAMIC WEIGHT

IA8-I8) COMPRESSOR BEARINGS TURBINE HOUSING

GENERAL AVIATION 0 0 00 0

TRUCK • • 0 0 •

OFF-HIGHWAY • • • 0 •

MARINE • • • 0 •

AUTOMOTIVE • • 0 0 •

LOCOMOTIVE • • • • •

GROUND POWER

(MOBILE) • • 00 *

(STATIONARY) • • • • .

MAJOR =m MODERATE =• MINOR =

Figure 57. ATDS Technology will have a Broad Market Impact

94
TABLE 61. OVERALL MARKET IMPACT SUMMARY.

Generalaviation*

Truck*

Marine

Automotive*

Locomot i ve

Off-highway

MEAN MEAN MEAN FUEL COST

CRUISE X CRUISE X HRS/ v . ,-,, c . S/YR/ UNITS/

SHP SFC YR l.^u/o.3 - UNIT f. yR - S X 106/YR

250

100

200

20

2000

200

0.42

0.35

0.35

0.55

0.35

0.50

300

1000

500

200

1500

500

10,000

400,000

100,000

200,000

1,000

100,000

58

2,585

646

81

194

925

TOTAL S5,414 X 106

•Weight-sensitive

1-oercent Fuel Savings = S54 Million/Year

NASA Scaled Centrifugal Compressor Program (Contract

NAS3-2243) demonstrated the scalability of an existing

25-lb/sec design to both 10- and 2-lb/sec. However, no

development activity is planned.

AFAPL Variable-Cycle Technology Propulsion System Assessment Program (Contract F33657-79-C-0726) developed

an extended surge margin in a 25-lb/sec centrifugal compressor through the use of variable geometry. No scaling

of the resulting design is planned.

TABLE 62. EXISTING TECHNOLOGY PROGRAMS.

COMPONENT

Compressor

Ceramic turbine

Foil bearing

Lightweight housing

PROGRAM

Scaled centrifugal

Variable Cycle Centrifugal

AGT101

Injection molding demo

ACT 101

TJE331 foil bearing

demonstrator

APU gas-lubricated

foil bearing

None

SPONSOR

NASA

AFAPL

DOE

AMMRAC

DOE

AFAPL

AFAPL

—

STATUS

In test

In test

Preliminary design

Hardware

Early testing

AF planning

Testing underway

—

95
o NASA/DOE AGT101 Program (Contract DEN3-167) entails

operation of a demonstrator engine by 1985, with both

ceramic turbine and foil bearing applications included.

However, the foil bearing design is a journal bearing

only, with the rotating group supported at the compressor

end with a ball bearing that also absorbs thrust. Again,

no provision exists for production development.

o AMMRC Low-Cost, Net-Shape Ceramic Radial Turbine Rotor

Program - Demonstrated the capability of injectionmolding the appropriate geometry; however, durability or

reliability were not taken into consideration.

o AFAPL TJE331 Foil-Bearing Demonstrator Program (not yet

contract status) - Though demonstration is planned, no

production development is planned as part of the contract. The TJE331 is a single-spool, expendable turbojet

engine. For this program, the design would be modified

to incorporate full foil bearing support of the rotating

group. Again, development for production is not part of

the planned program.

o AFAPL Gas-Lubricated Foil Bearing Development Program

(Contract F33615-73-C-205) - A nearly completed demonstrator program. Development for production is not planned. Incorporates a bearing system configuration like

that used in the AGT101 (i.e., ball-thrust and foil journals) .

It is apparent that none of these programs provide the essential turbocharger technology that would allow commitment to production turbocharger development. The prospects for success of each

program are good to the level of technology stated; that is, a feasibility demonstration of objectives. However, only a few of these

objectives coincide with the objectives of an advanced turbocharger

design. NASA modification or addition to several of these programs is possible, but even this would not ensure reaching the

long-range goals of the ATDS Program.

Formulation of component technology demonstration programs

specifically directed toward turbocharger development is necessary

to establish the ATDS turbocharger concept for production. The

programs listed in Table 63 are necessary to accomplish this. Detailed recommendations for each program are discussed in Section 5.

96
TABLE 63. CONCEPTS THAT MUST BE ADDRESSED BY FUTURE PROGRAMS

COMPONENT/PART NEED

Ceramic turbine rotor

Foil bearing

High-pressure-ratio

compressor

Lightweight housing

Turbocharger

Characterization and demonstration

Operating parameters

Flow range and efficiency improvement

Containment and acoustic investigation

Full-scale demonstrator

97
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5.0 TURBOCHARGER TECHNOLOGY - FUTURE PLANS

5.1 Summary

Advanced technology engines developed in response to NASA's

program for general-aviation aircraft will require turbocharging to

achieve design performance. Turbocharger pressure ratios of 5 to

8, required by these advanced technology engines at 25,000 feet, are

substantially higher than any existing turbochanrgers can provide.

Moreover, turbochargers designed for these new engines using

today's technology would be substantially heavier (85 versus 42 lb)

than any units currently acceptable for aircraft application. The

alternative of using two turbochargers in series to achieve the

necessary pressure ratios would provide acceptable performance but

would still be heavy and would make it harder to install the turbocharged engine in an aircraft nacelle. Therefore, new turbochargers will be needed that require the use of advanced technologies to support the advanced technology engines and to attain the

desired goals in improved fuel consumption and performance. These

advanced turbochargers require:

o Determination of the compound (engine and turbocharger)

cycles

o Definition and build of turbochargers appropriate to each

engine cycle

o Development of additional technology in air bearings,

ceramics for rotating aerodynamic components, and control of variable-geometry elements.

The Garrett Corporation has contributed substantial advancements to the state of technology in each of these areas. The

resulting technical capabilities and technology base will be used

in acquiring the advancements necessary for an advanced technology

turbocharger.

5.1.1 Cycle Definition

Cycle definition will require effort to coordinate requirements and available technology for the intermittent combustion (1C)

engine and turbomachinery portions of the complete compound engine.

AID has contributed to optimization of the overall cycle for its

many engine customers through its application of engineering expertise and custom aerodynamic development capabilities. GTEC has

developed an extensive capability for integration of combustion

cycles with turbomachinery aerodynamics, and has also developed

compound engines combining 1C engines with turbomachinery. These

technologies have direct application to the Advanced Technology

Engine Program.

99
5.1.2 Advanced Turbochargers

The assembly and test of turbochargers incorporating all the

advanced technologies is required to determine the proof of concept. The conceptual design portion of the advanced turbocharger

design study is an example of layout of an advanced technology test

unit for the RC2-32 engine. GTEC and AID routinely design and

build prototype and breadboard turbomachinery for specific application and laboratory investigations. While simulation of turbocharging (most applicable to a constant-pressure application) and

adaptations of off-the-shelf turbochargers for some individual

operating points can assist in investigating engine operating characteristics, the final testing of engine technology advancements

cannot be considered complete without including a complete, appropriately sized turbocharger. No program currently under way incorporates such high-pressure-ratio aerodynamics in a practical turbocharger, even if the recommended features of air bearings, variable

geometry, and ceramic turbine were deferred.

5.1.3 Required Technology Advances

5.1.3.1 Air Bearings

Air bearings have been employed in production bearing systems

for rotors experiencing moderate gas temperatures, and in development programs as hot-end bearings for gas turbine engines. Air

bearing demonstration and development programs currently under way

are shown in Figure 58. These programs and production experience

have shown that gas-lubricated foil bearings are capable of withstanding starting and shutdown friction at elevated temperatures,

and that bearings can be designed with adequate dynamic stiffness

at moderate temperatures. None of these programs, however, has yet

addressed the dynamics of a rotor fully supported on air-lubricated

journal and thrust bearings which are capable of withstanding large

changes in operating temperatures. This need is currently being

addressed at AID through technology programs for small turbochargers, and is also under consideration by the Air Force Aero

Propulsion Laboratory (AFAPL) for short-life applications. However, these programs do not fully address the requirements for long

life and the large sizes needed for an advanced turbocharger.

5.1.3.2 Ceramics

Ceramic technology efforts currently underway are shown in

Figure 59 and include:

o ACT Programs

o AMMRC/Garrett Radial Rotors Program

100
1980 1970 1975

I

1076 1977

I

1978 1979

I

I960 1961 1982

I

kGTP30

DEMO

BEARING DEVELOPMENT

ON RIGS

X

COMPANY-SPONSORED FOIL BEARING

(DEVELOPMENT. MATERIALS & ANALYTICAL)

SIZE: 025- TO 45-IN DIAMETER

LOAD CAPACITIES TO 32 PSI AT 500° F

VARIOUS CLOSED-CYCLE

PROGRAMS WITH FOIL

BEARING DEVELOPMENT

NASA 2-KW MINI-BRU

DESIGN. FABRICATE. ENGINE

TEST. ANALYTICAL TOOLS

USAF FOIL BEARING

PROGRAM PHASE I

(F33615-73-C-2058)

PREVIOUS FUNDED PROGRAMS

JFS100 DEMO

CRITICAL SPEED CONTROL

AIR CYCLE TURBOMACHINERY

• OBJECTIVE — DESIGN.

FABRICATE

AND RIG TEST 45' n 6.<T

FOIL BEARING FOR INTERMEDIATE THRUST ENGINES

• PROGRESS

• MAT1 INVESTIGATION

(NASA PGM VALIDATED

= 1 MAT! SELECTION)

I RIG TESTS ON JOURNAL

TO ~ SO PERCENT LOAD

CAPACITY

I THRUST BEARING

DEVELOPED

IMPROVED LOAD CAPACITIES/

HIGH-TEMPERATURE CAPABILITY

1985 1987

• VEHICULAR FOIL (JOURNAL)

BEARING DEVELOPMENT

• BEARING US INCH OIA.

• TEMP. TOLERANCE TO 1200° F

1

• USAF APU HOT END FOIL BEARING DEV

F33615-78-C-2044 (JOURNAL ONLY. 3.5 IN)

• DESIGN • FAB BEARING RIGS

• RIG TEST • ENGINE DEMO

USAF APU v

FLIGHT |

DEMONSTRATION/

3-5-IN BEARING

30-PSI LOAD

1000° F SOAKBACK

1200° F TEMP. TOLERANCE

' INTERMEDIATE SIZE BEARING APPLICATIONS

SECONDARY POWER SYSTEMS

• STRATEGIC BOMBERS

• TRANSPORT AIRCRAFT

• GROUND CARTS

• SMALL INDUSTRIAL ENGINES

[ADVANCED DEVELOPMENT

TEST TURBOCHARGER

FOIL BEARINGS

GOALS • 2 IN. JOURNAL. ~4 IN THRUST BEARINGS HIGH OPERATING & SOAKBACK

TEMPERATURES (DETERMINED BY COOLING DESIGN)

• STABLE ROTOR SUPPORT OVER WIDE RANGE OF OPERATING TEMPERATURES

USAF TJE331-1029

FOIL JOURNAL

BEARING

DEVELOPMENT

r

GOAL

DESIGN FABRICATE &

TEST 35 IN. DIA.

FOIL BEARING

o 23.3 PSI LOAD

o 1200°F TEMP,

o RIG TEST ONLY

USAF TJE331-1029

GAS TURBINE ENGINE

| GROUND AMBIENT

I PROPOSED PROGRAM

(UNCERTAIN FUNDING)

\

USAF

FLIGHT

// DEMONSTRATION /

»1..JL/

LARGE BEARING APPLICATIONS

• INTERMEDIATE THRUST ENGINES

• MARINE ENGINES

INDUSTRIAL PRIME MOVERS

• RPV .

• POWER GENERATORS

• EXPENDABLE ENGINES

• TANK ENGINES

I

NOTE:

DASHED BOX INDICATES INCOMPLETE PROGRAM

U.S NAVY-NAPC

ADVANCED APU/SPS

FOIL BEARING

DEVELOPMENT

• fiflfli

• 46-PSI LOAD

CAPACITY

• 1.75 IN. DIA.

BEARING

RIG TESTS AND GAS

GENERATOR DEMO.

USN ADVANCED

J APU DEMONI STRATION

• PROGRAM AOVANTA6ES

• ELIMINATES NEED FOR HIGH

ON ROLLING ELEMENT

BEARINGS

I VALIDATES SCALEABILITY

I USES HIGH-TEMP. MATERIALS

I CORRELATION OF ADVANCED

DESIGN (ELASTO-HYORODYNAMIC)

TOOLS

• SMALL BEARING APPLICATIONS

• ADVANCED APU FOR ADV. FIGHTERS/

V-STOL AIRCRAFT

• GENERATOR SETS

• CATHOOIC CORROSION UNITS

• CRUISE MISSILES

• GENERAL-AVIATION APUt

• GROUND POWER UNITS

• VEHICULAR

Figure 58. Foil Bearing Technology Roadmap.

101
1981 1982

CALENDAR YEAR

1983 I 1984 1985 1986 1987

DOE/NASA. GARRETT, AND ALLISON AGT PROGRAMS WILL RESULT

IN SHORT TIME OPERATION OF A PROOF-TESTED CERAMIC ROTOR

IN A 100-HP PROTOTYPE ENGINE

AMMRC/GARRETT RADIAL ROTORS

• MATERIAL PROPERTY DEVEL

• FABRICATION AND SPIN TEST OF

25 PROTOTYPE SMALL-SIZED

TURBOCHARGER ROTORS

DECISION OF AUTO

COMPANIES TO PROCEED

AUTOMOTIVE COMPANIES

FULL-SCALE DEVELOPMENT

AND RELIABILITY DEMO 1..J

TACOM/CUMMINS

LARGE TURBOCHARGER ROTOR

FEASIBILITY STUDY

PROGRAMS NEEDED TO DEVELOP A

LARGER TURBOCHARGER ROTOR

• FAB PROPERTIES TEST ROTORS

• LONG-LIFE DATA BASE

• THERMO/MECHANICAL DESIGN

INDUSTRY-FUNDED SMALL

TURBOCHARGER ROTOR PROGRAMS

• MATERIALS

• MANUFACTURING PROCESSES

NOTE:

DASHED BOX INDICATES UNFINISHED PROGRAM

r

"~NASA/DOE CERAMICriNITIATIVE~

• MATERIALS RELIABILITY IN TEST BARS

• PROPERTIES CORRELATION IN NETSHAPE COMPONENTS

Figure 59. Ceramic Rotor Technology Roadmap.

102
o TACOM/Cummins Large Rotor Study

o Garrett Small Turbocharger Programs.

o NASA/DOE Ceramics Initiative.

Though each of these programs will contribute technology that

will be useful in the advanced turbocharger technology development

program, none provides the long-term materials properties needed to

undertake the design of a turbine with adequate reliability for an

aircraft engine. Material properties that are both a function of

material processing as well as size will not be sufficiently

addressed in the turbine wheel size (i.e., 5.5-inch diameter)

required of the advanced turbocharger.

The AMMRC and Garrett company-funded programs are farthest

along in studying material properties in actual aerodynamic shapes,

but are directed at rotors of 2- to 3-inch tip diameter. Material

property development in the AMMRC/Garrett Radial Rotors Program is

limited to short-time properties. The ACT Programs, which address

an appropriate sized turbine, have met with success only by selecting samples with acceptable short-term properties. None of the

currently active programs, including the Cummins Feasibility Study,

address the determination of long-term material properties. Though

the ceramic initiative being pursued by NASA/DOE has recognized the

importance of both long-term property characterization and reliability design techniques, these aspects will not be addressed

until the late 1980's. This would delay the availability of the

technology advances demanded by the general-aviation market.

However, experience gained in the ACT and AMMRC programs indicates that the short-term properties of silicon nitride are good

enough to justify the design of a rotor for an advanced technology

turbocharger. Sufficient design latitude was found during the ATDS

conceptual design effort to give reasonable expectation that a

suitable design can be achieved once the long-term properties are

k nown.

5.1.3.3 Variable Geometry

Garrett has acquired substantial experience in several areas

related to control requirements for variable geometry. Turbine

engines manufactured by GTEC use integrated control systems for all

phases of operation that include some variable-geometry elements.

However, performance optimization for turbine engine control is

related only to the combustion process and the turbomachinery aerodynamics. The additional complexity of the 1C engine as the power

producer creates a new requirement for technology development. AID

has current experience in controls for exhaust bypass valves on

turbocharged 1C engines, but this control function has not been

103
integrated with other controls on the engine. Past AID programs

have controlled variable turbine inlet guide vanes to maintain constant compressor pressure rise, but again, the overall engine performance has not been thoroughly optimized.

5.2 Recommended Follow-On Effort

The requirements definition, technology survey, and studies in

connection with the conceptual design effort show the feasibility

of advancements in turbocharger technology for general aviation.

The key consideration in the advancement of turbocharger technology

is the integration of the engine and turbocharger technology demonstrations and eventual development activities shown in Figure 60.

This plan has a baseline goal of 1991 for an aircraft incorporating

advanced engine and turbocharger technologies, and is based on an

optimistic schedule that includes the development of an advanced

engine. Morever, proper timing of the advanced test turbocharger

programs is necessary so that capabilities are assured prior to

committing the turbocharger to the performance requirements of a

specific engine. It is apparent that to meet these goals, the following component technology and coordination programs must begin

immediately. Slippage in initiation of the component programs will

either lead to a delay of turbocharger availability, or force the

design to a lower technology level.

1983 1984 1985 1986 1987 1988 1989 1990

GOVERNMENT

INITIATIVE

REQUIRED

GOVERNMENT

INITIATIVE

REQUIRED

TURBOCHARGER PRELIMINARY DESIGN

TURBOCHARGER TECHNOLOGY DEMONSTRATION

HIGH-PRESSURE-RATIO- COMPRESSOR

FOIL BEARINGS

CERAMIC TURBINE

LIGHTWEIGHT TURBINE HOUSING

VARIABLE-GEOMETRY TURBINE NOZZLES

TURBINE AERODYNAMIC DEVELOPMENT

TURBOCHARGER TEST UNIT

PRODUCIBILITY EVALUATION

TURBOCHARGER

ENGINE

TURBOCHARGER DEVELOPMENT

" ^ E"NGIN?PR?LIMiNARY~DESIGN

__ ENGINE. ENABLING TECHNOLOGY

ENGINE DEVELOPMENT ~"

LATEST AVAILABILITY OF TURBOCHARGER COMPONENT TECHNOLOGY

THAT WILL ENABLE GOAL FOR 1991 AIRCRAFT INTRODUCTION

Figure 60. Integration of Engine and Turbocharger

Technologies/Development Activities.

104
The following recommendations describe the required programs,

and indicate whether current industry efforts will achieve suitable

results or whether additional efforts are required.

5.2.1 Technology Requirements Coordination

Coordination of the preliminary designs of engine, turbocharger, and related control components and systems, considering

continuing technology advances, is fundamental to ensuring that

support continues to be directed toward the most advanced technology programs.

Successful integration of engine, turbocharger, and control

system technology requires thoughtful assessment of the relative

difficulty of obtaining advances in each technology. Tradeoffs of

potential advances must be continually reviewed throughout preliminary design of the engine, including its appropriate turbocharger.

Technology and design coordination would define aerodynamic cycle

requirements, establish thermal loading of the turbocharger as

affected by mounting location, coordinate cooling provisions for

the engine and turbocharger, and optimize turbocharger mounting

dynamics and vibration response.

A program should be initiated to coordinate technology

advancement requirements for general aviation engines. This program should be conducted in conjunction with, and would aid in managing technology advancement programs for engine and turbocharger

technologies.

The design phase of each advanced technology engine should

include—at least quarterly—cycle design and technology reviews

attended by the engine and turbocharger contractors. At each

review, the more difficult requirements of the engine and turbocharger should be presented for discussion of alternative technologies. Following each such review, the engine and turbocharger participants would report their evaluations of the alternatives presented.

5.2.2 Advanced Component Turbocharger Testing

An important milestone in NASA's general-aviation engine technology advancement programs should be the buildup of several turbocharger units utilizing components derived from the separate advanced technology programs. Other program recommendations herein

reflect the technology advancements appropriate and necessary to

achieve this milestone. Significant technologies that should be

incorporated in the turbochargers include variable geometry of turbines and/or compressors, air bearings, and ceramic turbine rotors.

To date, these three concepts have never been combined in a single

turbocharger design. Each of these concepts by itself should improve turbocharger life and performance.

105
The initial advanced turbocharger configuration could be based

on the design generated in the ATDS for the RC2-32 rotary engine.

This design could be modified to incorporate the use of a variable

turbine inlet nozzle. This configuration could be established as

soon as the interface configuration, engine speed, and engine displacement are determined. Design efforts would establish configurations of the aerodynamic rotating group, air bearings, and variable turbine inlet nozzles. The turbine should be designed using

the best available information on properties of available ceramic

materials. It is expected that enough acceptable turbine wheels

can be produced to satisfy the requirements of the engine technology demonstration. Concurrent with the advanced turbocharger testing program, separate rotor testing should be performed to determine ceramic property data for correlation with test bar data.

Variable turbine inlet nozzles should provide some improvement

in engine efficiency and at the same time, provide better control

of the fuel/air ratio for the four-stroke cycle of the rotary

engine. This technology would also improve performance of the twostroke cycle diesel engine by maximizing the scavenging pressure

drop across the engine. The degree of improvement in each case

would depend greatly on the control logic selected. Each engine

and turbocharger designer would have the opportunity and responsibility to investigate logic alternatives, including parameters to

be sensed and the options of open-loop versus closed-loop control

of parameters of interest. While a schedule of vane position or

engine inlet manifold pressure versus operating conditions may be

sufficient to obtain acceptable operation of the engine, there are

opportunities for interactive optimization of fuel/air ratio or

even of fuel consumption at any operating condition.

Concurrent with the RC2-32 Advanced Turbocharger Program,

definition of the aerodynamic cycle for the 1992 Advanced Technology Diesel Engine (ATDE) should begin. The coordination program

recommended above would be instrumental in optimizing the complete

engine/turbocharger cycle, including selecting accessories related

to semi-independent operation of the turbocharger. Cooperative

evaluation of the requirements for achieving high pressure ratios

with good efficiency, compared with the benefits of minimizing

engine size and speed, would establish the degree of turbocharging

to be used. Cycle analysis and preliminary aerodynamic design

would then establish whether variable compressor geometry as well

as variable turbine inlet nozzles are desirable. During cycle

definition, substantial effort should be devoted by the turbocharger and engine designers to determine the optimal control logic

for the variable geometry features. Progress on the diesel engine

would not be impeded by a slightly later start, since laboratory

simulation of turbocharging would acceptably represent actual turbocharging for the two-stroke diesel engine during its preliminary

technology and cycle development.

106
Requirements for an advanced turbocharger for the advanced

technology spark ignition engine should be determined in conjunction with the technology test program for that engine.

A program should be initiated for advanced turbochargers for

the candidate advanced technology engines. Each turbocharger

should be developed in four phases:

o Preliminary design of the machine and its controls

o Rig testing as required to confirm design characteristics

o Construction and testing of the advanced turbocharger on

a hot gas test stand

o Delivery of the unit to the engine manufacturer for

testing.

5.2.3 Component Technologies

5.2.3.1 Air Bearings

Use of air bearings would eliminate turbocharger dependence on

the engine oil system. Engine wear particles could not harm the

turbocharger bearings. Weight of oil lines would be eliminated,

along with the problems of safely routing them to the turbocharger.

Heating of engine oil would be reduced, and oil drain provisions

would no longer constrain turbocharger mounting attitudes. To gain

these benefits in turbochargers for advanced technology engines,

technology development is required in areas of thermal control,

cooling, duty cycle determination, and bearing stability.

5.2.3.1.1 Air Bearing Temperature Control Technology

Successful application of air bearings to turbochargers may

require new techniques of temperature control in the bearing housing. To define technology requirements in this area, a comprehensive thermal study of a turbocharger bearing system is necessary.

The conceptual design developed by the ATDS is an appropriate model

for this study. Evaluation of various means of temperature control

at the bearing foils would assist in setting realistic targets for

foil coating temperature capability and contribute input to ongoing foil coating development and evaluation programs. Estimates

of bearing temperature changes in operation would also contribute

design input to the bearing dynamics technology program. Turbine

inlet temperatures considered should allow for the effects of

reduced engine cooling in future engine technology advancement programs. Program results would be sufficiently broad in scope to

benefit a variety of advanced engine technologies. No existing

program is structured to produce parametric results to guide application of bearing temperature control technologies over this broad

operating range.

A comprehensive thermal study of the ATDS bearing system

should be performed in support of the overall effort of designing

107
advanced turbochargers. Parametric studies should be made to evaluate the effects of changes in materials conductivity/ areas of

thermal flow paths, cooling air flow velocity, pressure, and temperature, and turbine inlet gas temperature on bearing component

temperatures.

5.2.3.1.2 Air Bearing Cooling Air Sources

Alternative means of providing air for bearing cooling should

be evaluated. The high compressor pressure ratio required in aircraft applications produces hotter and higher pressure air than

required for bearing cooling. Aircraft applications have unusually

difficult requirements in these respects, and there is no existing

program that will supply applicable information. Possible air

sources for cooling include the turbocharger compressor, either at

the normal discharge station or at an intermediate pressure bleed

location; mechanically or pneumatically driven air compressors or

fans; and aspiration of ambient air to augment a reduced compressor

bleed flow. Definition of air temperature and pressure would

establish these parameters for use in the bearing system thermal

study. An appropriate air source might also supply air to pressurize the cabin at less overall energy cost. Evaluations of possible air sources should be used in design and technology tradeoffs

by engine and turbocharger designers.

A program should be initiated to evaluate alternative air

sources for bearing cooling. This study should include cycle analysis and performance specifications of alternate cooling air supply

components, and comparisons of engine and turbocharger performance.

5.2.3.1.3 Air Bearing Duty Cycles

The air bearing duty cycles in 1C engine powered general-aviation aircraft need further definition. A study and experimental

program should characterize turbocharger and rotating group motion

during starting, running, and shutdown, and determine the greatest

likely bearing loads due to hard landings, taxiing, and combinations of flight maneuvers. A variety of aircraft missions should

be used to establish the time/temperature profile and load in situations that can cause contact between the rotor and bearings. The

resulting operating profiles can be combined with the improved

understanding of transient dynamics to establish more complete

requirements for foil design and coating materials. Evaluation of

the effects of mounting methods and location on the duty cycle

parameters would contribute to selection of an optimized installation by the engine and turbocharger designers. The operating profiles to be established are unique to aircraft applications, and at

present there are no programs to adequately define the applicable

duty cycle.

A program should be initiated to provide a comprehensive definition of the duty cycle of air bearings in 1C engine-powered

108
general-aviation aircraft, considering the frequency and duration

of all operating conditions over a variety of mission types.

5.2.3.1.4 Air Bearing Stability Analysis Technology

Garrett experience in application of compliant foil gas bearings covers a wide spectrum of rotor configurations, performance

requirements, and operating environments. In the course of performing each rotor dynamics analysis, the interaction of rotor and

bearings is examined using parametric techniques. Incomplete

knowledge of foil bearing dynamic properties and lack of an integrated model of the bearing and rotor prevents purely analytical

predictions of the susceptibility of the rotor to self-excited subsynchronous whirl. This whirl phenomenon is a common occurrence

during the development phase of rotor-foil bearing systems, especially when:

o Journals are closely spaced

o Significant vibration inputs exist that could excite resonant responses

o Large variations in temperature are present at the

bearings that cause significant changes in dimensions or

properties between various operating conditions.

If journal spacing can be reduced, the turbocharger is

shortened by the same amount and the engine installation is directly improved. In particular, externally induced vibration distinguishes turbocharger applications from the typical gas turbine

applications (for which most foil bearing development has been

done). Changes in power settings and altitude in aircraft applications cause heat input and cooling air-flow differences that significantly change bearing temperatures.

Several companies in the industry are devoting significant

effort to improving their analytical performance prediction capabilities for air bearings. These efforts are expected to assist in

development of optimized applications of air bearings in advanced

turbochargers.

5.2.3.2 Ceramics

Ceramic materials will contribute greatly to the success of

air bearings in turbochargers. Their light weight and low thermal

conductivity will reduce bearing loads and bearing temperatures,

significantly easing the demand for advancements in bearing technology. The light weight of a ceramic turbine wheel also allows

substantial weight reduction in the housings and structure that

must contain the products of any wheel burst. The reduction of

inertia improves transient response of the turbocharger; however,

this advantage is less important in general-aviation aircraft than

109
the direct weight saving in the turbocharger and mounting structures.

Current ceramic development programs have demonstrated adequate short-term behavior of ceramic materials in radial aerodynamic wheels. However, the scope of current programs does not

include establishing long-term behavior characteristics of these

materials. The results to date of these programs have shown the

need to improve the reliability and reproducibility of ceramic compositions. Knowledge of the long-term material behavior and

improved reliability is required to permit confident design of

rotating components. The ceramics initiative being pursued by NASA

and DOE recognizes the importance of long-term properties characterization and reliability design techniques, but will not address

these aspects until the late 1980's. This scheduling will delay

availability of the technology advances demanded by the generalaviation market.

A program should be initiated to determine the mechanical

properties (including long-term behavior) of currently available

sintered silicon-nitride ceramics. The results of these test bar

evaluations would be compared with the results obtained during

testing of actual ceramic rotors. A statistically iterative experimental procedure would be used to identify and reduce the variability in materials and processing, thus improving the reliability

of the ceramic rotor. This data would specifically benefit the

highly advanced technology engine programs within appropriate

schedule requirements. More advanced materials that will emerge

from the Ceramics Initiative will provide the increased temperature

capabilities necessary for turbochargers used on engines having

reduced cooling.

5.2.3.3 Variable Geometry

Use of variable geometry in the turbine nozzles will provide

the engine designer with control over the fuel/air ratio in the

engine. It also allows each operating condition to benefit from

turbomachinery operation closer to optimum. For four-stroke cycle

engines the benefits lie principally in controlling peak firing

pressure and thermal load on the engine. For two-stroke cycle

engines, scavenging pressures are improved and combustion efficiencies increased significantly. Optimization of variable geometry requires iteration of the turbine wheel design with nozzle

vane shape, location, motion, and engine and turbomachinery interactions, including scheduling or active optimization of turbine

nozzle position, fuel input, and such variables as fuel injection

or valve timing. No programs addressing technology of variablegeometry optimization for turbocharged engines are known to exist.

110
A program should be initiated to investigate interaction of

variable-geometry control with controllable functions on the advanced technology engines. The program should provide for alternate turbine designs; nozzle vanes and motion; aerodynamic analysis

of overall turbine performance; and simulation and confirmation

tests of turbine performance on the engines.

Ill
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6.0 CONCLUSIONS AND RECOMMENDATIONS

Use of turbochargers is necessary in advanced technology

engine development programs. In this vein, careful integration of

the turbocharger into each overall system must be accomplished.

Turbocharger development must lead or, at worst, pace that of the

engine in order to assure optimal system performance capability.

Turbocharger technology improvements can be developed and will

continue to draw on gas turbine technology as it advances. The

transfer of this technology will be motivated by the economic incentives tied to fuel prices, aircraft utilization, and market

penetration of the advanced engines.

Additional component technology demonstrations are required

before the advanced turbocharger can be committed to production

development. These include compressor and turbine aerodynamics,

advanced turbine materials, foil bearing dynamics and durability,

and lightweight structural acoustics and containment capability.

A large family of turbochargers is needed that will encompass

a range of flow capacities and pressure ratios that can meet the

diverse engine requirements for general-aviation aircraft. The desirable features of the ATDS are equally needed in nonaircraft applications, and the magnitude of this total market places great importance on the acquisition of these technologies. The evaluation

of the schedules for engine and turbocharger development indicates

that component technology demonstration programs should be initiated as soon as possible.

Turbocharger demonstration programs should be initiated that

support the engine technology demonstrators. Turbocharger design

and development should be carried out concurrently—and actually

integrated—with development of the engines for which they are intended.

Many of the turbocharger markets are characterized by pricequantity relationships that place great importance on volume production at the lowest possible cost. As a result, processes must

be used that do not have a development base in the gas turbine industry. The risks involved in advanced process development make it

difficult for any one company engaged in turbocharger production to

justify the total development effort needed for an advanced turbocharger design. Government sponsorship of the initial component

technology demonstration program will provide the impetus needed

for the turbocharger industry to undertake further development of

advanced turbocharger designs.

The relationships of the necessary program components were

shown in Figure 60. Once the enabling technology is made available, the producibility of engine and turbocharger designs will be

113
realistically evaluated by the industries involved. Provision of

information for this evaluation is the key activity by which NASA

can effectively support the nation's aircraft industry.

114
REFERENCES

1. Stuckas, K.J., "Advanced Technology Spark-Ignition Aircraft

Piston Engine Design Study," NASA CR-165162, Teledyne Continental Motors, Aircraft Products Div., November, 1980.

2. Brouwers, A.P, "150- and 300- kW Lightweight Diesel Aircraft

Engine Design Study," NASA CR-3260, Teledyne Continental

Motors, General Productions Div., April 1980.

3. "Design of Fluid Duct Bends for Low Pressure Loss," NASA Technical Brief No. 68-10395, October, 1968.

4. Miccio, J., "Variable Cycle Technology, Propulsion System Program," Garrett No: 21-3347(16), July 1983.

5. Badgley P., et al., "Advanced Stratified Charge Rotary Aircraft Engine Design Study", NASA CR-165398, Vol. 1, CurtissWright Corp., July, 1981.

6. Huggins, G., and Ellis, D., "Advanced General-Aviation Comparative Engine/Airframe Integration Study," NASA CR-165564,

September, 1981.

115
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APPENDIX I

CESSNA-PAWNEE SENSITIVITY STUDIES

(57 Pages)

117
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ATDS Airplane Sensitivity Charts

(Single Engine Airplane)

ORDINATE ABSCISSA

Takeoff Distance Versus

Maximum Takeoff Weight Versus

Rate of Climb at Sea Level Versus

Rate of Climb at 25,000 Ft Versus

Minimum Weight Required Versus

BEW Versus

Cruise Speed at 25,000 Ft Versus

Range Versus

Minimum Weight Required Versus

BEW Versus

Range Versus

Maximum TOGW Versus

Takeoff Distance Versus

Rate of Climb at Sea Level Versus

Rate of Climb at 25,000 Ft Versus

Cruise Speed at 25,000 Ft Versus

Minimum Weight Required Versus

BEW ' Versus

Takeoff Distance Versus

Rate of Climb at Sea Level Versus

Rate of Climb at 25,000 Ft Versus

Cruise Speed at 25,000 Ft Versus

Stall Speed Versus

Block Fuel Versus

Block Fuel Versus

Block Fuel Versus

Direct Operating Costs Versus

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Specific Fuel Consumption

Specific Fuel Consumption

Specific Fuel Consumption

AEngine Weight

ADrag

ADrag

ADrag

ADrag

ADrag

ADrag

TOGW

TOGW

TOGW

TOGW

TOGW

Specific Fuel Consumption

AEngine Weight

ADrag

AEngine Weight

119
120
J444U >..\^iii'i:~ i-rrn-.n

121
122
123
-WEIGHT-:

PAD RAN

124
125
126
127
128
129
130
131
132
133
QSSIAJlIIPM

134
135
GARRETT AtrVANCI

136
LOF. iGROSB

i.

. SENSITIV „.! ..__4

137
138
139
140
141
EFFECT OF DRAG, ON

FOK A CQNSfTANI, ! !i2ia0! LBSi 703

i RANGE

142
143
ATDS Airplane Sensitivity Charts

(Twin-Engine Airplane)

ORDINATE ABSCISSA

Takeoff Distance Versus

Maximum Takeoff Weight Versus

Rate of Climb at Sea Level Versus

Rate of Climb at 25,000 Ft Versus

Single Engine ROC at Sea Level Versus

Single Engine ROC at 5000 Ft Versus

Cruise Speed at 25,000 Ft Versus

Minimum Weight Required Versus

BEW Versus

Range Versus

Minimum Weight Required Versus

BEW Versus

Range Versus

Maximum TOGW Versus

Rate of Climb at Sea Level Versus

Rate of, Climb at 25,000 Ft Versus

Single Engine ROC at Sea Level Versus

Single Engine ROC at 5000-Ft Versus

Minimum Weight Required Versus

BEW Versus

Takeoff Distance Versus

Cruise Speed at 25,000 Ft Versus

Takeoff Distance Versus

Rate of Climb at Sea Level Versus

Rate of Climb at 25,000 Ft Versus

Single Engine ROC at Sea Level Versus

Single Engine ROC at 5000 Ft Versus

Cruise Speed at 25,000 Ft Versus

Stall Speed i Versus

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Brake Horsepower

Specific Fuel Consumption

Specific Fuel Consumption

Specific Fuel Consumption

AEngine Weight

ADrag

ADrag

ADrag

ADrag

ADrag

ADrag

ADrag

ADrag

Takeoff Gross Weight

Takeoff Gross Weight

Takeoff Gross Weight

Takeoff Gross Weight

Takeoff Gross Weight

Takeoff Gross Weight

Takeoff Gross Weight

144
ATDS Airplane Sensitivity Charts

(Twin-Engine Airplane Contd)

ORDINATE ABSCISSA

Block Fuel

Block Fuel

Block Fuel

Direct Operating Costs

Versus Specific Fuel Consumption

Versus AEngine Weight

Versus ADrag

Versus AEngine Weight

145
146
147
148
149
150
151
152
153
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171
172
173
EFFECT OF DRAG ON BLOCK FUEL

A PONSITANT. RELOAD RANGE

1400: LBSl 5013 NM

[TWIN ENGIN^ AIRPLANE

LTURBQCHARGER

; SEN^IT^ViTYj STUDY

ECJUIVALJENt F|_AT P.A^E

.4. ..j ^ I ,.J L- .L -

174
175
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APPENDIX II

ATDS AIRPLANE COST EQUATIONS

(7 Pages)

177
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T.TI r

pnrpAaro av

CMFCKEO BY

CESSNA AIRCRAFT CO

pATF

"ATE. ,._

Cessna.

flFV'S'ON

PAGE

REPORT I

MODEL

PAWNEE DIVISION WICHITA KANSAS

PRICING

The selling price (for purposes of estimating the value of new models) is

considered to be the sum of 3 components1. The base price - that which is manufactured by Cessna (also accounts for

minor purchased parts);

2. The poverplant contribution to the price - includes engine, propeller, and

governor; and

3. The part of the price attributable to optional equipment (of which avionics

account for about 50Z of the value).

The base price is computed by

Base Price

where

a(WEp)

b

(VMAX)

C

(Sw)

d

(CW)£

a - 7.268188MO"4

b « 1.06942

c - 1.05600

d • .65289

e - .22723

This is an emperical equation generated by applying a least squares regression

analysis to the existing 1981 Cessna fleet. (The constant is valid only for

pressurized aircraft.) The factors in the above equation are;

Wgp - Dry Empty Weight minus Weight of Powerplant (engine, governor, propeller)

- Maximum Speed (knots)

- Wing Area (ft2)

GW - Takeoff Gross Weight (Ibst

Values of these factors for each plane and the resulting base price are shown in

Table II-l.

The price Increment attributable to the p owe rp lane Is very difficult to estimate

since it involves estimating the OEM cost not only of the engine but the propeller

as well. Then the markup applied to the powerplant of that individual airplane

must be considered. When the engine is not only new but is of a type not previously

used in production aircraft the Job becomes virtually impossible to do w.'th any

degree of certainty. Based on existing products and the fact that these are

larger engines than now used in most Cessna products an increment of $35,000 per

engine was chosen for all engines in the study.

The price increment attributable to optional equipment varies widely depending on

the equipment chosen. Airplanes of this category can normally be expected to be

equipped with radar, alrconditioning, and a de-icing package In addition to the

usual avionics and interiors. Based on present Cessna prices for equipped planes

that include these features an increment of $48,000 was chosen for the single

engine planes and $82,000 for the twin engine planes.

The total estimated selling price is shown in Table AII-1.

179
TITLE

PREPARED BY .

CHECKED BY

DATE.

DATE_

Cessna.

PAGE

REPORT NO.

MODEL

CESSNA AIRCRAFT CO PAWNEE DIVISION WICHITA KANSAS

DIRECT OPERATING COST

The components considered in generating DOC and an outline of how they are

generated are shown in Table AII-2. The following discussion covers how these

values were calculated in this study.

For the engines considered herein no data was available to accurately estimate

the cost of either the engine periodic maintenance nor the reserve for engine

overhaul. Values of $9/engine and $8/engine were chosen which are in line with

the current values for the larger TSIO-520's.

Propeller overhaul, airframe maintenance and systems maintenance are calculated

as shown. The factors used in the emperical curve fits for the last two

components are shown in Table AII-3.

Hull and liability insurance rates are found in Table AII-4. The rate for the

single engine aircraft Is .0160 and for the twin is .0150. A utilization of

500 hours/year was assumed for both.

Fuel cost was calculated assuming' a cost of $1.70 per gal, a density of 6 Ib/gal,

a BHP in cruise of 250, and the SFC's shown in Table AII-3.

The oil consumption was assumed to be .1 GPH per hour per engine at a cost of

$6/gallon.

Depreciation is based on the total aircraft selling price. It is, therefore,

not Independent of the assumed price for the engine.

Reserves for avionics was based on assuming that avionics accounts for 507. of

the optional equipment price.

Table AII-5 shows the values of each component of the DOC and the total for

each plane.

180
TITLE x~1 \.

 r*GE

PREPARED BY OATC VCSSIla Ces * REPORT NO .

CHECKED BY DATE REVISION MODEL

CESSNA AlRCRATT CO PAWNEE DIVISION WICHITA KANSAS

TABLE AII-1

FACTORS USED IN COMPUTING AIRCRAFT PRICE

SINGLE ENGINE TWIN ENGINE

TS10-550 RC2-47 RC2-32 TSIO-550 RC2-47 RC2-32

WEp 1910.4 1630.0 1615.6 2314.5 2084 1 2038 3

VH^V 209 1 249.5 252.8 243.3 275.4 279.5

Sw 177 150 142.5 170 162.5 154.5

GW 4550 3900 3775 6625 5650 5375

Base Price 131677 116053 113171 203799 191644 182601

Powerplant Price 35000 35000 35000 70000 70000 70000

Optional Equip Price 48000 48000 48000 82000 82000 82000

Total 214677 199053 196171 355799 343644 334601

181
TABLE All-2

DIRECT OPERATING COSTS FOR GENERAL AVIATION AIRCRAFT

- 1981 Estimate

1) ENGINE PERIODIC MAINTENANCE

Use past experience (i.e. similar engine/airframe combination) or

engine manufacturer's estimate.

Otherwise use:

Number of labor hours for 100 hour inspection x labor rate

100

Then double this answer to account for parts.

Labor rate right now runs $20/hour S/E

$25/hour M/E

$30/hour Turboprops

*Turboprops must be considered under a different formula. Instead of

being inspected every hundred hours, they undergo a aeries of Hot

Section Inspections during the overhaul period. These are usually

of considerably greater time than 100 hours. For some engines the

work scheduled for each HSI is different as the time from last overhaul increases.

/filters. Igniters + \

E (cost of labor + cost of parts) for HSI's + misc. \labor not included In HSI

TBO

2) RESERVES FOR ENGINE OVERHAUL

The assumption (conservative) is made that every other overhaul will

require, instead of an overhaul, a remanufactured engine. Therefore:

(overhaul cost + cost of remanufactured engine)/2

TBO

*For Turboprops:

allowance for premature removal** of

engine and engine accessories*** and

overhaul cost (labor -4- parts) + engine components 

TBO

•Information not given to Curtiss-Vrlght by phone in January

** Allowance for engine removal amounts to 1/5 to 1/2 of engine overhaul cost

***Starter generator, etc.

****Turbines, nozzles, etc.

182
TABLE AII-2 (CONTINUED)

3) PROPELLER OVERHAUL

Propeller

Fixed Pitch fl£t

S/E Controllable < HPSE .60

VCenturion class .82 >

M/E controllable (per propeller) .90

4) AIRFRAME MAINTENANCE

This number Is based on a parametric fie of the available data.

DOC = 1.472 + .000534 TOGW - .000373 BHP (Total

+2.774 (only for twins) + 1.878 (only If pressurized)

5) INSURANCE (HULL + LIABILITY)

See attached charts

6) FUEL COST

DOC = price x gal per (present price of AV gas is calculated at $1.30 gal)

ga1 hour

7) OIL COST

DOC ° price x GPH used (present price of oil is $6/gal which also accounts

gal for cost of oil filter)

or alternately use

DOC = actual price x CPH used + cost of filter

gal (including 0 hrs between filter change

both consumed

and lost during

change)

8) DEPRECIATION

- Total equipped airplane price i.e. discounted to zero residual in

7.5 x utilization rate/year 7.5 years

9) RESERVES FOR AVIONICS

10% of total aviontc package (standard + optional)

1000

10) RESERVES FOR SYSTEMS MAINTENANCE

DOC - -.513 + 000803 TOGW + 1.109 (if pressurized)

Again this is a parametric fit of available data

183
TABLE All-3

FACTORS USED IN COMPUTING DIRECT OPERATING COST

SINGLE ENGINE TWIN ENGINE

TOGW

BHP 707

Price

Optional Equip Price

SFC

TSIO-550

4550

350

2U677

48000

.446

RC2-47

3900

320

199053

48000

.371

RC2-32

3775

320

196171

48000

.355

TSIO-550

6675

700

355799

82000

.446

RC2-47

•5650

640

343644

82000

.371

RC2-32

5375

640

334601

82000

.355

ItmipgilMOl Anjlf $U Plan

184
INSURANCE RATES AFPLIUVBLfc TO 198i CESSNA MODELS

October, 1980

TABLE AII-4

Pleasure & Business Rates For Well-Qualified Pilots:

Hull Value

$15,000

25,000

40,000

60,000

100,000

150,000

150,000

300,000

500,000

750,000

1 Mil

- 24,999

- 39,999

- 59,999

- 99,999

- 149,000

- 200,000

- 299,999

- 499,999

- 750,000

- 1 Mil.

- 1.5 Mil

Single Engine Rate

3.00Z

2.75

2.50

2.00

1.75

1.60

Hulti-Engine Rate

1.75Z

1.50

1.35

1.10

1.00

Legal Liability limit of $5,000,000, combined cingle limit.

Seats

A

5

6

7

8

9

10

11

Annual Premium

$ 575

675

725

825

975

1,075

1,175

1,250

185
OREPAREC tf"

CnECI LL BY_

DATE.

DATE.

CAGC „

v^essna.

BEViS'ON __

CESSNA A>RC»A*T CO PAWNEE DIVISION Wi,_H>T«

TABLE All-5

DIRECT OPERATING COST

SINGLE ENGINE

Engine Maintenance

Engine Overhaul

Propeller

Airframe

Insurance Liab

Insurance Hull

Fuel

Oil

Dep

Res for Avionics

Systems

Total

TSIO-550

9.00

8.00

.82

5.65

1.45

6.87

31.59

.60

57.25

2.40

4 25

127 88

RC2-47

9.00

8.00

.82

5.31

1.45

6.37

26.28

.60

53.08

2.40

3 73

117.04

RC2-32

9.00

8.00

.82

5.25

1 45

6.26

25.15

.60

52.31

2.40

3.63

114.89

TVIl> ENGINE

TSIO-550

18.00

16.00

.90

9.92

1.65

10 67

63.18

1.20

94.88

4.10

5.92

226.42

RC2-47

18.00

16.00

.90

8.90

1 65

10.31

52 56

1.20

91.64

4.10

5.13

210.39

RC7-32

18 00

16.00

.90

8.76

1.65

10.44

50 29

1.20

89.23

4.10

UJL_

205 08

186