What is the main focus of the NASA technical memorandum?

The memorandum focuses on the overview of aerospace gas turbine technology and its relevance to the development of automotive gas turbine engines.

What are the key technology areas discussed in the document?

The key technology areas include performance prediction techniques, compressor and turbine design, combustor technology, and materials technology.

What are the requirements for automotive gas turbine propulsion systems?

Requirements include efficient operation over a duty cycle, minimal impact of small-scale effects on efficiency, competitive manufacturing costs, acceptable acceleration, environmental acceptability, reliability, safety, and acceptable volume and weight.

What types of configurations are available for automotive gas turbine engines?

Configurations include single shaft, multiple shaft (free turbine), and three-shaft engines, each with different trade-offs in cost, performance, and transmission type.

What are the future performance targets for automotive gas turbines?

Future targets include increases in turbine inlet temperatures to the 1645 K (2500°F) range and improved part power performance.

An overview of aerospace gas turbine technology of relevance to the development of the automotive gas turbine engine

Document

NASA T M -7?849

N A S A T E C H N I C A L

M E M O R A N D U M

AN OVERVIEW OF AEROSPACE GAS TURBINE TECHNOLOGY

OF RELIVANCE TO THE DEVELOPMENT OF THE

AUTOMOTIVE GAS TURBINE ENGINE

by D. G. Evans and T. J . Miller

Lewis Research Center

Cleveland, Ohio 44135

TECHNICAL PAPER to be presented at the

Annual Meetirg of the Society of Automotive Engineers

Detroit, Michigan, February 27 - March 3, 1978

AN OVERVIEW OF AEROSPACE GAS TURBINE TECHNOLOGY OF RELIVANCE TO TlIE DEVELOPMENT OF THE AUTOMOTIVE GAS TURBINE ENGINE by D . C. Evans and T. J. Miller National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 44135 ABSTRACT The NASA-Lewis Research Center (LeRC) has conducted, and has sponsored with i n d u s t r y and u n i v e r s i t i e s , e x t e n s i v e r e s e a r c h i n t o many of t h e technology areas r e l a t e d t o g a s t u r b i n e propulsion systems. This aerospace-related technology has been developed a t both t h e com- ponent and systems l e v e l . and may have s i g n i f - i c a n t p o t e n t i a l f o r a p p l i c a t i o n to t h e auto- motive g a s t u r b i n e engine. This paper summarizes t h i s technology and lists t h e as- sociated references. The technology areas included are: system s t e a d y - s t a t e and tran- s i e n t performance p r e d i c t i o n techniques, compresscr and t u r b i n e design and performance p r e d i c t i o n programs and e f f e c t s of geometry, combustor technology and advanced concepts, and ceramic c o a t i n g s and materials technology.

Evans and ?!iller WE NATION HAS ESTABLISHED THE GOAL OF reducing its consumption of petroleum derived f u e l s . Since t h e t r a n s p o r t a t i o n s e c t o r con- sumes over SOX o r approximately 1.mtlol3 mega j o u l e s (18 QUADS) per year of t h e crude o i l used in t h e nation, it is a prime area f o r implementing conservation measures. For high- way v e h i c l e a p p l i c a t i o n s , t h e g a s t u r b i n e has been i d e n t i f i e d as a v i a b l e a l t e r n a t i v e pro- pulsion system (l)*. It has t h e p o t e n t i a l f o r improved f u e l economy over t h e present spark i g n i t i o n i n t e r n a l c o m b s t i o n engine while meeting f u t u r e emission standards, and has t h e inherent a b i l i t y t o o p e r a t e not only on petroleum derived f u e l s b u t a l s o on f u e l s derived from o t h e r sources.

Most of t h e technology base f o r t h e g a s t u r b i n e has evolved from over 30 years of ef- f o r t by t h e m i l i t a r y , NASA, and t h e industry t o meet t h e m i l i t a r y and commercial needs f o r Improved a i r c r a f t propulsion systems. For NASA, t h e s e technology a c t i v i t i e s have been conducted both in-house and through c o n t r a c t s and g r a n t s with Industry and u n i v e r s i t i e s by its lead c e n t e r i n propulsion technology, t h e L e w i s Research Center (LeRC). Because t h e automotive gas t u r b i n e w i l l b u i l d upon t h e aerospace base, i t is t h e purpose of t h i s paper t o d e s c r i b e aerospace r e l a t e d technology developed i n programs conducted by LeRC t h a t may help t h e gas t u r b i n e powered automotive v e h i c l e t o meet o r exceed t h e goal t h a t has been set. Although w e l l known t o t h e auto- mobile Industry, t h e general s t a t u s and char- acter of t h e automotive gas t u r b i n e and its d i f f e r e n c e s from t h e a i r c r a f t a p p l i c a t i o n rill be o u t l i n e d 70 provide a b a s i s f o r discussion.

I n s t a l l i n g a gas t u r h i n e i n an automotive v e h i c l e is not a new idea. Programs have been conducted by many firms over approximately t h e l a s t 30 y e a r s a s reported i n (1). The follow- ing i t 3 an excerpt from t h e reference: "Rover demonstrated t h e world's f i r s t gas t u r b i n e passenger c a r i n 1950. The Chrysler Corporation has also heen s e r i o u s l y pursuing development of passenger c a r gas t u r b i n e s s i n c e 1950, and tried out a proto- type 50-car f l e e t on t h e public i n t h e 1964- 1966 time period. General Motors announced its first automotive t u r b i n e i n 1954 and un- Evans and Miller v e i l e d its experimental F i r e b i r d I1 regener- *Numbers I n parentheses designate Ref- ences a t t h e end of paper.

a t i v e GTE sedan i n 1955. Ford, Williams Re- search, Volkswagen, and o t h e r s have l i k e w i s e embarked upon t u r b i n e passenger car develop- ment programs. None of t h e s e programs have, as y e t , r e s u l t e d i n 8 mass-production v e r s i o n of t h e automctive gas turbine."

U . S. manufacturers have been developing g a s t u r b i n e propulsion systems f o r many a p p l i - c a t i o n s . M o s t r e c e n t l y , t h e Chrysler Corpora- t i o n has been working with t h e government t o develop t h e g a s t u r b i n e engine and car sham i n Fig. I . S i m i l a r l y , D e t r o i t Diesel A l l i s o n (DDA) Division of General Motors is conducting a performance improvement program w i t h its heavy duty gas t u r b i n e engine i n s t a l l e d i n Greyhound buses and t r a c t o r trailer trucks.

A cutaway view of t h e DDA heavy duty engine is shown i n Fig. 2. The Ford Motor Co. con- ducted a p i l o t production program i n t h e e a r l y 1970's, i n s t a l l i n g t h e i r heavy duty g a s t u r - b i n e engine i n t r u c k and marine a p p l i c a t i o n s .

The engine w a s scheduled f o r productlo,? i n t h e mid-l970's, but regenerator problem d e f e r r e d t h e decision.

Each of these programs has contributed toward def ining t h e c u r r e n t s t a t e - o f - t h e - a r t of t h e automotive g a s t u r b i n e engine. Cur- r e n t l y these gas t u r b i n e engines are o p e r a t i n g a t t u r b i n e inlet temperatures of approximately 1310 K (1900O F), and are b a s i c a l l y con- s t r u c t e d of m e t a l l i c components (i.e., t u r - bine, combustor, e t c . ) w i t h t h e exception of ceramics being used i n some regenerator a p p l i - cat ions.

I n response to t h e n a t i o n a l goal t o re- duce t h e consumption of f u e l s derived from crude o i l and t o meet t h e f e d e r a l emission standards, t h e Division of Transportation En- ergy Conservation of t h e Department of Energy (DOE) is sponsoring a program t o develop arzd demonstrclte f u e l e f f i c i e n t g a s t u r b i n e powered highway v e h i c l e s . This program is being c a f r i e d o u t by t h e Government with t h e major U. S. automobile manufacturers. Pro- gram madagernent is under t h e d i r e c t i o n of DOE.

"Project management" r e s p o n s i b i l i t y f o r t h e v a r i o u s elements of t h e program has been d e l - egated t o LeRC.

Although t h e DOE program considers t h e e n t i r e propulsion system (engine and trans- Evans and Miller mission) t h i s paper will address only t h e en- g i n e r e l a t e d technology. S p e c i f i c a l l y , t h e a r e a s t o be discussed i n t h e paper a r e : steady s t a t e and t r a n s i e n t systems a n a l y s i s , compressors and turbine3, combustors, and ma- terials and coatings. Papers covering seals, bearings, and instrumentation technology de- veloped i n t h e LeRC managed programs are be- ing presented i n s e p a r a t e papers during t h i s ses s ion.

The following i n d i v i d u a l s contributed t o t h i s paper: John Klann and John Zeller, Sys- tem Performance Analysis Tech-iiques; Robert Wong and Arthur Glassman, Compressors and S r b i n e s ; David Anderson, Albert Juhasz, Donald Schultz, and Richard Niedzwiecki, Com- bustors; and Stanley Levine, Curt L i e b e r t , and Richard Ashbrook, Coatings and Materials.

DESCRIPTION OF THE AUTOMOTIVE GAS TURBINE ENGINE The gas t u r b i n e o p e r a t e s on a Brayton thermodynamic cycle. The design e f f i c i e n c y of t h i s c y c l e is b a s i c a l l y c o n t r o l l e d by t h e o p e r a t i n g temperatures of t h e c y c l e and t h e e f f i c i e n c y of t h e components used t o mech- a n i z e t h e cycle. Typically, a gas t u r b i n e propulsion system is w e l l s u i t e d f o r applica- t i o n s where t h e engine spends a l a r g e amount of time o p e r a t i n g a t or near 100% of its de- s i g n speed and design operating temperature (i.e., engines f o r subsonic a i r c r a f t , a u x i l - i a r y power u n i t s , e t c . ) . The automotive ap- p l i c a t i o n , however, imposes a d d i t i o n a l re- quirements on gas t u r b i n e propulsion system which must be met i f t h e system is t o be a v i a b l e a l t e r n a t i v e . The b a s i c requirements are : h e 1 e f f i c i e n t o p e r a t i o n over a duty c y c l e t h a t r e q u i r e s s i g n i f i c a n t o p e r a t i o n a t off-design p a r t speed and p a r t power condi- t ions.

- Minimum impact of small-scale e f f e c t s

on conponent e f f i c i e n c y .

* Competitive manufacturing, i n i t i a l and l i f e c y c l e c o s t s .

- Acceptable a c c e l e r a t i o n .

* Env ironmen t a l ac c ep ta b i l i t y (emissi c , ns and n o i s e ) .

- Acceptable r e l i a b i l i t y and s a f e t y .

* Acceptable volume and weight.

hl t i - f uel capa b i 1 i t y .

The gas t u r b i n e s y s t e m has shown a good po- Evans and Miller t e n t i a l f o r meeting these automotive require- ments.

I n t h e i d e a l Brayton c y c l e , t h e working f l u i d ( a i r ) is compressed from ambien; condi- t i o n s . Energy is added t o t h e working f l u i d i n t h e farm of h e a t , and energy is e x t r a e t e d from t h e working f l u i d t n an expansion pro- cess. The c y c l e can b e operated w i t h a num- ber of v a r i a t i o n s . For an automotive a p p l i - c a t i o n , however, a n open regenerated cycle using r o t a t i n g turbomachinery is commonly of The components used i n t h i s most i n t e r e s t .

caw are a compressor, combustor, h e a t ex- changer, and t u r b i n e . Multiple o p t i o n s eldsts i n t h e type of components used t o mechanize t h e c y c l e and t h e c o n f i g u r a t i o n i n which t h e s e components are arranged. The primary o p t i o n s a v a i l a b l e are:

- Compressor - a x i a l o r c e n t r i f u g a l

- s i n g l e o r m u l t i p l e s t a g e s

- f i x e d o r v a r i a b l e geometry

Heat Exchanger - s t a t i o n a r y recuptera-

t o r

- r o t a t i n g regenerator

* Turbine - r a d i a l o r a x i a l

- s i n g l e c: m u l t i p l e s t a g e s

- f i x e d o r v a r i a b l e geometry

* Engine S h a f t Arrangements - single o r

mu1 t i p l e

Engine Configurations - The arrangement

of engine components i n t o s i n g l e o r m u l t i p l e s h a f t configurations is a complex problem with trade-offs required between c o s t , per- formance, transmission type, and engine s i z e .

I n t h e s i n g l e s h a f t engine, t h e compressor and power t u r b i n e are mounted on a common s h a f t and t h e r e f o r e must o p e r a t e a t t h e same speed. This common s h s f t is then d i r e c t l y coupled t o a transmission. While t h i s engine configuration is r e l a t i v e l y simple, it re- q u i r e s a r a t h e r complicated continuously v a r i a b l e transmission (CVT) t o o p e r a t e over t h e speed range required f o r t h e automotive a p p l i c a t i o n . A s i n g l e s h a f t c o n f i g u r a t i o n that was t h e s u b j e c t of study i n ( 2 ) is shown schematically i n Fig. 3.

I n t h e m u l t i p l e s h a f t configuration, re- f e r r e d t o as a f r e e t u r b i n e engine, t h e com- pressor and compressor d r i v e t u r b i n e a r e mounted on a common s h a f t . This p o r t i o n of t h e assembly including t h e combustor, o f t e n r e f e r r e d t o as t h e g a s i f i e r assembly, is f o l - lowed by t h e pcwer t u r b i n e ( f r e e turbine) Evans and Miller which is mounted on a second s h a f t . I n t h i s configuration t h e g a s i f i e r t u r b i n e is coupled aerodynamically t o t h e power t u r b i n e and t h e power turbine is coupled t o t h e transmission.

This configuration can be operated over t h e automotive duty c y c l e using a conventional three-speed transmission. A schematic of a two s h a f t engine (Chrysler) is presented i n Fig. 4. The state p r o p e r t i e s shown i n t h e f i g u r e are r e p r e s e n t a t i v e of t h e operating conditions f o r t h i s engine. Turbine i n l e t temperatures are l i m i t e d t o approximately 1325 K (1925' F), t h e compressor pressure ra- t i o is approximately 4:1, and t h e regenerator o p e r a t e s t o a temperature of approximately 1017 K (1370' F) .

Another m u l t i p l e s h a f t g a s t u r b i n e con- f i g u r a t i o n is t h e three-shaft engine discussed i n (3) and shown schematically i n Fig. 5.

This configurations key f e a t u r e is t h e u s e of a t h i r d t u r b i n e on a s e p a r a t e s h a f t that can assist i n d r i v i n g t h e compressor, t h e v e h i c l e , and t h e a c c e s s o r i e s . The use of t h e t h i r d t u r b i n e according t o t h e r e f e r e n c e permits c l o s e t a i l o r i n g of t h e engines performance and operating characteristics t o t h e require- ments of t h e automotive duty cycle. While t h e t h i r d s h a f t adds complexfty t o t h e engine, t h e transmission requireme3ts can be s a t i s f i e d with a r e l a t i v e l y simple p l a n e t a r y gear box.

There are many v a r i a t i o n s p o s s i b l e i n t h e gas t u r b i n e engine. However, t h e t h r e e b r i e f l y described here, t h e s i n g l e , two- and three-shaf t c o n f i g u r a t i o n s a r e r e p r e s e n t a t i v e of t h e primary v a r i a t i o n s .

Future Requirements . Today with t h e heavy emphasis on f u e l economy, t h e primary performance t a r g e t s f o r t h e advanced automo- t i v e gas t u r b i n e can be summarized a s : (1) Increases i n t u r b i n e i n l e t tempera- t u r e t o t h e 1645 K (2500' F) range.

(2) Improved part power performance.

These perf orrnance f a c t o r s , along w i t h the manufacturing c o s t , must be considered i n t h e development of these engines. The high l e v e l nf t u r b i n e i n l e t temperature places require- ments on t h e hot s e c t i o n c o a p n e n t s well be- yond t h e technology i n c u r r e n t automotive gas t u r b i n e engines where these components a r e made of metal a l l o y s . Because component cool- ing is not c u r r e n t l y being considered because of !ts impact on c o s t , t h e higher temperature requirements w i l l r e q u i r e development of ce- ramic m a t e r i a l s f o r t h e s e components. A pre- liminary a n a l y s i s of t l p o t e n t i a l b e n e f i t s Evans and Miller and problems a s s o c i a t e d with the introduction of ceramics t o an advanced gas t u r b i n e engine is presented i n ( 4 ) .

The second t a r g e t of good p a r t power per- formance w i l l r e q u i r e development of t h e o f f - design performance c a p a b i l i t y of t h e turbo- machinery. One of t h e f a c t o r s being consid- ered f o r increased p a r t power performance is t h e a d d i t i o n of v a r i a b l e geometry t o t h e t u r - bomachinery. Another f a c t o r required is higher o p e r a t i n g temperature heat exchangers, which is another area where high temperature ceramic materials are being considered.

APPLICABLE AEROSPACE TECHNOLOGY T h e following s e c t i o n s present a summary of t h e t u r b i n e engine technology t h a t has been developed through aerospace-related pro- grams t h a t have t h e p o t e n t i a l t o help meet t h e goals and requirements f o r t h e automotive gas t u r b i n e engine.

SYSTEM PERFORMANCE ANALYSIS TECHNIQUES -

Both steady-state and t r a n s i e n t performance p r e d i c t i o n s programs and a n a l y s i s techniques have been developed and used extensively a t LeRC as an e f f e c t i v e t o o l t o a n a l y t i c a l l y de- termine o r e v a l u a t e t h e performance charac- teristics of varicjcls gas t u r b i n e systems.

They are described a s follow?:

Steady-State Performance Program - As

discussed previously, there a r e many p o s s i b l e gas t u r b i n e c o n f i g u r a t i o n s t h a t could be b e n e f i c i a l i n an automotive a p p l i c a t i o n . I n o r d e r t o idencify these and determine t h e most promising c o n f i g u r a t i o n s , each one must be evaluated on its own merit. One of t h e primary c o n s i d e r a t i o n s is t h e f u e l economy p o t e n t i a l when o p e r a t i n g a t design and o f f - design c o n d i t i o n s over a d r i v i n g o r duty cycle. To perform t h i s t a s k , a generalized computer code which was developed f o r a n a l y s i s o f a i r c r a f t gas t u r b i n e s has been modified f o r automotive use. The code is r e f e r r e d t o as t h e NAW/NASA Engine Program (NNEP), and is Jcscribed i n (5). I t has evolved from a series of previous Government and i n d u s t r y - developed a i r c r a f t gas t u r b i n e computer codes i n t o a v e r s a t i l e and advanced assessment t o o l .

The c u r r e n t version of NNEP i s t h e r e s u l t of a j o i n t e f f o r t of :he Naval A i r Development Center and LeRC. and is a nonproprietary code t h a t is f r e e l y a v a i l a b l e t o t h e i n d u s t r y . Evans and Miller W i t h some modifications and a d d i t i o n s , NNEP is being applied a t L e R C t 3 automotive gas- t u r b i n e a n a l y s i s .

Since NNEP is an outgrowth of many pre- vious gas t u r b i n e computer codes, It ie a highly s o p h i s t i c a t e d and generalized t o o l .

Through d a t a Input statements, t h e user b u i l d s t h e engine configuration by numbering and nam- i n g t h e type of components, d e f i n i n g t h e com- ponent performance c h a r a c t e r i s t i c s , i n d i c a t i n g component interconnections, and d e f i n i n g t h e engine design conditions. Each i n i t i a l calcu- l a t i o n i n NNEP is f o r an engine design p o i n t condition. Furthermore, NNEP allows more than one engine arrangement t o be s p e c i f i e d a t a time. Thus, while running t h e code, flow paths o r mechanical arrangements can b e switched t o simulate v a r i a b l e c y c l e engines ( f o r a i r c r a f t engin- a n a l y s i s ) o r power trans- f e r among s h a f t s ( f o r automotive engine anal- y s i s ) .

The types of flow and mechanical compo- nent a v a i l a b l e i n M N E P include: i n l e t s , ducts, compressors, combustors, t u r b i n e s , heat ex- changers, water i n j e c t o r s , flow s p l i t t e r s , flow mixers, nozzles, s h a f t s , and loads. Per- formance of each of t h e s e components is han- dled i n s e p a r a t e NNEP subroutines. The com- p r e s s o r , duct, and t u r b i n e subroutines allow t h e s p e c i f i c a t i o n of engine bleed flows. The t u r b i a e subroutine a l s o has b u i l t - i n cooling options.

Additional NNEP i n p u t s include c o n t r o l s a n d j a r optimization v a r i a b l e s . Controls are used t n balance ( o r unbalance) engine flow condrtlons and/or s p e c i f y t h e d e s i r e d er i n e operating conditions. The optimization vari- a b l e s can be assigned t o maximize or minimize engine performance parameters a t design o r off-design o p e r a t i n g conditions. The code is l i m i t e d t o a maximum of 60 t o t a l components, c o n t r o l s , and optimization v a r i a b l e s . As a r e s u l t , n e a r l y any t y p e of gts-turbine con- f i g u r a t i o n and its o p e r a t i o n a l c o n t r o l s can be synthesized i n NNEP.

Off-design component performance is de- scribed i n NNEP through t h e use of input ta- b l e s o r maps. The t a b l e s a r e normaiized t o the component design conditions. NNEP can accommodate three-dimensional component maps.

For example, a configuration w i t h both v a r i - a b l e t u r b i n e otid compressor geoaetry could be syntheoized, and its v a r i a b l e geometry s e t t i n g could be optimized f o r f u e l consumption a t Evans and Miller each engine o p e r a t i n g speed.

Recent a d d i t i o n s t o t h e NNEP are a com- ponent s i z i n g and a weight subroutine. These 7 subroutines, constructed from a i r c r a f t design procedures, have demonstrated accuracies within 210 percent. Unfortunately, these routines are not ccrrently applicable to automotive gas turbine analysis .

There have been two additions to NNEP’s basic capabilities to facilitate automotive gas turbine analysis. These additions include subroutines for generating preliminary com- pressor and turbine (turbomachinerp) design characteristics, and for outputting engine performance maps for use with a Driving-Cycle Analysis computer code. The preliminary tur- bomachinery design subroutines are based on synopses of models for radial-flow compressors, radial-flow turbines, and sxial-flow turbines.

Use of these subroutines in NNEP produces design-point component efficiency predictions, rotor dimensions, and mean-section velocity diagrams. Thus, tu:bomachinery design trade- offs and their effects on :ngine performance can be investigated directly rather than on a parametric basis.

The subroutine which has been added for outputting engine performance maps allows two transmission options which affect the form of outputting. Either a continuously-variable speed-ratio (CVT) or three-speed transmission may be specified. With the CVT option, the engine performnnce map is outputted as a single operating line. The three-speed option results in a complete matrix of engine oper- ating conditions.

One example of the intended use for NNEP in automotive analysis is a current effort aimed at making a preliminary definition of an Advanced Gas Turbine Powertrain. The fuel economy potentials of a matrix of Advanced Gas Turbine configurations is being screened with NNEP in combination with the Driving Cycle Code. Multiple-shaft arrangements are being studied with both the CVT and three- speed transmissions. Single-shaft configura- tions are being studied with a CVT.

For each configuration and set of oper- ating constraints, NNEP is being used to ex- amine engine performance as a function of de- sign point compressor pressure ratio. Results from NNEP are then being used in the Driving Cycle code to find the best design conditions for v.?hicle fuei economy. Comparisons among Evans and Miller the screening study configurations will be used to identify promising arrangements to be carried into more detailed conceptual design studies.

Transient Performance Prediction and

Analysis - In addition to steady-state turbine

engine performance, tranaient engine perfor- mance is of importance to overall propulsive system operation. Transient conditions exist when a change in output power is deliberately requested or when external engine conditions change. Whether a turbine engine is being used in an aircraft applicatioc or an automo- tive application. a knowledge of its transient performance is necessary. The manner in which a turbine engine will perform transiently is today dictated by a closed loop (feedback) control system. The controller must regulate performance at an output power condition and be able to quickly and predictably take it from one operating point to another. It must do this while avoiding conditions of over- speed, turbine overtemperature, combustor blow-out, etc.

To understand the dynamic interactions of aircraft turbine engine system components, in- cluding the controller, and to assist in de- signing a control strategy which will guar- aritee some specified performance, computer- ized simulations of this complex system are being used. These simulations are detailed accurate analytical representations or models of the engine systems. These simulations have been accomplished using both digital and hybrid (analog and digital) computers. Digi- tal simulations of the transient and steady- state characteristics of complex systems are useful when the simulation must be utilized by several different organizations. Universal software languages (FORTRAN and C S M P ) enable the simulation to operate on different types of computing equipment. Digital simulations of engines, however, do not normally run in real-time without major simplifications. Hy- brid engine simulations, however, can aperate in real time ( 6 to 8). The value of a hybrid simulation for the development of controllers for advanced technology turbine engines will be the subject of the next few paragraphs.

The trend in controllers for aircraft turbine engines is toward an electronic digi- tal computing device. This trend is due to the high computational requirements of control laws for advanced engines and the potential Evans and Miller cost benefits such a device may yield. A digital computer controller is a sampled-data Controller (control update inputs at distinct intervals of time after sampling new perfor- mance information)and relies on a s t o r e d s o f t - ware ?rogram t o accomplish its c o n t r o l a c t i o n .

Some method of debugging, r e f i n i n g p.r ,eces- s a r y , and f i n a l i z i n g t h i s c o n t r o l soA ..,'are is required. I n a d d i t i o n , i t is valuable t o e v a l u a t e t h i s software i n real time s i n c e a closed loop c o n t r o l is s e n s i t i v e t o timing considerations, e s p e c i a l l y when t h e c o n t r o l is designed t o e x e r c i s e v a r i o u s p r i o r i t y l e v e l s of p r o t e c t i o n and r e g u l a t i o n .

Real-time hybrid computer simulations of t h e engine process a s a n approach t o control- l e r design and evaluation has been used i n two s u c c e s s f u l programs (9 and 10). Figure 6 shows how such simulation c a p a b i l i t i e s are u t i l i z e d . The hybrid computing system models a l l a s p e c t s of t h e propulsion process includ- ing t h e a c t u a t o r s and :ensors.

Figure 7 is a p a r t i a l d e s c r i p t i o n of how t h e component c h a r a c t e r i s t i c s (compressor maps, burner e f f i c i e n c y , e t c . ) are connected t o generate a t r a n s i e n t engine simulation r e p r e s e n t a t i v e of engine performance charac- t e r i s t i c s . When a real-time simulation has been designed, its q u a l i t y is v e r i f i e d by com- paring engine t r a n s i e n t performance informa- t i a n obtaineu from t h e 3imulatton t o a c t u a l dat, from t h e cngine. Figures 8(a) and (b) are samples of such a v e r i f i c a t i o n f o r an F-100 a i r c r a f t -urbofan engine simulation.

Both curves sho!T c l o s e correenondence between simulated and a :\.ual engine response t L an instantaneous I ement of t h e power l e v e r frm.

i d l e t o i n t e r m r . i l a t e power.

T h e control ..i:orichms a r e programmed i n fixed point assembly language ( t c accomplish t h e computing speed required) qn a s e p a r a t e d i g i t a l c o n t r o l computer represientative of t h e c a p a b i l i t i e s t h a t would eventrrally e x i s t i n t h e a c t u a l computer hardware c o n t r o l l i n g t h e real engine. After s u i t a b l e r e i b e m e n t of t h e c o n t r o l algorithms and their scftware imple- mentation, t h e c o n t r o l conip--ter c.gn be switched over t o o p e r a t e t h e a c t u a l engine hardware.

A development technique u t i l i z i n g r e a l time simulations can save much time and avoid t h e risks involved with developing c o n t r o l l a w on t h e a c t u a l engine hardware. Engine running can be m i n i m i z e d O r a t l e a s t d e f e r r e d Miller Evans and u n t i l a h i g h degree of confidence i n t h e con- t r o l l e r ' s ;]hi1 i t y t o guarantec s a f e t r a n s i e n t and s t e a d y - s t a t r engine operation has been 1 Q achieved.

The a p p l i c a t i o n of e l e c t r o n i c c o n t r o l s to advanced a i r c r a f t engines is very close to r e a l i t y . A u t m t i v e t u r b i n e s w i l l have t h e same d i f f i r u l t c o n t r o l t a s k , and a d i g i t a l e l e c t r o n i c c o n t r o l can help s a t i s f y t h e stringent c o n t r o l requirements. The use, therefore, of simulations and e s p e c i a l l y real- t h e simulations to develop t h e d i g i t a l c q u - ter c o n t r o l algorithms would prove q u i t e b e - f i c i a l . When improvesents i n t h e engine components are required. or t h e need for con- t r o l refinements arises, a s i m u l a t i o n can e f - f e c t i v e l y and e f l i c l e n t l y be used to deter- mine t h e e f f e c t s i n engine performance which can be expected. S e n s i t i v i t y s t u d i e s of corn- ponent production t o l e r . n c e s can also b e ac- complished through parametric s t u d i e s v i t h these silllations.

CollPBEsSOR AND TURBINES - Advances in t h e

technology of turboaachinery have been di- rected toward developing a b e t t e r understand- i n g of t h e fundamental f l a v processes thatoc- cur i n coeeptessors and t u r b i n e s from which ipr proved design methods and performance predic- t i o n techniques have been developed. The scope of t h e s e a c t i v i t i e s has included both the areas o f b a s i c experimental and a n a l y t i c a l research, as w e l l as f u l l scale r i g and engine tests of corepressors and t u r b i n e s f o r various aerospace and. more r e c e n t l y , automvtive ap- p l i c a t i o n s Those a s p e c t s of t h e work which are f e l t to be most a p p l i c a b l e to t h e auto- motive gas t u r b i n e engine are described 1 7 t h e following s e c t i o n s .

P e r f o G n c e P r e d i c t i o n Programs - The

a b i l i t y to p r e d i c t a n a l y t i c a l l y t h e design point performance f o r a c e n t r i f u g a l compres- sor, r a d i a l turbine, o r axial flow t u r b i n e has been w e l l e s t a b l i s h e d . The methods and com- p u t e r programs developed a t LeRC and through LeRC sponsorship are described i n (11 t o 16).

With a p p r o p r i a t e design input InformatLon such as r o t a t i v e speed, flow, i n l e t s t a t e condi- t i o n s , and power or p r e s s u r e r a t i o , t h e pro- grams can b e used t o c a l c u l a t e a n d l o r optimize t h e design point e f f i c i e n c y , t h e number of s t a g e s , t h e Ciaensions of t h e flow p a t h an- nulus, and t h e v e l o c i t y diagrams. For a x i a l flow t u r b i n e s , t h e program of (11) is simpli- f i e d t o c a l c u l a t e t h e flow c o n d i t i o n s a t t h e Evans and Miller mean blade height only. Also, for any given turbine, a l l s t a g e s have t h e same shape of diagram, with t h e shape depending on t h e s t a g e work f a c t o r and degree of r e a c t i o n s e l e c t e d .

The d e s i g a p o i n t program of (12) calcu- lates t h e radial v a r i a t i o n s of flow angle and v e l o c i t y fnnn blade hub t o t i p . Both f r e e vortex and nonfree v o r t e x d e s i g n s can be gen- erated or d y e e d . Loss c o e f f i c i e n t s are either input to t h e programs or c a l c u l a t e d in- t e r n a l l y to t h e program by t h e Pethod de- s c r i b e d i n (13).

SiPrilar methods and programs f o r calcu- l a t i n g t h e design p o i n t performance f o r cen- t r i f u g a l compressors and r a d i a l inflow t u r - bines have also been developed, and are described i n (14 t o 16).

The need f o r c a l c u l a t i n g t h e off-design performance c h a r a c t ? r i s t i c s of t h e s e campo- n e n t s over t h e i r p o t e n t i a l o p e r a t i n g range vas noted previously i n t h e s e c t i o n titled System Performance Analysis Techniques. S u c c e s s f u l a e t h o d s have been developed and coded f o r axial and radial turbines. and they are de- scribed i n (17 and 181, r e s p e c t i v e l y . The r a d i a l t u r b i n e program is based on a meanline a n a l y s i s of the flow and is f o r a single-stage turbine. The axial t u r b i n e program 2s a p p l i - c a b l e t o t u r b i n e s having up t o e i g h t s t a g e s .

The program can be run as a meanline a n a l y s i s , or can allow f o r r a d i a l v a r i a t i o n s i n l o s s and flow conditions. Tvo loss o p t i o n s are pro- vided: a k i n e t i c energy c o e f f i c i e n t - i n l e t re- covery c o e f f i c i e n t method, and a t o t a l pres- s u r e loss c o e f f i c i e n t method. The a n a l y s i s is a p p l i c a b l e from z e r o t o design speed, and t h e work done may wary up t o t h e maximum as limited by discharge annulus area choking.

The value o f an off-design performance program depends upon its a b i l i t y t o a c c u r a t e l y p r e d i c t performance over a wide range of con- d i t i o n s . An experimentally determined per- formance map, as reported i n (19) f o r a s i n g l e - s t a g e a x i a l flaw t u r b i n e , is shown i n Fig. 9.

Data were obtained over a range of speed from 40 t o 100 percent of design, and f o r p r e s s u r e r a t i o s of 1.4 to 2 . 0 .

The program of (12) (with c o e f f i c i e n s e l e c t e d t o match t h e design-point perionnance) was used t o p r e d i c t t h e t u r b i n e work and flow fcr t h e same range of conditions. Over t h e e n t i r e map. t h e pre- d i c t e d performance was w i t h i n 1 percent of t h e experimentally obtained values. Thus, t h e v a l i d i t y of t h e program was demonstrated. It Evans and Killer should be noted, however, t h a t such good agree- ment may not be experienced f o r a l l t u r b i n e s .

The methods being developed t o p r e d i c t 1 2 t h e off-design performance of c e n t r i f u g a l compressors, such as t h e method described i n (201, are still evolving. The e f f i c i e n c y v a r i a t i o n s p r e d i c t e d by t h e method of (20) have not proven to b e realistic as y e t . and experimentally-obtained values must be used.

Blade Design Prograem - I n o r d e r t o mln-

Imlee blade losses, i t is necessary to calcu- late a c c u r a t e l y t h e flow c o n d i t i o n s for a given geometry so that t h e flow d i s t r i b u t i o n throughout t h e b l a d e passage can be c o n t r o l l e d to avoid c o n d i t i o u s associated w i t h high losses. A typical f l a u passage f o r an axial and radial f l a w rotor is shown i n Figs. 10 and 11, r e s p e c t i v e l y . Gradients in v e l o c i t y occur across t h e passage from blade-to-blade and from hub-to-tip, Fig. 10. or hub-to- shroud, Fig. 11, as t h e r e s u l t of g r a d i e n t s i n static p r e s s u r e necessary t o t u r n t h e flow o r s a t i s f y r a d i a l e q u i l i b r i u m requirements. I n o r d e r to calculate t h e f l o w f i e l d i n these passages, quasi-three-dimensional a n a l y s i s techniques using meridional plane s o l u t i o n s together w i t h blade-to-blade plane s o l u t i o n s have been developed. "bo b a s i c computation methods are employed: a s t r e a w f u n c t i o n method covering t h e e n t i r e blade passage, and a velocity-gradient method covering t h e guided or covered p o r t i o n of t h e channel passage.

The programs are described i n ( 2 1 and 22).

r e s p e c t i v e l y . An example of t h e measured vs.

predicted v e l o c i t y d i s t r i b u t i o n using t h e f i r s t method is s h a m i n Fig. 12(b) f o r t h e stator b l a d e p r o f i l e shown i n Fig. 12(a). The p r e d i c t i o n s a g r e e q u i t e w e l l over most of t h e blade s u r f a c e .

I n another program, described i n (231, t h e f l o x i n j u s t t h e meridional ( r a d i a l - a x i a l ) plane of a n a x i a l , mixed, or r a d i a l flow ma- chine can he determined a n a l y t i c a l l y . The program has been extended, (24) t o allow solu- t i o n s t o be made i n annular passages without blades. T h i s program as w e l l as t h e progralss of (11 t o 13, 17, and 19 t o 23) were used to design t h e turbomachinery and ducting s e c t i o n s upstream and downstream of t h e power t u r b i n e of t h e Upgraded automotive gas t u r b i n e engine described i n (25) and shown previously i n Fig. 4.

Small-Scale E f f e c t s - Uppermost i n im- Evans and Miller

portance i n t h e a p p l i c a t i o n of turbomachinery t o t h e a u t o m t i v e g a s t u r b i n e engine I s t h e p o t e n t i a l e f f e c t of small s c a l e on perfor- mance. I n numerous programs conducted a t LeRC, t h e e f f e c t s of Reynolds number and s i z e on performance have becn i n v e s t i g a t e d f o r both curial and r a d i a l flow compressors and t u r - bines. Some of t h e more e i g n i f i c a n t r e s u l t s are summarized i n t h e following paragraphs.

I n (26) an e x t e n s i v e study w a s made t o c o r r e l a t e the e f f e c t s of t u r b i n e s i z e and Reynolds number to t u r b i n e l o s s e s . The per- formance of a t o t a l f 19 single s t a g e a x i a l flaw t u r b i n e s , ranging in s i z e from 10.2 t o 35.6 ~ I B ( 4 t o 14 in.) In diameter, and from

104 t o 2x106 in Reynolds number w e r e included

i n t h e study. The r e s u l t s i n d i c a t e d that t h e r e w a s an effect on performance due t o Reynolds number below 2X105. T h i s is shown on Fig. 1 3 by t h e negative s l o p e c h a r a c t e r i s - tics of t h e curves f o r t u r b i n e s 3, 5, and 6.

Above t h i s Reynolds number, however, t h e ef- f e c t w a s n e g l i g i b l e . The results a l s o showed a d i f f e r e n c e i n t h e l o s s parameter between s e v e r a l of t h e t u r b i n e s at t h e same Reynolds number, and t h i s w a s a t t r i b u t e d t o geometric f a c t o r s . Of t h e v a r i o u s f a c t o r s considered, s t a t o r t h r o a t area appeared t o c o r r e l a t e t h e l o s s parameter most c l o s e l y , and t h i s corre- l a t i o n is shown i n Fig. 14.

S i m i l a r programs have a l s o been conducted a t LeRC on r a d i a l inflow t u r b i n e s (27 and 281, where t h e e f f e c t s of Reynolds number, diam- eter, shroud clearance. and r o t o r configura- t i o n on performance were i n v e s t i g a t e d . The e f f i c i e n c i e s were a f f e c t e d by Reynolds number v a r i a t i o n s over t h e range of v a l u e s l n v e s t i - gated, and t h e s e results a r e shown i n Fig. 15.

However, t h e r e w a s no s i g n i f i c a n t d i f f e r e n c e i n performance due t o d i f f e r e n c e s i n r o t o r diameter between t h e t h r e e diameters i n v e s t i - gated of 8.89, 11.68, and 15.29 cm (3.50, 4 . 5 9 , and 6.02 i n . ) . The performance w a s a l s o g e n e r a l l y i n s e n s i t i v e t o shroud c l e a r - ances and blade loading over t h e r a d i a l por- t i o n of t h e roto:, but was found t o be sensi- t i v e t o t h e geometry of t h e r o t o r exducer and e x i t d i f f u s e r duct.

The e f f e c t s of v a r i a t i o n s i n Reynolds number on t h e performance of a small c e n t r i f - ugal and a x i a l flow compressor has a l s o been i n v e s t i g a t e d and compared (29 and 30). Both compressors were designed f o r t h e same a p p l i - c a t i o n , and a r e shown i n Figs. 16 and 1 7 . Evans and Miller The c e n t r i f u g a l compressor had a t i p diaml-ter of 15.2 c m ( 6 i n . ) . The a x i a l flow compres- s o r had s i x s t a g e s and a t i p diameter of 1 4 9.4 c m ( 3 . 7 i n . ) . A comparison of l o s s e s w i t h percent change i n i n l e t p r e s s u r e (shown as per- cent of design Reynolds number i n Fig. 18 f o r t h e two compressors), showed t h e c e n t r i f u g a l compressor t o b e less s e n s i t i v e t o reductions i n Reynolds number than t h e axial flow c o r pressor. I f similar t r e n d s i n performance oc- cur with down-sizing, as might be required f o r an automotive a p p l i c a t i o n . t h e r e s u l t s suggest an axial flow compressor would s u f f e r a l a r g e r reduction i n e f f i c i e n c y than a c e n t r i f u g a l compressor.

Variable Geometry - The e f f e c t of varying

t u r b i n e s t a t o r o r compressor d i f f u s e r s e t t i n g angles to extend t h e off-design o p e r a t i n g range and e f f i c i e n c y of t u r b i n e s and compres- sors has been explorec i n s e v e r a l programs.

These i n v e s t i g a t i o n s have been d i r e c t e d toward such aerospace a p p l i c a t i o n s as space power generation and jet engines f o r supersonic f l i g h t , but have a l s o included t h e e f f e c t s of v a r i a b l e geometry on automotive g a s t u r b i n e compressor and t u r b i n e performance.

The results of a recent i n v e s t i g a t i o n of t h e v a r i a b l e geometry powe; t u r b i n e f o r t h e sixth-generation Chrysler Baseline engine, (31), are c u r r e n t l y being evaluated. Refer- ence (32) d e s c r i b z s t h e measured performance, v e l o c i t y diagrams, and e x i t d i f f u s e r charac- t e r i s t i c s of t h e t u r b i n e o p e r a t i n g a t its design s t a t o r chord s e t t i n g angle of 3S0 from t. .lbential. I n i t i a l r e s u l t s d e s c r i b i n g t h e performance a t off-design s t a t o r s e t t i . l g angles are shown i n Fig. 19. The sharp varia- t i o n i n s t a g e e f f i c i e n c y w i t h s t a t o r s e t t i n g angle is evidept. The e f f e c t of s t a t o r end- w a l l clearance, necessary t o allow f o r actua- t i o n of t h e s t a t o r s , is a l s o noted on t h e f i g u r e .

T1.e flow mechanisms which cause Chis v a r i a t i o n i n e f f i c i e n c y have been t h e s u b j e c t of rrumerous i n v e s t i g a t i o n s . The results a r e 3ummarized i n (33) f o r a t u r b i n e having t h e same design p r e s s u r e r a t i o , and s t a t o r and r o t o r r e a c t i o n c h a r a c t e r i s t i c s a s t h e Chrysler Baseline power t u r b i n e . A cross-section of t h e s t a g e i s shown i n Fig. 20. From t h e ex- p e r i m e n t a l r e s u l t s , a l o s s breakdown a n a l y s i s was made which showed t h e r e l a t i v e e f f e c t s of varying t h e s t a t o r angle on s t a t o r , r o t o r , and incldence l o s s e s . The r e s u i t s a r e shown Evans and M i l l e r i n Fig. 2 1 f o r t h e t h r e e s t a t o r s e t t i n g angles 8 h m in Fig. 20 (design, 30% open, and 30% closed from design) which corresponds t o an angle v a r i a t i o n from design of approximately

- 7 O and +a0. The e f f e c t of varying t h e t u r -

b i n e p r e s s u r e rat?;., on e f f i c i e n c y and losses is a l s o shown on t h e f i g u r e . From t h e results, it is evident t h a t considerable changes occur i n t h e t u r b i n e when v a r i a b l e geometry is im- posed. The results emphasize t h e need t o de- termine thesc changes r e l a t i v e t o t h e intended a p p l i c a t i o n such that t h e b e s t design can ul- timately be made.

COMBUSTORS - Ongoing programs i n combus-

t i o n research include t h e f i v e basic areas of combustor aerodynamics, l i n e r cooling, f u e l preparation, a l t e r n a t e f u e l s , and exhaust emissions. Much of t h e e f f o r t has, of course, been d i r e c t e d toward t h e advancement of air- c r a f t gas t u r b i n e combustion systems, but be- cause of lts fmdsmental nature, it is equally a p p l i c a b l e t o t h e automotive gas t u r b i n e com- bustor. I n a d d i t i o n t o t h i s b a s i c research, s e v e r a l exploratory i n v e s t i g a t i o n s of various advanced combustor concepts have demonstrated performance c h a r a c t e r i s t i c s which are f e l t t o b e p a r t i c u l a r l y a t t r a c t i v e t o t h e automotive a p p l i c a t ion.

Basic Research - Combustor aerodynamics

are t h e aerodynamics a s s o c i a t e d w i t h t h e de- s i g n and a n a l y s i s of t h e combustor geometry and its e f f e c t on air f l o w d i s t r i b u t i o n , pres- s u r e l o s s e s , and e x i t temperature p r o f i l e s .

Through t h e use of a c o w u t e r program devel- oped under LeRC sponsor:ihip ( 34). extensive analyses of t h i s compler problem have been made. An example of t h e use of t h e program as a design o r a n a l y s i s t o c l is described i n (35) The e f f o r t s i n l i n e r cooling have been d i r e c t e d towards t h e development o f a n a l y t i c a l

temperature prediction programs ( 3 5 ) , and im-

proved f i l m cooling techniques. A perforated- s h e e t l i n e r has been developed which is one of t h e simplest and lowest c o s t f i l m cooled li- n e r s t o f a b r i c a t e . I t has been t e s t e d i n a t u r b o j e t combustor r i g and compared t o a con- ventional stepped-slot film cooled l i n e r ( 3 7 ) .

No adverse e f f e c t s were found using t h e perforated-sheet l i n e r r e l a t i v e t o its f i l m cooling e f f e c t i v e n e s s , combustion e f f i c i e n c y , pressure drop, o r e x i t temperature p r o f i l e .

I n v e s t i g a t i o n s i n t o t h e design and per- formance of various advanced f u e l i n j e c t i o n Evans and Miller and f u e l preparation techniques have been con- ducted and extensively documented. Two injec- t i o n techniques, t h e a i r - b l a s t and a i r - a s s i s t nozzles, (38 and 3 9 ) have been found t o have s u p e r i o r emission c h a r a c t e r i s t i c s compared t o more conventional f u e l i n j e c t i o n techniques.

Fuel premixing and prevaporizing techniques have been i n v e s t i g a t e d f o r various advanced combustor concepts such as the c a t a l y t i c com- bustor described i n t h e next s e c t i o n . Funda- mental i n v e s t i g a t i o n s of premixivg- prevaporizing combustors (40 t o 421, have es- t a b l i s h e d that t h i s system can achieve very l o w NO, levels, Fig. 22. Also, t h e use of f u e l s t a g i n g between two or more i n j e c t o r s has been e f f e c t i v e f o r c o n t r o l l i n g t h e perfor- mance and emissions of a combustor over t h e t h r o t t l a b l e range of an engine.

E f f o r t s a t LeRC o r under LeRC sponsorship on a l t e r n a t e f u e l s range frcim t h e study of t h e t e c h n i c a l problems of converting s h a l e o i l o r coal-derived syncrudes t o t u r b i n e f u e l , t o e v a l u a t i o n s of t h e thermal s t a b i l i t y and test- i n g of t h e s e as w e l l as o t h e r a l t e r n a t e f u e l s i n combustor r i g s and engines. Comprehensiire s t u d i e s of t h e r e f i n i n g requirements and ther- m a l s t a b i l i t y c h a r a c t e r i s t i c s t h a t have been made t o d a t e are given i n (43 and 44). The comparative r e s u l t s of running a combustor on Jet A and Diesel number 2 f u e l s ( 4 5 ) showed comparable l e v e l s of N O , and CO emissions be- tween t h e two f u e l s , but approximately twice the unburned hydrocarbons and smoke number with Diesel number 2.

The i n i t i a l emphasis i n t h e emission in- v e s t i g a t i o n s was t o develop and demonstrate the technology required t o reduce t h e HC, CO, and N O , p o l l u t a n t s i n c u r r e n t and f u t u r e air- c r a f t engines. The r e s u l t s of a p o r t i o n of t h i s broad based program a r e summarized i n ( 4 6 ) . Other a r e a s of emissions i n v e s t i g a t i o n s have includcd t h e e f f e c t s of water i n j e c t i o n (471, exhaust gas r e c i r c a l a t i o n (48) heat pipe regenerators (49), t h e e f f e c t of f u e l temper- a t u r e on NOx formation (50) and exhaust odors, (51). Referring t o t h i s latter reference, t h e recent r e s u l t s from a series of combustor r i g tests run with f i v e d i f f e r e n t f u e l s showed t h a t odor i n t e n s i t i e s were relaLed t o t h e con- certtration of oxygenates i n the exhaust, Fig. 23, wtich were i n t u r n proportional t o ‘.he e f f i c i e n c y of t h e combustion process, re- g a r d l e s s of t h e f u e l used.

Advanced Concepts - The high I n l e t and Evaris and Miller

o u t l e t combustor temperatures 1 squired f o r advanced a i r c r a f t and regenerative automotive gas t u r b i n e engines have resulced in both fa- vorable and unfavorable e f f e c t s on t h e com- b u s t i o n process. The h i g h temperatures are conducive t o high combustion e f f i c i e n c i e s and hence l o w hydrocarbons and carbon monoxide emissions. However, t h e high temperatures a l s o i n c r e a s e t h e formation of NO,, and sev- eral unique and advanced combustor concepts have been under i n v e s t i g a t i o n t o minimize t h e s e N O , formations.

The p r i n c i p a l of t h e LeRC developed swirl- can combustor f o r a i r c r a f t a p p l i c a t i o n , shown i n Fig. 24(a), is t o d i v i d e t h e flame zone i n t o many small zones through t h e use of many small combustor modules, c a l l e d swirl-cans.

T h i s approach provides l e a n e r combustion through premixing, and reduced hot g a s resi- dence times. Typical r e s u l t s have shown a 30 t o 35 percent reduction i n :he NOx emis- s i o n s , Fig. 25, compared t o conventional com- b u s t o r s o p e r a t i n g at c o n d i t i o n s representa- t i v e of commercial jet a i r c r a f t engines.

A second concept, premixed-prevaporized combustion which w a s noted i n t h e previous s e c t i o n , atomizes, vaporizes, and mixes t h e f u e l and a i r t o allow combustion t o t a k e place at l e a n e r f u e l - a i r r a t i o s . The p r e d i c t e d N O , emission levels f o r t h i s type of combustor f o r a t y p i c a l current and p r o j e c t e d f u t u r e auto- motive gas t u r b i n e engine are shown i n Figs.

26(a) and (b). One a p p l i c a t i o n of t h i s con- cept is c u r r e n t l y under development f o r t h e Chrysler Upgraded Gas Turbine engine, Fig. 24(b).

Two a d d i t i o n a l concepts under investiga- t i o n are t h e c a t a l y t i c combustor and t h e multi-element combustor, shown schematically i n Figs. 24(c) ar.d ( d ) . These two combustor concepts a r e described i n more d e t a i l i n t h e following s e c t i o n s .

C a t a l y t i c Combustor - The combustion

process i n t h i s concept is heterogeneous and r e q u i r e s a prevaporized, premixed f u e l - a i r preparation system as shown enlarged i n Fig. 27. The mixture is then reacted i n a c a t a l y s t bed where t h e catalyzed s u r f a c e of t h e bed allows conbustion t o take place a t equivalence r a t i o s w e l l below t h e l e a n flam- mability l i m i t s ( 5 2 ) . The maximum combustion temperature reached during t h e r e a c t i o n pro- c e s s is only s l i g h t l y above the t u r b i n e i n l e t temperature. As a result, low NO, is p o s s i b l e Evans and Miller with high combustion e f f i c i e n c y . Preheating of t h e bed during cold s t a r t i n g is requ:red t o i n i t i a t e the process f o r t h e c a t a l y s t s inves- t i g a t e d t h u s f a r .

The concept has been under experimental investigation in combustion rigs for several years. Most of the catalyst test elements consisted of a cylindrically-shaped block of ceramic honeycomb material coated with a noble metal catalyst. The results obtained thus far are presented in (52 and 53). Based on these results, the predicted NOx emission levels, as shown in Fig. 26 for the autono- tive application, are considerably below either the goal levels or the premixed- prevaporized combustor levels.

Multi-Element Combustor - The principal

of this concept is to reduce the formation of N O , by removing sufficient temperature from the flame to prevent the adiabatic flame tem- perature from being reached, and by increasing the hot gas velocities through the combustor by the use of turbulent flame holding to re- duce the flame residence time. The concept is shown schematically in h ~ g . 28(a) for a single combustor element. Unmixed fuel anu air enters a round stepped combustor passage located adjacent to a dilution air passage intrigally cast in a ceramic block. Premix- ing occurs at the inlet of the combustor pas- sage, and homogeneous combustion occurs in the downstream portion of the passage. The steps provide high turbulence and good flame holding characteristics. Heat is transferred from the flame to the surface walls of the combustor passage, and hence to the dilution air through the ceramic interface between the two passages. A photo of the flame in an ex- perimental single-element combustor is shown in Fig. 28(b). A complete combustor would consist of multiple elements of combustor passages interspersed between dilution air passages as shown in Fig. 24(d). As engine power or speed is varied, fuel to the individ- ual combustor passages is staged or regulated to vary the overall fuel-air ratio of the con- bustor. Emission measurements from the single element tests, extrapolated to a multi-element comhustor sized for the automotive application is shown in Fig. 26. Compared to the premixed-prevaporized and catalytic combustors, the results are very encouraging thus far.

COATINGS AND MATERIALS - Cersmic Thermal

Barrier Zoatings - A ceramic coating system Evans and Miller

has been evolved for cooled metal engine com- ponents such as rocket nozzles, combustor liners, turbine blades and vanes. The coat- ing serves to insulate the metal parts from t h e hot combustion gasses through its low thermal conductivity (approximately 1.5 W / m K) and high r e f l e c t i v i t y (from two t o four times t h a t of a t y p i c a l high temperature metal a l l o y ) . A s a r e s u l t , the metal temperature of t h e cooled coated p a r t s , o r t h e amount of cooling a i r required t o cool the parts is re- duced.

The coating concept i s i l l u s t r a t e d i n Fig. 29 f o r a t y p i c a l cooled metal surface.

Both a metallic bond coat and a ceramic in- s u l a t i n g l a y e r are applied t o t h e s u r f a c e by plasma spraying. The most recent and success- f u l coating system c o n s i s t s of an i n s u l a t i n g l a y e r of y t t r i a s t a b i l i z e d z i r c o n i a ceramic (ZrO2-12Y2o3) 0.025 t o 0.076 c m (0.01 t o 6.03 in.) t h i c k applied over a nickel-chromium- aluminum yttrium a l l o y ( N i C r A l Y ) base coat 0.013 c m (0.005 in.) t h i c k .

A paper given a t t h i s :onference last year ( 5 5 1 , described t h e sj'stem i n d e t a i l in- cluding its performance and d u r a b i l i t y i n tests on t u r b i n e blades, vanes, and combustor l i n e r s . Because of considerations of small s i z e and low c o s t f o r an automotive gas t u r - b i n e a p p l i c a t i o n , t h e use of the concept may be l i m i t e d t o such components a s t h e combustor l i n e r , its interconnecting duct t o t h e tur- bine, and vane shrouds. The coating is cur- r e n t l y being considered f o r the vane shrouds and p o r t i o n s of t h e combustor f o r an army tank engine.

In r i g t e s t s of a combustor from a com- mercial j e t engine, Fig. 30 and (56), a coated l i n e r operated a t metal temperatures 130 t o 210 K (230" t o 380" F) cooler than an uncoated l i n e r over t h e operating range of t h e combus- t o r . Smoke concentrations and soot were a l s o reduced due t o t h e coating r e f l e c t i n g a l a r g e r amount of t h e i n c i d e n t r a d i a t i o n back t o t h e flame. The r e f l e c t e d energy reduced t h e amount of unburned hydrocarbons which could a l s o be b e n e f i c i a l t o i n c r e a s i n g combustion e f f i c i e n c y and reducing emissions. This par- t i c u l a r combustor, however, already had a high e f f i c i e n c y of 0.999, and no s i g n i f i c a n t change i n e i t h e r t h e e f f i c i e n c y or emissions was ob- served.

Ceramic Materials Development - 21. i t i c a l

to meeting t h e needs f o r f u e l e f f i c i e n t high Evans and Miller performance gas t u r h i n e engines of the f u t u r e is t h e a v a i l a b i l i t y of high temperature low cost m a t e r i a l s f o r t h e burner and t u r b i n e components. I n t h e case of t h e automotive gas t u r b i n e engine, t h i s a l s o includes t h e regen- e r a t o r component. Ceramics o f f e r t h e b e s t po- t e n t i a l f o r meeting these requircments because of t h e i r low c o s t and low d e n s i t y ( p r o j e c t e d to be about 1/10 t h e c o s t of superalloys and 1/3 the d e n s i t y ) , as w e l l a s t h e i r high s t r e n g t h at o p e r a t i n g temperatures up t o 1670K (2600' F) and higher. For t u r b i n e b l a d e s where t h e primary stress r e s u l t s from c e n t r i f - ugal f o r c e , t h e i r low density and high strength-to-density r a t i o makes them particu- l a r l y a t t r a c t i v e . However, t h e l a c k of duc- t i l i t y and very low impact r e s i s t a n c e of ceramics probably won't permit t h e i r use in hot s e c t i o n components of a i r c r a f t engines u n t i l t h e problem of impact f a i l u r e s of t h e ceramic p a r t s and t h e e f f e c t on f l i g h t s a f e t y can be solved. There is, however, a much g r e a t e r likelihood t h a t they w i l l see service i n automotive t u r b i n e s i n t h e foreseeable fu- t u r e . A s noted previously, ceramics are al- ready i n use i n automotive regenerators, and a l s o in experimental t u r b i n e s where an engine demonstration of an all-ceramic t u r b i n e has r e c e n t l y been completed.

The key elements i n t h e study and devel- opment of ceramic m a t e r i a l s technology as i t a p p l i e s t o t h e p a r t i c u l a r components of t h e engine are summarized i n t h e following para- graphs. Currently the most promising s t r u c - t u r a l ceramics appear t o b e Si3N4 and SIC.

Extensive screening s t u d i e s of 35 d i f f e r e n t ceramics i n t h e LeRC Mach 1 burner r i g , Fig. 31 and ( 5 7 1 , have shown t h a t Si3N4 and SIC based ceramics have t h e most favorable re- s i s t a n c e t o thermal shock. T h e s e ceramics a r e a l s o much more oxidation r e s i s t a n t than super- a l l o y s . Figure 32 shows a ceramic and a cooled superalloy blade which had been sub- j e c t e d t o i d e n t i c a l exposures i n t h e Mach 1 burner r i g a t 1470 K (2190" F). T h e ceramic blade shows l i t t l e e f f e c t of t h e exposure, but t h e a l l o y blade is badly cracked and eroded.

The Si3N4 and SIC ceramics a l s o have ex- c e l l e n t high temperature creep rupture prop- erties, a s may be seen i n Fig. 33. The f i g u r e compares commercially a v a i l a b l e hot pressed Si3N4 and experimental Si3N4 and SIC t o the s t r o n g e s t known c o n v e n t i m a l l y c a s t vane a l - loy, WAZ-16, and t h e oxide dispersion strength- Evans and Miller ened (ODS) a l l o y s .

Combustor l i n e r s , t u r b i n e vanes, and in- terconnecting ducts which a r e r e l a t i v e l y low stressed components w i l l have t o withstand peak c y c l e o p e r a t i n g temperatures up t o 1640 K

(2500" F) , as w e l l as rapid changes i n gas

temperatures due t o er-gjne l i g h t - o f f and shut- down t r a n s i e n t s , and t h r o t t l e excursions, To meet these requirements, p a r t s such as s t a t o r vanes may t y p i c a l l y have t o o p e r a t e at stresses up t o 50 MN/m2 (7000 p s i ) , which is c l o s e t o t h e c u r r e n t c a p a b i l i t y of t h e commercially a v a i l a b l e Si3N4 shown on Fig. 33. Rotor blades blades, which w i l l o p e r a t e a t somewhat lower temperatures, may t y p i c a l l y have t o oFerate a t stresses up t o 207 MN/m2 (30,000 p s i ) which is c l o s e t o t h e c u r r e n t c a p a b i l i t y of t h e experi- mental S i c shown on t h e f i g u r e .

Work under LeRC sponsorship is c u r r e n t l y d i r e c t e d towards improving t h e creep-rupture s t r e n g t h of Si3N4 by reducing t h e a l k a l i metal, F u r t h e r 02, and densifying a d d i t i v e s (58).

improvemect3 i n s t r e n g t h and impact r e s i s t a n c e , as w i l l b e required f o r t u r b i n e blade applica- t i o n s , are a l s o being i n v e s t i g a t e d through im- proved ceramic processing procedures, and t h e development of energy absorbing crushable sur- f a c e l a y e r s (59). Large improvements i n t h e creep r u p t u r e and l i f e p r o p e r t i e s of SI3N4 have been achieved through t h e use of higher p u r i t y s i l i c o n powders, reduced Ca c o n t e n t , and t h e s u b s t i t u t i o n of Z r 0 2 f o r NgO as a densifying a i d , Fig. 34 and ( 5 9 ) . Current work is a l s o showing (60) t h a t Si3Nq can b e d e n s i f i e d t o 95% of i t s t h e o r e t i c a l value without t h e a i d of a d d i t i v e s through t h e use of hot i s o s t a t i c pressing t o p r e s s u r e s up t o 276 MN/m2 (40,000 p s i ) . This pressure is 10 times t h e p r e s s u r e necessary t o achieve t h i s l e v e l of d e n s i t y when a d d i t i v e s a r e used.

This should permit a reduction o r e l i m i n a t i o n of the a d d i t i v e s which reduce high temperature s t r e n g t h , and should lead t o f u r t h e r improve- ments i n the high temperature s t r e n g t h prop- e r t i e s of t h e m a t e r i a l .

a-phase Si3N4 I n c r e a s i n g t h e p u r i t v of powders (61) combined with t h e a p p l i c a t i o n o f t h e crushable energy absorbing l a y e r s noted i n (59). have provided an i n c r e a s e i n impact s t r e n g t h . The use of a porous layer of reac- t i o n s i n t e r e d Si3N4, t h e t p s t all-around ap- proach t o d a t e , has r e s u l t e d i n a b a l l i s t i c impact r e s i s t a n c e of 1 1 . 4 j o u l e s ( 8 . 4 ft l b ) , which i s a n i n c r e a s e o f 6 over t h e unprotected Evans and Miller Si3N4.

I n v e s t i g a t i o n s have been conducted i n t o t h e problem of i n t e r f a c i n g ceramics w i t h metals (e.p;. . ceramic b l a d r s t o metal disks).

A r l y work, (62 and 63) recognized t h e need t o .ccinnmodate t h e l a c k of d u c t i l i t y of ceramics by employing generous r a d i i and cushioning in- t e r f a c e s between the b l a d e s and d i s k t o pre- vent stress concentrations, as w e l l as t o pre- vent chemical r e a c t i o n s between t h e two mate- rials. Successful tests were conducted i n a i r c r a f t engines up t o f u l l power without b l a d e r o o t f a i l u r e s . I n t e r f a c e s made of a porous o r s c r e e n material were p a r t i c u l a r l y b e n e f i c i a l . Ductile s h e e t metal used as a compliant l a y e r at t h e i n t e r f a c e h a s been used i n more recent i n v e s t i g a t i o n s , (64) and Fig. 35, and have been adapted f o r use i n t h e ARPAINAVSEA-AiResearch Ceramic Gas Turbine Engine Demonstration Program.

An o v e r a l l summary of t h e c u r r e n t s t a t u s of advanced high temperature t u r b i n e material, c o a t i q s , and technology r e l e v a n t t o t h e air- c r a f t gas t u r b i n e engine is given i n ( 6 5 ) .

S U M M A R Y The automotive gas t u r b i n e engine is an a t t r a c t i v e a l t e r n a t i v e powerplant because of its low emissions and a l t e r n a t e o r multi-fuel c a p a b i l i t y . With t h e performance improvements p r o j e c t e ? through t h e a p p l i c a t i o n of currently- evolving and advanced technology, i t may be- come a competitive a l t e r n a t i v e f o r t h i s appli- c a t i o n .

I n t h i s paper, t h e a u t h o r s and coniribu- t o r s present a summary of t h e c u r r e n t develop- ments and f u t u r e requirements f o r t h e automo- t i v e gas t u r b i n e zngine, as w e l l as an over- view of t h e aerospace-related r e s e a r c h and technology developments i n s e v e r a l of t h e areas whSch may have a s i g n i f i c a n t impact on meeting t h e s e f u t u r e requirements f o r t h e automotive a p p l i c a t i o n . The areas discuesed are system a n a l y s i s , comprersors and tur- bines, combustors, coatings, and naterials.

The NNEP s t e a d y - s t a t e system performance a n a l y s i s model, and a t r a n s i e n t performance simulation technique using a hybrid computer have been developed and used extensively a t LeRC as an e f f e c t i v e t o o l t o a n a l y t i c a l l y de- termine and e v a l u a t e v a r i o u s t y p e s of gas t u r - b i n e systems. The NNEP program is c u r r e n t l y Evan6 and Miller being ueed i n a screening study t o p r e d i c t t h e f u e l economy p o t e n t i a l f o r a matrix of advanced automotive gas t u r b i n e engine configurations. 23 Computer codes have also been developed and used e x t e n s i v e l y at LeRC as w e l l as by t h e ,as t u r b i n e i n d u s t r y t o p r e d i c t and optimize t h e pcrformance, aerodynamic requirements, and geometric c o n f i g u r a t i o n s f o r a x i a l and ra- d i a l flow comprec i o r s and t u r b i n e s . Good agreement between t h e a n a l y t i c a l l y p r e d i c t e d and experimentally measured performance maps and blade s u r f a c e v c l o c i t y d i s t r i b u t i o n s have been obtained i n most cases. T h e a p p l i c a t i o n of t h e v a r i o u s codes range from nonbladed un- n u l a r passages t o s i m p l i f i e d qean-nassage- height blade and s t a g e a n a l y s i s , t o quasi- three-dimensional free-stream and blade chan- n e l flow a n a l y s i s where r a d i a l v a r i a t i o n s i n v o r t i c i t y , work, and l o s s e s are accounted f o r .

Experimental e v a l u a t i o n s of t h e perfor- mance .And l o s s mechnnisms t h a t occur i n com- p r e s s o r s and t u r b i n e s due t o small s c a l e and v a r i a b l e geome’.ry e f f e c t s emphasize t h e need t o a c c u r a t e l y p r e d i c t and design f o r t h e s e e f f e c t s t o a s s u r e minimum impact on perfor- mance f o r t h e automotive aDplication.

Research i n fundamental combustor tech- nology has l e a d t o the develorment o f : a computer rode which can b e used as a design o r a n a l y t i c a l t o o i LC, d e s c r i b e t h e aerody- namic performance of combustors; a n e f f i c i e n t low-cost p e r f o r a t e d s h e e t l i n e r ; f u e l nozzle designs and f u e l preparation techniques such as premixing and prevaporizing, which have achieved very low NOx omission l e v e l s ; and a broader technology base f o r low emission com- bustiot- fundamentals, and f o r t h e r e f i n i n g , charac.-tization and emission l e v e l s of a l - t e r n a t e f u e l s . The performance c h a r a c t e r i s t i c s of Revera1 advanced combustor designs employ- i n g m u l t i p l e s w i r l cuprs, c a t a l y s t , and multi- p l e element concepts show considerable promise f o r meeting t h e performance and emission g o a l s f o r advanced automotive a p p l i c a t i o n s .

The development and u s e of a ceramic thermal b a r r i e r coating on a combustor l i n e r r e s u l t e d i n a l a r g e reduction i n l i n e r temper- a t u r e , and e x h i b i t e d t h e p o t e n t i a l t o reduce the N O , emission l e v e l s f o r some a p p l i c a t i o n s .

Research ef a r t s i n s t r u c t u r a l ceramic materials have lead t o improved processing, design techniques, a n ’ m a t e r i a l s p r o p e r t i e s .

T h e s e e f f o r t s h a v e b t r , l d i r e c t e d toward t h e Evans and Miller c r i t i c a l goal of f a b r i c a t i n g hot s e c t i o n com- ponents of ceramic m a t e r i a l s t o meet t h e c o s t and performance requirements for advanced automotive R ~ S t u r b i n e engines. Increases i n t h e high temperature stress r u p t u r e p r o p e r t i e s bve heen made through improved material pur- ities and improved densifying a i d s . or through t h e use of i s o s t a t i c p r e s s i n g p r e s s u r e s of up t o 10 tiaes t h e p r e s s u r e necessary t o achieve t b e same level of material d e n s i t y when dens- f€yfng aids are used.

The use of porous sur- f a c e l a y e r s have demmstrated a s i x f o l d in- crease i n iqmct r e s i s t a n c e over unprotected ceramic surfaces. A l s o , t h e use of compliant =tal i n t e r f a c e l a y e r s have been developed which reduce stress concentrations and chemi- c a l r e a c t i o n s between ceramic and -tal parts.

1. "Should Zr'e Have A N e w Engiae, A n Autm- b i l e Power Systems Evaluation, Volume 11." Jet Propulsion Lab report JPL-SP-43-17; SM-SP-600, August 1975, pp. 5-1 t o 5-44.

2. J . L. Klanu, and R. Tew, "Analysis of Regenerative Single-Shaft Certmic Gas Turbine Engines a . d Resulting Fuel Econorny i n a Compact Car." N S A RI X-3531, 1977.

3. S . 0. Kronograd, 'Three-Shaft Automotive Turbine Transmission of t h e K l T Type-Performance and Features." Paper 77-GT-94 presented a t ASME Gas R - b i n e and Products .Chaw, Philadelphia, March !.977.

4. S . H . Nosek, "Ceramics t o r t h e Advanced

Autamotive Gas Turbine Engine - A Look a t t h e

Single-Shaft Engine." NASA Tn X-73651, 1977.

5 . L. H. Fishbach, and H . J. Caddy, "NNEP -

The Navy N A S A Engine Program." N A S A TM X-71857, 1975.

6. J. R. Stuch, and W. H. Bruton, "Real- Time Simulation of t h e TF30-P-3 Turbofan Engine Using a a y b r i d Computer." NASA TM X-3106, 1974.

7 . J. R. Stuch, and K. Seldner, "Rea\-Time Simulation of t h e FIOO-PW-100 Turbofan Engine Using t h e Hybrid Compcter." N A S A Tn X-3261, 1975.

8 . J . R. Szuch, K. Seldner, and D . S .

Cwynar, )* Development and V e r i f i c a t i o n of a Real-Time, Hybrid Computer S i m l a t i o n of t h e FlOO-PW-lOO(3) S e r i e s I1 Engin?." N A S A TP-1034, 1977.

9 . J. R . Stuch, C. S k i r a , and J. F. soeder, "Evaluation of an FlOO Multivariable Control Using a Real-Time Enriqe Simulation." Paper Evans and Miller 77-834 presented a t 1 and S A E T h i r t e e n t h Pro- pclsion Conference, urlando, F l o r i d a , J u l y 1977.

N A S A TM X-73648. 2 5 10. L. 0 . B i l l i g , ' I n t e g r a t e d Propulsion Control System (IPCS), Volume I, Summary."

AFAPL-Iyf-76-61, Auqust 1976; AD-AO33062.

11. A. J . Glassman, "Computer Program f o r Preliminary Design Analysis of Axial-Flow Tur- bines .le NASA TN 0-6702, 1972.

12. A. F. Carter, and F. K. Lenherr, "Ana- y s i s of Geometry and Design-Point Performance of Axial-Flow Turbine Using S p e c i f i e d Meridional Velocity Gradients ." NASA CR-1956, 1969.

13. A. F. Carter, M . P l a t t , and F . K.

Lenherr, "Analysis of Geometry and Design-Point

Performance of Axial-Flow Turbines. P a r t I -

Development of t h e Analysis PIethods and t h e Loss C o e f f i c i e n t s Correlation." NASA CB-1181, 1968.

14. H . R . G s ~ v ~ s , "Analytical C o r r e l a t i o n o f C e n t r i f u g a l Compressor Design Geometry f o r n3xiaa1~1 E f f i c i e n c y w i t h S p e c i f i c Speed." NASA TN 0-6729, 1972.

15. J. E . Rohlik, "Analytical Determination of Radial Inflow Turbine Design Geometry f o r Maximum Efficiency." NASA TN 0-4384, 1968.

16. A. J . Glassman, "Coraputer Program f o r Design Analysis of Radial-Inf low Turbines .I1 NASA TN 0-8164, 1976.

17. E. E . Flagg, "Analytical Procedure and Computer Progran f o r Determining t h e Off-Design Performance of Axial Flow Turbines . I 1 N A S A Ch-710, 1967.

18. C. A. Wasserbauer, and A. J. Classman, "Fortran Program f o r P r e d i c t i n g Off-Design Per- formance of Radial-Inf lw Turbines .** NASA TN 0-8063, 1975.

19. W. J. Whitney, E. M. Szanca, B. Bider, and 0 E. Monroe, "Cold-Air I n v e s t i g a t i o n of a Turbi-le f o r High-Temperature-Engine Application.

111 - Overall Stage Performance." NASA TN

0-4389 , 1968.

N. R. Galvas, "Fqrtran Program f o r Cal- 20.

c u l a t i n g Total-Efficiency-Specific Speed (213r- a c t e r i s t i c s of C e n t r i f u g a l Compressors .I' NASA TM X-2594, 197:.

21. T. Katsanis, "Fortran Program f o r Cal- c u l a t i n g Transonic V e l o c i t i e s on a Blade-to- Blade Stream S u r f a c t of a Turbomachine." NASA TN D-3427, 1969.

22. T. K i t s a n i s , "Fortran Program f o r Quasi-Three-Dimensional C a l c u l a t i o n of Surface V e l o c i t i e s and Choking Flow f o r Turbomachine Evens end M i l l e r Blade Rows . I 1 NASA TN D-6177, 197 1.

23. T. Katsanis, and W . D. McNally, "For- t r a n Program f o r Calcul..ting V e l o c i t i e s and Streamlfnos on t h e Hub-Shroud Mid-Channel Flow 26 Surface of an Axial- o r Mixed-Flow Turbomachine, I - User's b n u a l . " NASA T N D-7343, 1973.

24. T. #atsattis, rmJ U. D. Mdsally, "Be- vised Fortran Program f o r Calculating Velocities ead S t r e m l i n e s on the Hub-Shroud Hidchannel Stream Surface of an Axial-, Radial-, o r Xixed-

Flav Turbaeachine o r Annular Duct, I - User's

tQanue1." NASA TN D - 8 4 3 0 , 1977.

25. C. A. P a l l , J. I. Gumaer, and T. H .

Sebes tyen, "The ERDA/Qlrys ler Upgraded Cas Tim- bine Engine Objectives and Design." Paper 760279 presented a t SAE Automotive Engineering Congress and Exposition, Detroit, Feb. 1976.

26. D. E. Holeski, and W. L . Stewart, 'study of XACA and NASA Single-Stage Axial Flow m r b i n e Performance a s Belated to Reynolds h m - ber and Geometry.'' Journal of Engineering f o r Power, Vol. 86, J u l y 1964, pp. 296-298.

27. H. G. Kofskey, and C. A . Wasserbauer, "Experimental Evaluation of a 3.50-Inch Radial Turbine Designed f o r a 10-Kilowatt Space Power System." NASA TN D-5550, 1969.

28. H. E. Rohlik, and H . G . Kofskey, "Re- cent Radial Turbine Research a t the NASA Lewis

Research Center ." Paper 72-GT-42 presented a t

ASME Gas Turbine and Fluids Engineering Confer- ence, San Francisco, March 1972.

29. L. J. Heidelberg, C. H. Ball, and C.

Weigel, "Effect of Reynolds Number on Overall Performance of a 6-Inch Radial Balded Centrigu- g a l Compressor." NASA TN D-5761, 1970.

30. L. J. Heidelberg, and C. L. Ball, "Effect of Reynolds Number on Overall Perform- ance of a 3.7-Inch Diameter Six-Stage Axial-Flow CaPnpressor .'' NASA TN D-6628, 1972.

31. P. R. Angell, and T. Golec, "Upgrading Automotive Gas Turbine Technology; An Experi mental Evaluation of Improvement Concept."

Paper 760280 presented a t S A E Automotive Engin- eering Congress and Exposition, Detroit, Feb.

1976.

32. M. G . Kofskey, and W. J . Nusbaum, "Cold-Air Performance of a Free Power Turbine Designed f o r a 112-kW Automotive Gas Turbine

Engine; I - Design S t a t o r Blade Chord S e t t i n g

Angle of 35 Degrees." NASA TP-1007, 1977.

33. T. P. Moffitt, W . J. Whitney, and H. J .

Schum, "Performance of a Sinkle-Stage Turbine as Affected by Variable S t a t o r Area."

N A S A T M

X-52553 , 1969. Evans and Miller

34. "Computer Program for the Analysis of Annular Combus t o r . Northern Research and En-

gineerLng Corporation, Vols. I and 11 ." N A S A 27

CR-72374 and N A S A CR-72375, 1968.

35. R. R. Tacina, and J. Grobman, "Analysis of Total-Pressure Loss and Airflow D i s t r i b u t i o n f o r Annular Gas Turbine Cmbustors." NASA TN D-5385 , 1969.

36. C. T. Norgran, "Comparison of Primary- Zone Combustor Liner Wall Temperatures with -1- c u l a t e d Predictions." NASA "F4 X-2711, 1973.

3 7 . J. S. Fear, "Preliminary Evaluation of a Perforated Sheet Film-Cooled Liner i n a Turbo- jet Combustor." N A S A R ( X-52705, 1969.

38. D. B r i e h l , and L. Papathakos, "Use of an Air-Assist Fuel Nozzle t o Reduce Exhaust Emissions from a Gas-Turbine Combustor a t Siau- lated I d l e Conditons." NASA TN D-6404, 1971.

39. R. D. Ingego, and C. T. Norgren, "High- Pressure Combustor Exhaust Emissions w i t h Im- proved Air-Atomizing and Conventional Pressure- Atomizing Fuel Nozzles. NASA TN D-7154, 1973.

40. A. H. Lefebrre, #'Lean Premixed/ Prevaporfzed Combustion." NASA CP-2016, 1977.

41. C. J. Marek, and L. C. Papathakos, "Exhaust Emissions from a Premixing, Prevapori- z i n g Flane Tube Using Liquid Jet A Fuel." N A S A T M X-3383, 1976.

42. D . Anderson, "Effects of Equivalence Ratio and Dwell Time on Exhaust Emissions from an Experimental Premixing Prevaporizing Burner."

NASA TM X-71592, 1975.

43. A . C. Antoine, "Synthesis and Analysis of Jet Fuels f r m Shale O i l and Coal Syncrudes."

N A S A TM X-73399, 1976.

44. T. W. Reynolds, "Thermal S t a b i l i t y of Some A i r c r a f t Turbine Fuels Derived from O i l S h a l e and Coal ." NASA TM X-3551, 1977.

45. R. D. Ingebo, and C. T. Norgren, "Com- b u s t o r Exhaust Emissions with A i r - A t m i s i n g Splash-Groove Fuel I n j e c t c r s Burning J e t A and Diesel Number 2 Fuels." NASA TM X-3?55, 1975.

46. R. A . ?m!ey, and E . E . Kernle, "Tech-

nology f o r Reducing A i r c r a f t Engine P o l l u t i o n ."

NASA T M X-71670, 1975.

47. N . R . Marchionna, L. A . Diehl, and A . M. Trout, "The E f f e c t s of Water I n j e c t i o n on Nitric Oxide Emissions of a Gas Turbine Camous- t o r Burning A S T M Jet-A F u e l . " NASA T M X-2958, 1973.

48. C. J. Merek, and R . P. Tacina, "Effect of Exhaust Gas R e c i r c u l a t i o n on Emissions from a Flame-Tube Combustor Using Liquid Jet A Fuel.'' Evans and Miller NASA T M X-3464, 1976.

49. G . A . K r s f t , "Preliminary Evaluation of a Heat Pipe Heat Exchanger on a Regenerative Turbofan." NASA TM X-71853, 1975. 28 50. N . R. Marchionna, "Effect of Increased of Oxides of N i - Fuel Temperature on Emissions trogen from a Gas Turbine Combustor Burning ASTM Jet A Fuel." N A S A TM X-2931, 1974.

51. H . F. Butze, and D. A. Kendall, "Odor Intensity and Characterization Studies of Exhaust From a Turbojet Engine Combustor TM N A S A X-7 1429 , 1973.

52. D. N. Anderson, R. R Tacina, and T. S.

Nroz, "Catalytic Combustion Tcc cne htomotive Gas Turbine Engine." N A S A TM X-73589, 1977.

53. D . N. Anderson, 'lEmi~sionsand Perform- ance of Catalysts f o r Gas Turbine C a t a l y t i c SOL- bustors." N A S A T M X-73543, 1977.

54. D. N. Anderson, "Performance and Emis- sions of a C a t a l y t i c Reactor with Propane, Diesel, and Jet A Fuels." CoNS/1011-20; N A S A TM X-73786, 1977.

55. F . Stepka, C. H. Liebert, and S .

Stecura, " S u m r y of N A S A Research on Thermal-

B a r r i e r Coatings ." Paper 770343 presented a t

SAE International Automotive Engineering Con- gress, Detroit, Feb. 1977.

56. H . F. Butze, and C. H. Liebert, "Effect of Ceramic Coating of JT8D Combustor Liner on Maximm Liner Temperature and Other Combustor Performance Parameters." N A S A I'M X-73581, 1976.

57. W. A . Sanders, and H . P . Probst, "Be- havior of Ceramics a t 1200' C i n a Simulated Gas Turbine Environment .'I Paper 740240 presented a t SAE A~iomotiveEngineering Congress, Detroit, Feb. 1974.

58. T . Vasilos, and R . ?i. Cannon, Jr., "Im- proving the Toughness of Refractory Compounds .I1 AVSD-OlO8-76-RR, Avco Corporation, L o w e l l , Nov- ember 1975; a l s o NASA CR-134813.

59. R. L. Ashbrook, "Improved Performance of S i l i c o n Nitride-Based High Temperature Cer- amics .'I NASA TM-73791 , 1977.

60. P . F . Sikora, and H . C . Yeh, "Consoli- dation of S i l i c o n Nitride Without Additives."

NASA TM X-73693, 1976.

61. W. H . Rohodes, and R . M . Chnnon, J r . , "High-Temperature Compounds for Turbine Vanes .I* AVSD-0336-72-CR, Avco Corporation, Lowell, Massachusetts, September 1972; a l s o N A S A CR- 120966.

62. J . C. Freche, "Further Investigation of Gas Turbine with NBS Body 4811C Ceramic Rotor

Blades." NACA RM-E49L07 , National Advisory Evans and Miller

Committee f o r Aeronautics, March 1950.

63. G . C . Deutsch, A . J. Mayer, and G . M.

A u l t , "A Review of the Development o f Cermets." 29 AGARD-185, Paris, Advisory Group f o r Aeronautical Research and Development, 1958.

64. C . Calvert, "Ceramic Blade-Metal D i s k Attachment." East Hartford, Connecticut, Pratt 6 Whitney Aircraft, NASA Contract NAS3-19715 ( i n process).

65. J . C. Freche, and G . M. Ault, "Pregress i n Advanced High-Temperature Turbine Materiels, Coatings, and Technology." NASA TI4 X-73628, 1977.

Evans and Miller REGENERATOR COMP TURBINE (117 r P O w E R TURBINE UI) I I

I I 1

PLANETARY GEAR BOX r -1 I

l i

figure 5. - Three shaft turbine-transmission system.

COMMANDED REAL-TIMI: CONTROL SIMULATION' VAR IAB US

------

HYBRID COMPUTER

I

DIGITAL ---- ANALOG r - L - 7 ACTUATORS

i- -- -

ACTUATOR L - , 1 , , , 1

-i-=

I + I

1 I

DIGITAL I f%NCTlON I I ~ G ~ N E I

I ACTUAL ENGINE

I GENERATION I

CONTROL

I

I SENSOR

I

L-,J

------ J

SENSED ENGINE VARIABLES

Figure 6. - Use of real-time. hybrid computer simulations for digital control development

Figure 7. - Block eagram of F l O O mathematical model.

-

HYBRID SIM.

- - -- EXPERIMENTAL

I I I

(a) COMPRESSOR SPEED.

I I I I

. 1 5 10 15 20 25 0 TIME, SEC 0)) BURNER PRESSURE.

Figure 8. - Transient verification of F U N simulation.

Idle-to-intermediate power lever slam. Altitude = 30 OOO ft, Mach number = 0.7.

NOTE: THE PREDICTED VALUES USING THE PROGRAM OF REF. 16 AGREE WITHIN 1% OF M E MEASURED VALUES.

DESIGN TOTAL EQUIVALENT PRESSURE P 6 8 10 12 14 16 18 2&1# MASS FLOW-SPEED PARAMETER.

(B HrDmCsec

I I I I I I I

. 3 . 4 .5 .6 . 7 .8 .9 l ! O MASS FLOW-SPEED PARAMEIER. &-raG/rec2 Figure 9. - Experimentally determined turbine performance map.

PARALLEL TO AXIS

_ _ ---

Figure 10. - Typical axial flow blade passage.

- HUB-TO-SHROUD

STREAM SURFACt SlREAM SURFACE (a) HUB-TO-SHROUD

,- BLADE-TO-BLADE

SURFACE M I *- W Q) BLADE-TO-BLADE SURFACE OF REVOLUTION.

Figure 11. - Typical radial flow blade passages.

PRESSURE SUCTION SUR FACE (a) BLADE AND CHANNEL PROFILES.

A SUCTION SURFACE EXPERIMENTAL DATA 0 PRESSURE SURFACE EXPERIMENTAL DATA CALCULATED BY PROGRAM (REF. 2l) 0 . 2 . 4 6 . 8 1.0

RELATIVE SURFACE LENGTH

(b, SURFACE VELOCITIES.

Figure 12 - Comparison of experimental and computed

surface velocities around stator blade.

IT' IR 6,. ._ IDENTIFICATION NUMBER 1

-1

E

Y

ORIGfNAL PAGE d

o z a

OF POOR QUA.LITY

v) m

s

01- I I ' I I I ' I I ' I l 1 ' ! ' 1 ' ' " 1 L l l i . l

1 6 1 o7

104 105 REYWLPS NUMBER, Wfpr,

FigLre 13. - Variation of loss pdrameter w i t h Re)

nolds n1Jmber.

STATOR THROAT AREA, Ath. INCHES SQUARED

I 1 I I

1 ; - 5 10-4 10-3 10-2 10-1 STATOR Tt!ROAT AREA, Ath, METER SQUARED

Figure 1 4 . - Variation of loss parameter w i t h stator throat

area.

-

3 . .U 8.89

2 .go- DESIGN

e

, , , , , , I I] 1 I

3 0 4 0 60 80 loo m m

PERCENT DESIGN REYNOLDS NUMBER

Fiqure 18. - Loss as function of Oenent design Reynolds number

at design speed.

- W l l H STATOR VANE END

CLEARANCES

m

''OD-

--- STATOR 'JANE END CLEARANCES I

Y FILLED c u

.w - *

-

F r DESIGN

I 1

a 3 n 3 5 4 0 4 5 5 0 ST.4iOR CHORD SElTING ANGLE FRGvi TANGENTIAL deg

Figure 19. - Effect of stator setting angle and

end clearance on Chrysler baseline power turbine performance.

Figure 2 0 . - Stator and rotor blade profile of LeRC variable

geometrytuurbine. ref. 3 3 .

O DESIGN SPECIFIC WORK Q U -

STATOR LOSS; I ROTOR INCIDENCE LOSS- # ROTOR L O S S WORK OUTPUT .7 (a) 1 3 % STATOR AREA TURBINE.

.70 @) DESIGN SlATOR AREA TURBINE.

1 . q

’ -1

I

I I I I i

.7 1.2 1.4 1.6 1.8 2.0 22 TOTAL PRESSURE RATIO. pvpp IC) 70% STATOR AREA TURBINE.

Overview

Key points

Document

NASA T M -7?849

N A S A T E C H N I C A L

M E M O R A N D U M

AN OVERVIEW OF AEROSPACE GAS TURBINE TECHNOLOGY

OF RELIVANCE TO THE DEVELOPMENT OF THE

AUTOMOTIVE GAS TURBINE ENGINE

by D. G. Evans and T. J . Miller

Lewis Research Center

Cleveland, Ohio 44135

TECHNICAL PAPER to be presented at the

Annual Meetirg of the Society of Automotive Engineers

Detroit, Michigan, February 27 - March 3, 1978

- Minimum impact of small-scale e f f e c t s

- Compressor - a x i a l o r c e n t r i f u g a l

- s i n g l e o r m u l t i p l e s t a g e s

- f i x e d o r v a r i a b l e geometry

Heat Exchanger - s t a t i o n a r y recuptera-

- r o t a t i n g regenerator

* Turbine - r a d i a l o r a x i a l

- s i n g l e c: m u l t i p l e s t a g e s

- f i x e d o r v a r i a b l e geometry

* Engine S h a f t Arrangements - single o r

Engine Configurations - The arrangement

SYSTEM PERFORMANCE ANALYSIS TECHNIQUES -

Steady-State Performance Program - As

Analysis - In addition to steady-state turbine

CollPBEsSOR AND TURBINES - Advances in t h e

P e r f o G n c e P r e d i c t i o n Programs - The

Blade Design Prograem - I n o r d e r t o mln-

Small-Scale E f f e c t s - Uppermost i n im- Evans and Miller

104 t o 2x106 in Reynolds number w e r e included

Variable Geometry - The e f f e c t of varying

- 7 O and +a0. The e f f e c t of varying t h e t u r -

COMBUSTORS - Ongoing programs i n combus-

Basic Research - Combustor aerodynamics

temperature prediction programs ( 3 5 ) , and im-

Advanced Concepts - The high I n l e t and Evaris and Miller

C a t a l y t i c Combustor - The combustion

Multi-Element Combustor - The principal

COATINGS AND MATERIALS - Cersmic Thermal

Barrier Zoatings - A ceramic coating system Evans and Miller

Ceramic Materials Development - 21. i t i c a l

(2500" F) , as w e l l as rapid changes i n gas

Autamotive Gas Turbine Engine - A Look a t t h e

5 . L. H. Fishbach, and H . J. Caddy, "NNEP -

Performance of Axial-Flow Turbines. P a r t I -

111 - Overall Stage Performance." NASA TN

Flav Turbaeachine o r Annular Duct, I - User's

Research Center ." Paper 72-GT-42 presented a t

Engine; I - Design S t a t o r Blade Chord S e t t i n g

X-52553 , 1969. Evans and Miller

gineerLng Corporation, Vols. I and 11 ." N A S A 27

nology f o r Reducing A i r c r a f t Engine P o l l u t i o n ."

B a r r i e r Coatings ." Paper 770343 presented a t

Blades." NACA RM-E49L07 , National Advisory Evans and Miller

I I 1

l i

------

I

DIGITAL ---- ANALOG r - L - 7 ACTUATORS

i- -- -

ACTUATOR L - , 1 , , , 1

-i-=

I + I

1 I

DIGITAL I f%NCTlON I I ~ G ~ N E I

I ACTUAL ENGINE

I GENERATION I

I

I

I

I SENSOR

I

------ J

Figure 6. - Use of real-time. hybrid computer simulations for digital control development

-

- - -- EXPERIMENTAL