Study of LH2-fueled topping cycle engine for aircraft propulsion

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Technical Memorandum

ASS

'hopping Cycle

Study of a LH2-Fueled

Aircraft Propulsion

Engine for

(NASA-TM-83466) STUDY OF 1H2-FUELLC ICEPING N83-34942 CYCLE ENGINE FOR AI RCEA FT EECP ULS IO 14 (NASA) 21 P HC AG 2 /14F AO i CSCL 21E Unclas G3/07 42052

George E. 'gurney and Laurence H. Fishbach

Lewis .Research Center

Cleveland, Ohio

STUDY OF A LH2-FUELED TOPPING CYCLE ENGINE ORIGNAL F'*' FOR AIRCRAFT PROPULSION OF POOR QUALI V By George E. Tuiney and Laurence H. Fishbach National Aeronat,;.ics and Space Administration Lewis Research Center Cleveland, Ohio 44135 Abstract P total pressure, psia An analytical investigation was made of a topping cycle aircraft engine system which uses a heat transfer rate in heat Btu cryogenic fuel. This systeii consists of a main QHX-1 exchanger NX-1, hir turboshaft engine which is mechanically coupled (by cross-shafting) to a topping loop which aug- heat transfer rate in heat Btu ments the shaft power output of the system. QHX -2 exchanger HX -2, -lir The thermodynamic performance of the topping BtuR cycle engine was analyzed and compared with that S entropy,

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of a reference (conventional-type) turboshaft engine. For the cycle operating conditions w SP shaft power output of engine, selected, the performance of the topping cycle horsepower Shp) engine in terms of brake specific fuel consumption (bsfc) was determined to be about 12 percent bet- T `R temperature, ter than that of the reference turboshaft engine.

Engine weights were estimated for both the m W air flow rake a topping cycle engine and the reference turboshaft engine. These estimates were based on a common shaft power output for each engine. Results lb m indicate that the weight of the topping cycle bleed air flow rate, W bleed hr ngine is comparable to that of the reference e turboshaft engine.

lb W f fuel flow rate, hr Nomenclature.I lb bsfc brake specific fuel consumption, W• flow rate of component i in flow hr mixture (equation (7)), i lb fuel h hr P- y specific heat ratio for air C p,a specific heat of air O R— n m- thermal efficiency m at constant pressure, X fraction of total fuel which is turbomachine efficiency e burned in burner B2 Btu Subscripts: H enthalpy, m N 0 refers to atmospheric free-stream conditions heat of combustion of hydrogen fuel, All comb 51500 lb 19 flow station designations 1, 2,..., 3 and 4) m (see Figs.

& Btu id ideal enthalpy of component i in H i gas mixture, m j refers to flow station j Btu enthalpy of gas mixture H 1 refers to flow station 1 mix m 3 3 (equation (7)), ` introduction K conversion factor (equations (4) and (6)), The problem of meeting the nation's energy needs in the coming years has become a subject of 3.927 x 10_ 2h% increasing concern. The escailating worldwide i I; demand for energy is now causing an accelerated Ma Mach number depletion of some of the world's key energy resources.

In the foreseeable future, this problem could become acute for some resources, This paper is declared a work or the U.S.

Government and therefore Is in the public domain.

i a

particularly petroleum - which is the basic re- this study, assessments were also made of the source for production of all aviation fuels. weight associated with this engine system. Re- sults of the thermodynamic analysis and the engine Figure 1 (constructed from published datal) weight analysis are described in this report.

shows projected depletion trends in petroleum This engine study is based on the use of liquid resources for three different annual consumption growth rates. It is apparent from this figure hydrogen fuel. However, in principle, the topping that the world's oil resources are limited and cycle engine could operate with other cryogenic could be depleted in a relatively short period of fuels, such as liquid methane.

time. When crude oil production rates are con- sidered, it appears that supply shortages could develop long before the crude oil reserves are Description of Topping Cycle Engine depleted. For instance, Fig. 2 shows an estimate As an aid in explaining the topping cycle of the maximum proruction rate of crude oil in future years. The maximum production rate shown engine, consider first a simplified version of a here was obtained from published estimates ) and hydrogen fueled turboshaft engine such as shown in represents the maximum rate at which crude oil can Fig. 3. Air entering this engine is compressed by be supplied to the consumer market. At a zero- the compressor (Cl) and is then mixed with the hydrogen fuel pumped from the storage tank. The demand growth rate, Fig. 2 indicates that short- mixture is burned rind the combustion products are ages may develop in about 32 years. And with a 2 expanded through the turbine to produce work. The percent growth rate, shortages may be felt in about 15 years. work produced by , he turbine drives the main com- pressor and fuel pump and also powers an external The depletion of crude oil resources has load. The external load in this system could be a prompted the NASA and o'.hers to investigate fan, a propeller or some other propulsive compo- nent.

alternate energy resources for the production of aviation fuels. Studies conducted by NASA have Figure a concluded that liquid hydrogen (LH2), liquid shows the same turboshaft engine with i methane (LCH4) and synthetic aviation kerosene a topping loop added. In this system, a fraction of the compressor airflow from the main engine is (or synJet) are the three most promising alternate bled off at an interstage point and fed to the fuels for aviation. All of these fuels can be topping loop. The topping loop in this system produced synthetically from coal - one of our most abundant fossil fuel resources. operates as an auxiliary power unit. Tta bleed air entering this loop is precooled and compressed Liquid hydrogen (LH 2 ) and liquid methane in two separate processes. Preceoling is accom- (LCH 41 are cryogenic fuels which have been plished with hydrogen-to-air heat exchangers as studied extensively for both subsonic and super- shown in rig. 4. Precooling is advantageous in sonic transport aircraft 3, . The advantages and that it lowers the air temperature ahead of the compressors thereby reducing the compression work.

disadvantages associated with th- use of cryogenic fuels for transport aircraft are well known.

The total engine fuel flow is fed directly into Storage and handling are considered to be a prime drawback. But the cryogenic state of these fuels the top,,ing loop burner. The topping loop burner can be used advantageously to improve the perfur- operates highly fuel-rich with an equivalence mance of aircraft propulsion systems. For exam- ratio (ER) ranging somewhere between about 2,5 and ple, cryogenic fuels can be used effectively as a 4.5. The equivalence ratio is defined as follows: heat sink to reduce or eliminate the need of com- pressor bleed air for to rhine cooling. Other Actual fuel-to-air ratio ER schemes for improving engine performance have been to c iometr c ^ fuel-t -_ ratio investigated . They include: using cryogenic fuel as a heat sink to precool the compressor air- Thus, only a fraction of the total fuel flow is flow, expanding heated fuel through an auxiliary burned in the topping loop, turbine to produce additional shaft work and pre- heating a portion of the fuel before burning in a The combusticn products from the to p ping loop combustor.

include hydrogen - the unburned fuel, nitrogen and water vapor. (All oxygen is consumed in the fuel- In this report, a topping cycle gas turbine rich burning process.) The fuel-rich combustion engine is introduced which uses cryogenic fuel and products are expanded through the topping loop which combines all of the above mentioned schemes turbine and ducted directly into the main engine for augmenting engine performance. This engine burner. There, the fuel mixture is combined with (hereafter referred to as the "topping cycle the remaining eA r from the main engine, burned engine") consists of a main turboshaft engine completely and expanded through the main engine which is augmented by a direct-coupled secondary turbine. Typically, the main engine burner oper- ates at an equivalence ratio (ER) ranging from power generation loop (or topping loop). The top- N ina about 0.2 to 0.5.

loop operates with precooled compressor air and a fuel-rich combustor. The fuel-rich combus- As indicated in Fig. 4, the topping loop and t tion products are expanded through an auxiliary turbine in this loop and are then fed into the the main turboshaft engine are interconnected by a burner of the main turboshaft engine where the cross-shaft so that power produced by the topping excess fuel i s burned completely loop turbine is supplied to the compressors and to the main engine load. Thus, the topping loop The study presented in this report daals pri- operates as an auxiliary powerplant and produces marily with the thermodynamic performance charac- additional power which is transferred to the sys- teristics of this engine system. As a part of tem load. The performance benefits of this cycle

OF POOR QUALITY

are described in the Results section of this Figure 5 is a temperature-entropy diagram for report.

the topping cycle engine operating at the assumed design point conditions of Table 1. This figure shows the prQrcures and temperatures at each of Analytical Approach the major points throughout this engine. As indicated in Fig. 5, both burners in this engine The primary purpose of this study was to evalu- system operate at a design point temperature level ate and assess the thermodynamic performance of of 3000° R. This temperature level is considered the topping cycle engine. But, in additi to the to be representative of an advanced technology thermodynamic or cycle analysis, engine component engine.

and system weights were also evaluated.

The bleed air entering the topping loop is As a first step in the approach to this study, taken from an interstage point of compressor C1.

a reference or "baseline" engine was established The pressure of the bleed air was assumed to have to serve as a standard by which the topping cycle a design value of of 81.0 psia. This pressure engine could be compared. The baseline engine is level was chosen to provide nearly-equal heat the hydrogen fueled turboshaft engine depicted in transfer rates in the individual heat exchangers Fig. 3 and described in the preceding section. of the topping loop.

For the thermodynamic analysis, analytical In the analysis of this engine system, a pres- models representing both the baseline engine (Fig.

sure matching constraint was imposed such that the 3) and the topping cycle engine (Fig. 4) were exhaust stream pressure from turbine T2 was equal developed. All engine performance calculations for to the stream pressure in the main engine burner this study were based on cruise flight conditions 81.

at Mach 0.80, 35,000 feet altitude.

Cycle Analysis. Normally, in cycle analysis, To syst:matically evaluate the thermodynamic the conventional index used in rating the perfor- performance of the topping zycle engine, appropri- mance of aircraft propulsion systems is the s•.

ate equations were developed for both the topping "specific fuel consumption". For turboshaft cycle and baseline eng;.:9s. Digital codes were engines, such as the baseline engine and the then formulated for cz1 ,7 1ilating the performance r on- topping cycle engine, the brake specific fuel characteristics of these engines. Fluid thermo- su,nption, bsfc, is used and is defined as: dynamic properties for, the analysis were taken from three sources: Properties for parahydrogen, Fuel flow rate Wf rogen and oxygen were taken from Hendricks et.

( 1 ) osfc = Net shaft power delivered to load SP M . And properties f r air and steam were obtained from Fishbach^ and Hendricks et. a18.

Another index used to indicate aircraft propul- sion system performance is the thermal efficien- For the purpose of estimating engine system cy. The thermal efficiency, q, is directly re- weight , a digital computer code, called lated to the bsfc. For turboshaft engines, the WATE-2 , was used. This program is capable oP thermal efficiency is defined as: estimating weights and dimensions of individual components (such as compressors, turbines, Net shaft power deliver:d to load burners, etc.) as well as total engine system - Chemical energy input weights. The WATE-2 program was used in this study to estimate weights of both the baseline SP engine and the tapping cycle engine. A reference (2) 6 (W f) 0 shaft power outpv t of 10,000 horsepower was

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assumed as a basis fo • estimating and comparing the weights of these engine systems.

In equations (1) and (2), the net shaft power delivered to the load, SP, represents the total The analytical procedure along with the assump- turbine power developed less the power absorbed by tions g rid equations used in this study are the compressor(s) and pump. For the baseline described in the following paragraphs.

engine (Fig. 3), the net shaft power delivered to the load is: Thermodynamic Analysis Baseline Power developed Power consumed engine = by - by Cycle Conditions and Component Characteristics.

SP turbine T1 compressor C1 Cycle operating conditions and component charac- teristics were established for the purpose of Power consumed (3) providing a consistent set of reference conditions by for use in computing and comparing th y ; performance pump P of the baseline engine and topping cycle engine.

Table I is a listing of the assumed cycle Or, in terms of the respective flow rates and operating conditions kdesign poihit conditions) and enthalpies, equation (3) becomes: the component characteristics. As indicated in this table, all components which are common to Baseline both of these engines (such as inlets, com- engine = K[(W a + W f ) (H 4 — H5) pressors, burners and turbines) have identical SP

rli characteristic values. The component efficiencies

- W a (H 2 - 11 1) - Wf(H7 - H 6)] (4) listed in Table 1 are considered to be representa- tive of current day aircraft engine technology.

rf

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And for the topping cycle engine (Fig. 4), the net (8) Mat) YIY-

shaft power delivered to the load is: P 1 o P (1 + 3

o

Topping Power developed Power developed cycle = by + b

T 1

+ -2— = T o ( 1 ) (9) turbine T1 engine turbine T2 SP The enthalpy and entropy of air at station 1 were then obtained from an air prgperties subroutine Power consumed Power consumed formulated from data in NNEP . In functional - by - by notation, the enthalpy and entropy relationships compressor C1 compressor C2 are as follows: H1 ° (5) f ( P 1+ T 1) (10) and Power consumed Power consumed S1 = f ( P 1, - by T 1) (11) - by compressor C3 pump P (2) Turbomachinery - The fluid conditions at the outlet of a turbomachinery comr)onent (such as Expressed in terms of the respective flow rates a compressor, turbine, or pump) were determined as and enthalpies, equation (5) becomes: described below.

Topping Designating the outlet and inlet flow stations cycle = K [(W

a + Wf)

H5) + (H 4 - (W bleed + Wf) as "j" and " j -1 11 , respectively, the outlet fluid engine x (HH ) - (W - W ) (H H ) conditions were computed as follows. For a con- SP 14 15 a bleed 2 - 1 stant entropy ;p rocess, the ideal enthalpy at sta- H1) - - W bleed (H 8 - Wbleed (H 10 - H9) (6) tion j (Hj , id) is given by: - W bleed (H 12 - H 11 ) - W f (H 7 - H6)]

Vj, Sj- 1 )

H j,id ° f (12)

Some of the enthalpy terms in equations (4) and For the compressors and pump, the actl-el (6) represent a weighted average for two or more enthalpy at flow station j is: component fluids. For example, the enthalpy at flow station numbers 14 (i.e., H14) is the aver- age enthalpy per pound of fluid mixture containing = Hj-1 + (Hj,idC- Hi-I) (13) nitrogen, hydrogen, and water vapor. The enthalpy

H

of the mixture is dependent upon the composition of the flow mixture. An appropriate expression And for the turbines, the actual enthalpy at sta- was used to compute the enthalpy per unit mass of tion j is: flow mixture. In equation form, the enthalpy of a unit mass of fluid consisting of a mixture of n (14) + E(Hj,id - Hj-1) Hj " Hj -1 component gases is: Fluid temperatures (corresponding to the actual i Component n enthalpies and pressures at station j) were then obtained directly from a subroutine of fluid prop- E(Wi Hi) erties.

i = Component 1 (3) Burners - The stoichiometric fuel-to-air (7) H mix i Component n mass ratio with hydrogen fuel is .029152. In order to limit the burner outlet temperature to a Wi reasonable preset value, the engine burners must

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be operated either fuel-rich (ER > 1.0) or air- i = Component i rich (ER < 1.0). As stat-' in the section Description Of Topping Cycie Engine, the topping Eng ine Component Analysis. In order to solve loop burner (B2) operates fuel-rich; and the main equations 4 and (6), the respective enthalpies engine burner (B1) operates air-rich.

in these equations must be determined. Starting at the engine inlet, the enthalpy terms were The rate of heat release in the respective evaluated from thermodynamic relationships ap- burners is directly proportional to the rate at plied to the specific components in these sys- which fuel is reacted.

tems. Pressures at the inlet and outlet of each component were determined from the design point (a) Baseline engine - In the baseline engine pressures and pressure ratios listed in Table (Fig. 1), all fuel entering the main burner (B1) I. The equations and procedure applied to the is reacted with excess air. And the rate of heat specific components in the engine systems are release is given by: presented in the paragraphs which follow.

Rate of heat release = ' (15)

i f aHcomb

(1) En g ine inlet - Beginning with free- in burner B1 stream c.,aditior,s at 35,000 feet altitude and assuminc, an inlet recovery of 1.0 as stated in The mass ratio of fuel flow to total air Table I, the total pressure and total tempera- flow, W f, required to achieve a specified ture at the inlet (station 1) were computed as follows: Wa

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value o • ` burner outlet temperature, T4, 19 ob P00R Q"rd burned completely. The mass flow rate tained from the heat balance equation. That is, of fuel entering the main burner (B1) is equal to (1 - X) W f . And the rate of heat release in (16) burner 81 is given by the following equation.

H 3) Wf ° Hconb ° ( Wa + W f) ( H4 - Rate of heat release o (1 - X)(Wf)(aHcomb) (20) Rearranging equation (16) and solving for the in burner 81 fue,-to-air mass flow ratio gives: b,f (H4 - H3) The mixed-mean temperature out of burner B1, i.e., T4 of Fig. 4, has a specified design point (17)

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value of 3000 R. The total airflow required to achieve this specified temperature is obtained The enthalpies of the flow mixtures, i.e., H3 from an energy balance around burner 81. Thus, and ri4, were calculated from the relationship the total airfluw (including bleed a;r to the given by equation (7). The value of H4 in equa- topping loop) is given by the following equation: tion (17) is dependent on the temperature and pressure at flow station 4 and also on the fuel- (Wf) (AFomb) X) (I - W o W (L1) to-air mass flow ratio. Thus, an iterative proce- a H3) y f dure was used to solve equation (17) for the fuel- (H4 - to-air mass flow ratio.

An iterat i ve procedure was used to solve equa- (b) Topping cycle engine - In the topping tion (21). fr,9 procedure involved assuming a cycle engine (Vg. 4), burner B2 operates fuel- value of total airflow, Wa. This, in turn, rich and the rniin burner, Bl, operates air-rich.

fixes the composition of the flow mixture entering Representing the fuel fraction burned in burner B2 and leaving the burner. Then, the respective by X, the rate of heat release in burner B2 is mixture enthalpies (113 and H 4 ) were computes, given by the equation from the relationship given by equation (7). This procedure was repeated, as required, until a solu- Rate of heat release a X(Wf) ( nH (16) comb) 1 tion of equation (21) was obtained.

in burner B2 (4) Heat Exchangers - The function of the heat.

The products of combustion from burner B2 con- exchangers in Vie topping loop is to reduce the sist of water vapor, nitrogen and hydrogen - the temperature of the airflow entering the ,orn- unburned fuel. (All oxygen supplied with the pressors. The air temperature at the outlet of bleed air entering 82 is reacted.) The mixed-mean each heat exchanger was assumed to have a design temperature out of burner B2, i.e., T4 of point value of 600' R. And each heat exchanger Fig. 4, is specified (see Table I) an^ has a de- was assumed to have a design point pressure ratio sign point value of 3000' R. By rearrangement of of 0.99 (see Table I).

the heat balance equation for hurner B2, we ob- tained the foliowing expression for the mass frac- With reference to Fig. 4, the rate of heat tion of total fuel flow which is burned in this transfer from the bleed airflow to the hydrogen burner.

flow is given by the following equations: 1 + W bleed For heat exchanger HX-1, (H14 - H13) W 1 f C X = (19) T9 AH comb p , a dT (22) An iterative procedure was used to solve Q HX-1 ° W bleed f C equation (15). The first step in the procedure was to assume a value for the fuel-to-bleed air T8 mass flow ratio, Wf . W hen, since all oxygen And for heat exchanger HX-2, Wbleed present in the bleed air entering burner B2 is T11 consumed, we independently computed a value for X from a mass balance around this burner. Using the p , a dT initial values of W f and X, we then (23) Q HX-2 ° Wbleed f C Wbleed

calculated the composition of the combustion prod- T10

ucts leaving burner B2. The enthalpy of the com- The mixed-mean temperature of the hydroger, flow bustion product mixture, H14, was then calcu- from heat exchangers HX-1 and HX-2 (i.e., T20) lated from equation (7) for a preselected design print temperature of 3000' R. This procedure was was determined by equatirg the total heat transfer repeated, as required, until a solution of equa- rate (Q X_ plus QHX-2) to the sensible heat Gained by the hydrogen flow.

tion (19) was obtained.

an analy- The fuel-rich combustion products from burner In a previously putiished report 10 , sis was presented for the heat exchangers in the B2 are expanded through the topping loop turbine topping cycle engine. In that analysis, a partic- and fed directly into the main engine burner ular compact heat exchanger core configuration was (B1). There, the excess fuel in the mixture is considered. Heat exchanger dimensions, surface combined with the remaining air from the main

areas and weights were estimated for a reference Results topping cycle engine system which produces a net shaft pt .Per of 10,000 horsepower.

Thermodynamic Performance In this section, the thermodynamic performance Weight Analysis of the topping cycle engine is presented and co^m•- pared with that of a baseline (or reference) Engine system performance and engine system turboshaft engine which operates under the same weight are key factors to consider in the cycle conditions. The engine system component comparison of aircraft propulsion systems. But ch.racteristics and cycle conditions (design some of the schemes proposed to improve aircraft operating conditions) used in this study were engine system performance also result in increases introduced in the Thermodynamic Analysis section in the system weight and complexity. In view of of this report.

this, an engine system weight study was planned as a part of the overall topping cycle engine system Figure 6 depicts the performance of the topping analysis. The procedure used in this weight study cycle engine and baseline engine for the design is described below.

operating con^.'itions given in Table I. As V lus- trated in Fig. 6, the topping cycle engine has a A digital computer program, called WATE-210, significant performance advantage over the base- was used to estimate the weights of both the line engine. At the design conditions given in baseline and topping cycle engines. In order to Table I, the thermal efficiency and brake specific provide a common basis by which these engine fuel consumption of the topping cycle engine are weights could be compared, a fixed shaft power about 12 percent better than that of the baseline output of 10,000 horsepower was assumed for each engine.

engine system.

:,overal of the design parameters in Table I The engi,ie weight program (WATE-2) is designed were varied individually for the purp o se of to functio l in conjunction with a specifl y engine examining the topping cycle engine pe-formance at cycle analysis program, known as the NNEP I . As other cycle conditions. The effects of these a prerequisite to operating the WATE-2 program, a individual parametric changes on englna system digital simulation must be made for each engine performance are presented in the following system using the NNEP. Input to the NNEP includes paragraph,.

engine thermodynamic data and component data along with logic and control data which describe the Effect of Com reGsor Pressure Ratio. Figure 7 order, function, and arrangement of components in

for the

shows the thermal e f ficiency and

the engine systen.

baseline and topping cycle engines against the pressure ratio of the main compressor (Cl). Data The thermodynamic output from the NNEP (which in this figure were computed based on a constant includes temperatures, pressures, flow races, design point Durner outlet temperature of rotor speeds, shaft powers, etc.) is fed .,erectly 3000° R. With reference to Fig. 7, the topping into the WATE-2 program. There, weights and cycle engine shows a significant performance dimensions are computed for each of the major com- advantage ove r the baseline e,igine throughout the ponents in the engine system. The total engine range of compressor pressure ratios shown. The system weight is then determined by summing the relative difference in the curves of Fig. 7 ranges weights of the components along with calculated from about 12 percent at the design point pressure weights of structural elements (such as frames, ratio (PR a 43.5) to about 20 percent at a pres- cases and support hardware) which are included in sure ratio of 15.0.

the engine system.

Effect of Burner Temperature. The design point A separate analysis was made to estimate the value of burner temperature Table I) was assumed weights and dimensions of the heat exchangers to be 3000° R. Cycle calculations were made at (HX-1 and HX-2) in the topping cycle engine. This other values of burner temperature for the purpose was necessary because the heat exchanger analysis of determining the effect on engine performance.

subprogram in WATE-= is limited to compact heat In Fig. 8, the baseline engine thermal efficiency exchangers in which both working fluids are air.

and bsfc are presented against main compressor The details of the heat exchanger analysis are (C1) pressure ratio with burner temperature as a presented in another reportl0.

parameter. The trends shown in this figure are typical of a conventional turboshaft engine. That In the preceding paragraphs, we have given a is, cycle performance improves with increasing brief description of the procedure used to burner temperature; and the optimum or near- 2stimatc: engine system weights. The accuracy of optimum compressor pressure ratio (as shown by the engine weight estimates from the WATE-2 program, dashed line in Fig. 8) increases with burner tem- as stated 9 , is generally within t 10 percent. perature.

Results of the weight calculations for the base- line and topping cycle engines are presented and A similar plot showin g the effect r,; burner compared in the Results section of this report.

temperature on the topping cycle engine perfor- mance is depicted in Fig. 9. In the topping cycle engine, both burners, i.e., B1 and B2 of Fig. 4, are assumed to operate at the same constant tem- perature. A comparison of data in Figs. 8 and 9 indicates that the topping cycle engine perfor-

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mance is significantly better than the baseline

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engine performance at all values of burner temper- (P16 - P 19 ) and - P17)) was assumed equal (P 16 ature. With the main compressor operating at the design point pressure ratio of 43.5, the relative P 16 P16 difference in performance of these engines is to 0.01.

approximately 12 percent for ail constant values of burner temperature. At lower values of main The sensitivity of the topping cycle engine compressor pressure ratio (e.g., near 15), the performance to heat exchanger pressure losses is relative difference in performance of these shown in Fig. 11. At the design point burner engines is on the order of 15 to 20 percent for temperature of 3000° R, the data in Fig. 11 the range of operating temperatures shown in Figs.

indicates that the relative change in performance B and 9. Figure 9 also indicates that at the is about i percent for each 5 percent change in burner design point temperature of 3000 0 R. the heat exchanger pressure ratio (or 5 counts of optimum pressure ratio of the topping cycle engine pressure drop).

is near the selected design point pressure ratio of 43.5, as given in Table I.

Effect of Compressor Pressure Ratio in Topping Loop. The effect of the topping loop compressor Air in Toppinq Ef f ect of Precoolin q Comp ressor pressure ratio on engine performance is shown in Loop. Precooling of the bleed air in the topping Fig. 12. As indicatedin this figure, cycle per- O it accomplished with hydrogen-to-air heat formance improves with increases in the topping exchangers (HX-1 and HX-2) as shown in Fig. 4. As loop compressor pressure ratio. There is, how- stated in the section Description of Topping Cycle ever, a pract i cal upper limit to the compressor Engine, the purpose of precooling is to reduce the pressure ratio. At the design point, the pressure work requirements of compressors C2 and C3. The ratio across each compressor in the topping loop effect of precooling is shown in Fig. 10. In this was equal to 6.10. (See Table I). This pressure figure, the thermal efficiency and bsfc of the ratir was chosen as a design value so as to limit topping cycle engine are presented against burner the absolute pressure in the topping loop to a temperature for different values of compressor value of about 3000 psia.

inlet temperature. As show e r, the relative perfor- Yp mance of the topping cycle engine increases with From Fig. 12, it appears that addition., the level of precooling. At a burner temperature Improvements in performance may be realized by of 3000' R, i.e., at the design value, the rela- i ncreasing the compressor pressure ratio above the tive performance gain from precooling is on the design point value of 6.10. However, a relatively order of 1 percent for a change in compressor air small increase in the compressor pressure ratio inlet temperature of 100 ` R.

results in a large change in absolute pressure in the topping loop. For example, increasing the From the data in Fig. 10, it would appear that pressure ratio of each compressor in the topping further improvements in performance could be ob- loop from 6.10 to 7.0 results in a pressure tained by precooling the bleed air to a increase of about 1000 psia in the topping loop.

temperature below the design point value of 600 R. However, there is a practical limit to the amount of precooling. At the design point Discussion of Topping Cycle Engine Performance operating conditions, the heat transfer rates in both of the heat exchangers are approximately In the fora;oi n g presentation, it was shown equal and the design point effectiveness of each that the efficiency and/or bsfc of the topping heat exchanger is in the neighborhood of 65 cycle engine is about 12 percent better than that percent. Additional precooling would require a of a baseline turboshaft engine operating at the higher value of heat exchanger effectiveness same cycle conditions. Before proceeoing, it is which, in turn, would result in physically larger important to note that a significant fraction of heat exchangers. Another point which should be the net shaft power output of the toppini c;Icle considered in relation to the precooling is that a Engine is generated by the auxiliary toppir;g loop.

small amount of moisture (water vapor) is normally At the design operating conditions, the topping present in the ingested airflow, even at high loop delivers about 22 percent of the total shaft altitude flight conditions. The likelihood of the power output of this engine system. Thus, the entrained moisture condensing and freezing on the efficiency with which this loop functions can have air-side surfaces of the heat exchangers would be a significant effect on the overall performance of expected to increase w i th the amount of the topping cycle engine. In this section, we precooling. To prec"sude this possibility arid also shall consider the individual factors which Con- to have the heat exchangers operate at a tribute to the overall performance gain assoc'ated reasonable value of effectiveness, a minimuri biped with this engine system.

air temperature of 600' R was assumed as a reasonable condition for this study.

Precooling of Compressor Airflow in Topping Lgop. The precooling of airflow upstream of each Effect of Heat Exchanger Pressure Losses. At compressor in the topping loop reduces the work the design point, the pressure losses in the heat input to the compressors. Along with the com- exchangers were assumed to be 1 percent of tte pressor air precooling, the fuel is also pre"-ated inlet pressure. Thus, the design value for rsach somewhat in this process. The combined effect of of the individual ratios of pressure loss p r ecooling the compressor airflow and preheating to inlet pressure (i.e., (P 8 - P9)1 the fuel results in a cycle performance improv- (P 10 - '111)1 P8 P 10 ement of about 1 percent per 100° R change in the comp —ir inlet airflow temperature.

Utilization of Fuel as a Working Fluid. A significant net gain is real zed from the use of

fuel as a working fluid in this cycle. The of the computed weights for the baseline and topping cycle engines indicates that they are heated, unburned hydrogen fuel passing through the nearly the same.

topping loop turbine generates a large fraction of the topping loop turbine power output. At the de- In comparing these engine weights, it is sign conditions, only about 6 percent of the mass flow through the topping loop turbine is unburned important to note that the specific power of the hydrogen fuel. But, because hydrogen has a high topping cycle engine i s significantly greater than specific heat (nearly 15 times that of air), the that of the baseline engine. (Specific power as turbine work produced by the unburned hydrogen used here is dufined as the ratio of engine shaft amounts to about 48 percent of the total work power output to total engine airflow.) For a produced by this turbine. On the other hand, the fixed value of shaft power output, the total air- pumping power required toboost the pressure of flow of the topping cycle engine was calculated to the cryogenic fuel (which is stored as a saturated be about 28 percentage points less than that of liquid) is relatively small. At the design condi- the baseline engine. The lower airflow (or higher tions, only about 2.5 percent of the power output oeciflc power) results in smaller and lighter- of the topping loop turbine is used to drive the weight engine components. Thus, even though the topping cycle engine has an auxiliary flow loop turbopump. Thus, in this cycle, the tonping loop which contains additional components, the esti- operates very efficiently. In terms of the over- all performance gain associated with the topping mated total weight of this engine system is com- parable to that of the baseline turboshaft engine.

cycle engine, approximately 80 percent of this gain can be attributed to the efficient way in which the fuel is used as a working fluid in the Summary of Results topping loop of this engine.

The thermodynamic performance of the topping cycle engine was analyzed and comb. •ed with that of a Engine system Weights reference turboshaft engine which operates under The results of the weight studies for the base- the same cycle conditions. For the cycle oper- line engine and topping cycle engine are described ating conditions selected, the performance of the ^, topping cycle engine (i.e., the efficiency and/or below. Weight estimates obtained from the WATE-2 specific fuel consuoption) was dete^mined to be program and presented herein are based on a common about 12 percent better than that or a reference shaft power output of 10,000 horsepower for each turboshaft engine.

engine system. The procedure used in estimating engine system weights is described in the section The improved overall performance of this engine Weight Analysis.

can be attributed to the fact teat the auxiliary Baseline Engine. The calculated weights for topping loop in this system operates at a high the baseline engine are summarized in Table II, level of efficiency. In the auxiliary topping loop, cryogenic fuel (stored as a saturated The baseline engine is actually a two-spool liquid) is boosted to a high pressure by a turbo- engine. And the turbom-chinery weights listed in Table II include the individual high and low spool pump. Because the fuel has a relatively small components. (The requirement of two spools is specific volume, the power expended in pumping the dictated by the relatively large overall pressure fuel is also relatively small, even for the large ratio of the engine.) pressure change. The high pressure-low tempera- ture fuel is also used to precool the bleed air- The compressor and turbine weights listed in flow in this loop before compression. This, in Table II include the weights of the disc, blades, turn, reduces the work required to compress the stators, connecting hardware, frame and case. bleed airflow. Finally, the high pressure fuel is But, as indicated in Table II, inlet and exhaust heated and a portion of it is then expanded system weights are not included in the total through an auxiliary power turbine to produce engine system weight. The total weight of the additional shaft power. The combination of these baseline engine as estimated from the WATE-2 processes result in a significant improvement in program is shown in Table II to be 850 lbm. the overall performance of this engine.

Engine. The calculated weights The results of the weight comparison between Topping_ CXcle -- g cycle engine are listed in Table for the topping the topping cycle engine and the baseline turbo- III. The main gas turbine engine of this system shaft engine indicate that these engines have also has two rotating spools. nearly the same weight. The weight comparison was based on a common shaft power output of 10,000 The compressor and turbine weights in Table III horsepower for each engine.

include the same hardware items as described for It is concluded, therefore, that the topping Gearbox weights and inlet the baseline engine.

cycle engine has significantly better thermo- and exhaust system weights are not included in the The weights of dynamic performance than that of the baseline total topping cycle engine weight.

the heat exchangers (HX-1 and HX-2) were obtained turboshaft engine. And the estimated weight of the topping cycle engine is comparable to that of from a previously reported studyl0.

the baseline turboshaft engine.

Because of the relatively low mass flow rates in the topping loop, the compressors in this loop Keferences (C2 and C3) were assumed to be centrifugal ma- '` chines. The total weight of the topping cycle 1. "The Global 2000 Report to the President: engine as estimated from the WATE-2 program (and Entering the Twenty-First Century," Council is 865 lbm. A comparison listed in Table III) I , i ._.__.--, --^-s-ssz-=c=+^•-_-^^^=^_-....-tea —., -.""gym:+;,-r..,_v...<f.^,^.,^..<,s...._....

on Environmental Quality, Washington, D.C., 1980.

ORIGINAL PAC.;. .-

^. Witcofski, R. D., "Comparison of Alternate

OF POOR QUALM

Fuels for Aircraft," NASA TM-80155, 1979.

Brewer, G. D., and Morris, R. E, "Study of LH- Fueled Subsonic Passenger Transport Aircraft," + ockhoed-California Co., Burbank, CA, LR-27446, Jan. 1976. (NASA CR-144935.)

4. Whitlow, J. B., Jr., Weber, R. J., and Civinskas, K.C., "Preliminary Appraisal of Hydrogen and Methane Fuel in a Mach 2.7 Supersonic Transport," NASA TMX-68222, 1973.

5. Miller, B. A., "Analysis of Several Methane - Fueled Engine Cycles for Mach 3.0 Flight," NASA TND-4699, 1968.

6. Hendricks, R. C., Baron, A. K., and Peller, I. C., "GASP - A Computer Code for Calculat- ing Oe Thermodynamic and Transport Proper- ties for Ten Fluids: Parahydrogen, Helium, Neon, Methane, Nitrogen, Carbon Monoxide, Oxygen, Fluorine, Argon, and Carbon Dioxide," NASA TND-7808, 1975.

7. Fishbach, L. H., and Caddy, M. J., "NNEP - The Navy-NASA Engine Program," NASA TMX-71857, 1.975 8. Hendricks, R. C., Peller, I. C., and Baron, A. K., "WASP - A Flexible Fortran IV Computer Code for Calculating Water and Steam Proper- ties," NASA TND-7391, 1973.

Onat, E., and Klees, G. W., "A Method to 9.

Estimate Weignt and Dimensions of Large and Small Gas Turbine Engines," NASA CR-159481, 1979.

Turney, G. E., and Fishbach, L. H., "Analysis 10.

of a Topping Cycle Aircraft Gas Turbine Engine System Which Uses Cryogenic Fuel," NASA TP-2219, 1983.

ORIGINAL PAW 15

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POOR QUALWY

TABLE I. — DESIGN POINT OPERATING CONDITIONS AND COMPONENT CHARACTERISTICS FOR BASELINE ENGINE AND TOPPING CYCLE ENGINE AT FLIGHT CONDITIONS OF MACH 0.6, 35,000 FEET Topping cycle Baseline engine engine Component or cycle parameter: 1.0 110 Inlet recovery, P1/Po Burner temperatures: Burner-81,Tq °R 3000 R ---- 3000 Burner—B2,T1 4 Efticiencies: Hydrogen pump—P .60 .60 .88 .88 Main compressor—C1 .88 Compressor—C2 ---- ---- .88 Compressor—C3 .86 Main turbine—T1 .86 ---- .86 Turbine—TG 1.0 1.0 Burner-81 1.0 1.0 Burner—B2 - Component pressure ratios: Main compressor—C1, P2/N 1 43.5 43.5 Compressor—C2, P10/Pg ---- 6.1 Compressor—C3, 6.1 P ---- 12 /P 11 Burner—B1, P 4 /P3 .95 .95 Burner—B2, .95 P 14 /P13 ---- Fuel storage (sat. liquid): 14.696 14.696 Pressure, P6, psia R 36.608 36.608 Temperature, T6, —•--- 81.0 Bleed air pressure, P8, psia °R ---- 600 Compressor—C2 inlet temp., Tg, R ---- 600 Compressor—C3 inlet temp., T 1 1, S.27 55.27 Turbine—T1 discharge press., P 5 , psia Heat exchanger pressure ratios: Air side, P9/P8 and P .99 ---- il /P 10 .99 Hydrogen side, P17/P16 and P

lg /P 18 ""'

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OF POOR

TABLE II. - SUMMARY OF COMPONENT WEIGHTS AND TABLE III. - SUMMARY OF COMPONENT WEIGHTS AND ENGINE SYSTEM WEIGHT FOR ADVANCED 10,000 ENGINE SYSTEM WEIGHT FOR 10,000 HORSEPOWER HORSEPOWER BASELINE TURBOSHAFT ENGINE TOPPING CYCLE ENGINE Component Weight, Component Weight, lbm lbm Main gas turbine engine: Compressor (Cl): Compressor (Cl): LPC 271 LPC HPC 82 HPC 101 Burner (81) Burner (B1) 81 Turbine (T l): (T1): Turbine HPT 52 HPT 69 LPT 100 LPT Shafting 22 Shafting 28 Controls and accessories 75 Controls and accessories 99 Main gas turbine engine 664 Totai engine Topping loop: Compressors (C2) 40 Compressors (0) 20 Burner (82) 22 Turbine (T2) 57 Heat exchanger(HX-1) Heat exchanger ,X-2 20 Shafting Controls and accessories Topping lo tip Total Engine

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ORIGIV I M 7t ' ' 1 14-v OF POOR QI AL c .60 z PERFORMANCE GAIN o 127o W .50 .10 - _. -- .12 z J 40 T .13 Uj .14 w a w U- TOPPING .30 U °f BASELINE CYCLE c.a — ENGINE ENGINE I W N W Q co Figure 6. - Comparison of baseline engine and topping cycle engine performance at design point operating conditions - Mach 0. 8, 35000 ft.

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OF POOR QUALITY TOPPING LOOP COMPRESSOR INLET TEMPERATURES, OR .100 .48 600 .105 .

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2200 2400 2600 2800 3000 3200 V4 BURNER TEMPERATURE, T 4 AND T 14, OR Figure 10. - Topping cycle engine thermal efficiency and bsfc against burner temperature with topping loop compressor inlet temperatures as a parameter.

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.95 w .110 x`-.90 .44 ^ it .115 — J m O DESIGN lOR I 42 o REFERENCE) POINT L 120 I NOTE: ALTITUDE = 35000 ft MACH = 0.80 32200 ' 2400 2600 2800 3000 3200 3400 BURNER TEMPERATURE, T AND T 14, OR l Figure 11. - Topping cycle engine thermal efficiency and bsfc against burner temperature with )eat exchanger pressure ratio as a parameter.