Effect of the Fuel Type on the Characteristics of Air-Breathing Engines.pdf

Combustion, Explosion, and Shock Waves, Vol. 37, No. 4, pp. 387{394, 2001

Eect of the Fuel Type

on the Characteristics of Air-Breathing Engines

V. V. Shumskii 1

UDC 629.7.036:536.46

Translated from Fizika Goreniya i Vzryva, Vol. 37, No. 4, pp. 25{33, July{August, 2001.

Original article submitted May 25, 2000.

Experimental studies of air-breathing engines operating on gaseous and liquid fuels

are analyzed. It is shown that, for ight conditions with a Mach number 5 at the

ground level, the results on operation and thrust{economic characteristics, which

were obtained for hydrogen-powered engines, may be used directly in development of

the ducts of hypersonic air-breathing engines operating on liquid organoborn fuels.

Gaseous, liquid, or solid fuels may be used in hy-

personic air-breathing engines depending on their pur-

pose. High-enthalpy wind tunnels used for experimental

studies can ensure full-scale parameters of the incoming

air ow, which allows testing models with both gaseous

and liquid fuels. Nevertheless, researchers often pre-

fer to use gaseous hydrogen even for models of those

air-breathing engines that operate on liquid fuels under

ight conditions. Owing to the high reactivity of hydro-

gen, this allows one to perform heat addition to the test

gas comparatively easily and to study the characteris-

tics of the cycle on a convenient model fuel. However, it

is not clear to which extent the results on operation and

thrust{economic characteristics obtained for hypersonic

air-breathing engines working on gaseous hydrogen may

be used for analysis of engines operating on a liquid fuel.

The main features of operation and thrust{

economic characteristics, which were obtained in study-

ing small-scale models of air-breathing engines tested

both on gaseous and liquid fuels [1{4], are analyzed in

the present paper.

The models consisted of an inlet, a combustion

chamber, and a nozzle. The outer surface of the mod-

els had the form of a cylinder to avoid generation of

a longitudinal force by disturbances in the test section

(the models were not located completely in the Mach

diamond). In this case, the outer surface experienced

only the action of the friction force and the wave drag

force acting on the front part of the cowl lip, which was

located within the Mach diamond in balance tests. To

reduce the wave drag, the minimum possible excess of

the outer diameter of the model (d m ) over the diameter

of the air stream entering the model (d 0 ) was chosen

from the constructional reasoning. The results consid-

ered below were obtained for two axisymmetric models:

1) d m = 30 mm and d 0 = 23 mm; 2) d m = 84 mm and

d 0 = 73:5 mm. Descriptions of the models, their loca-

tion in test sections of wind tunnels, and test conditions

can be found in [1] and [4], respectively.

The models had dual-mode combustors to perform

heat addition to supersonic and subsonic ows in one

duct. The combustors had no igniters or special ame

stabilizers.

In the combustor with d m = 30 mm (Fig. 1), the

fuel could be injected either in the rst row through

eight uniformly distributed orices (nozzles) at an an-

gle ' 1 , or in the second row through six uniformly dis-

tributed nozzles at an angle ' 2 , or in both rows simul-

taneously. In the latter case, the ratio of fuel injection

through the rst and second rows was 60 : 40. The ori-

ce diameters were 0:2 mm for liquid fuels and 1 mm

for gaseous hydrogen. The same gure shows the dis-

tribution of the cross-sectional area of the duct (cross

sections 2, 4, and 5) along the model and the ducts

used for the numerical analysis (dashed lines). The dif-

ference between the real and calculated ducts is that

the numerical conguration does not contain a number

of constructional features than can be found in actual

model ducts, such as injectors, struts, etc. Nevertheless,

the basic geometric relations (f 2 = F 2 =F 0 , F ch = F 5 =F 2 ,

F 5 =F 0 , and F a =F 0 , where F i is the area of the corre-

1 Institute of Theoretical and Applied Mechanics,

Siberian Division, Russian Academy of Ssiences,

Novosibirsk 630090.

0010-5082/01/3704-0387 $25.00 c 2001 Plenum Publishing Corporation

387

388

Shumskii

Fig. 1. Cross-sectional areas of the combustor (cross

sections 2, 4, and 5) along the model with d m =

30 mm (the dashed lines show the combustor duct

used for the numerical analysis).

Fig. 2. Cross-sectional areas (cross sections 2, 4,

and 5) along the combustor of the model with d m =

84 mm.

sponding cross section of the model duct) of the model

and numerical scheme were identical.

The experiments were performed in a high-enthalpy

facility [2, 3], both with parameters of the incoming air

ow decreasing during the test regime (pulsed mode)

and with constant parameters of the air with a Mach

number of the incoming ow M in = 5 and reproduction

of air parameters for altitudes H = 0{25 km. Typical

test parameters are given in [1, 3, 4].

To eliminate the eect of model dimensions on the

measurement results, the inuence of the fuel type was

analyzed for a model whose dimensions were approx-

imately three times greater (see Fig. 2, where d 0 =

73:5 mm and d m = 84 mm) as compared to that shown

in Fig. 1, where d m = 30 mm.

Two types of injectors were used in this model:

| Injectors with counterow injection of the fuel

(Fig. 3); they were mainly used for injection of gaseous

hydrogen and only in some experiments for injection of

the liquid fuel;

| Injectors with fuel injection at an angle of 135

to the model axis (Fig. 4); they were used for injection

of the liquid fuel.

Each injector covered 1/48 part of the combustor cross

section. For the combustor with F ch = 2:4, the min-

imum cross-sectional area is located at a distance of

53 mm downstream of cross section 4, where the struts

become thicker. However, since this blockage in the

combustor is close to injectors, the combustion com-

pleteness is not large at this length. Therefore, the

throat section of the duct both for

Fig. 3. Injectors with counterow fuel supply: 1) air

ow direction; 2) 48 orices 0.3 mm in diameter for

hydrogen injection or 48 orices 0.2 mm in diameter

for liquid fuel injection.

F ch = 2:4 and

Fig. 4. Layout of liquid fuel injection in the com-

bustor: 1) air ow direction; 2) 48 orices 0.2 mm in

diameter.

F ch = 2:15 is actually cross section 5.

Eect of the Fuel Type on the Characteristics of Air-Breathing Engines

389

OPERATION PROCESS

Completeness of Combustion Based on

Balance-Measurement Results. The procedure of

determining the combustion eciency based on balance-

measurement results can be seen from Fig. 5, which

refers to the model of diameter d m = 30 mm operating

on gaseous hydrogen. In Fig. 5, X and F m are the drag

and thrust forces, respectively, measured by the bal-

ance, q in is the dynamic pressure, and T in is the static

temperature of the incoming ow. The balance mea-

sures the total force (X or F m ) applied to all surfaces of

the model (both outer and inner ones) in experiments

without hydrogen injection into the model (curve 1 in

Fig. 5) and in experiments with injection of hydrogen

with an air-to-fuel ratio 0:6 (curves 2 and 3 in

Fig. 5). After that, the portion of this force applied to

the inner surfaces only is revealed (force R).

In Fig. 5, the total force coecient C F m measured

by the balance is plotted by curve 3, which passes

through experimental points. The internal thrust of

the model, i.e., the force applied to the inner surfaces

of the model duct only is greater than the force F m

measured by the balance; the dierence equals the drag

of the cowl lip, which consists of the wave drag of the

cowl lip (X wave ) and the friction drag on the outer sur-

faces of the cowl (X fr ). Both components can be calcu-

lated: C X wave is determined from the two-dimensional

ow past the compression surfaces of the inlet and the

front part of the cowl lip and C X fr is calculated from

the at plate data.

If we add the drag coecient of the cowl lip and

the thrust coecient, which is lost because of the heat

losses to the walls, to the thrust coecient of the model

C F m measured by the balance (curve 3 in Fig. 5), then

region 5 yields the coecient of internal thrust gen-

erated in experiments by the model duct with actual

values of combustion eciency. Thus, region 5 covers

experimental values of the internal thrust coecient C R

of the model, which were obtained by correcting the

thrust coecient C F m measured by the balance. The

values of C R from region 5 may be compared now with

the calculated values of C R .

There is some uncertainty in comparison of experi-

mental and numerical values of C R ; it is associated with

the approximate knowledge of the level of losses in the

model, which may be characterized by the momentum-

loss coecient (or velocity coecient ' n ) at the model

exit. However, parametric calculations with these losses

varied showed that the uncertainty in the losses had

a weak eect on determination of the combustion e-

ciency. The reason is that the momentum of air leaving

the model in experiments for an incoming-ow Mach

Fig. 5. Thrust coecient: curve 1 shows the drag

of the model in experiments without hydrogen injec-

tion; curve 2 shows the change in the force applied to

the model due to hydrogen burning; curve 3 refers to

experimental values of C F m for = 0:55{0:65; curve

4 shows the drag of the cowl lip and the loss of thrust

due to heat transfer; region 5 covers the coecient

of internal thrust C R of the duct.

Fig. 6. Internal thrust coecient of the model oper-

ating on the liquid fuel: the curves refer to the cal-

culation for h:ch: = 0:9, ' ch = 0:98, and = 0:8 (1)

and 0.7 (2); the points refer to the experimental data

for ' 1 = 135 (3) and 45 (4).

number M in = 5 was signicantly greater than the mo-

mentum of air entering the model. Momentum losses

(even those determined rather roughly) were only an

insignicant fraction of the dierence between the in-

put and output momenta, which is the governing factor

for the model thrust.

The combustion eciency of hydrogen in the

model, which was determined by the balance, was 0.9{

0.95 for the entire range of temperatures (280{180 K)

of air incoming onto the model.

The experiments on the model with d m = 30 mm

with injection a liquid organoboron fuel in the front

390

Shumskii

row [5, 6] (for constant parameters of air incoming onto

the model) showed that the combustion eciency deter-

mined by force measurements was 0.85{0.95. Points 3

and 4 in Fig. 6 are the experimental values of the inter-

nal thrust coecient (the force measured by the balance

plus the drag of the outer surfaces of the model but mi-

nus the increment in C R , which may be obtained due

to heat losses to the duct walls); h:ch: is the total pres-

sure recovery coecient due to hydraulic losses in the

combustor (local losses and friction losses). Comparing

points 3 and 4 with the calculated curves 1 and 2, we can

see that the experimental data correspond to the heat

released during combustion of the liquid organoboron

fuel with a combustion eciency of 0.7{0.8. Taking

into account that 15{20% of heat released during com-

bustion of the fuel with = 1 was lost to model-duct

walls in the high-enthalpy wind tunnel [7], the physical

completeness of combustion of the liquid fuel is rather

high: at least 0.85{0.95. High values of the combus-

tion eciency of the liquid fuel were observed within

the entire examined range of F ch = 2:85{1:96 for the

free-stream temperature varied within 50 C and com-

bustor lengths within 130{180 mm.

The identically high combustion eciency in exper-

iments with gaseous hydrogen and the liquid fuel, and

also for the models with d m = 30 and 84 mm was a

consequence of the following factors:

| reproduction of full-scale parameters of air incom-

ing onto the model for conditions with M in = 5 at low

altitudes, which ensured self-ignition and high burning

rates of the fuels considered for moderate combustor

lengths (130{180 mm);

| eective mixing of fuels with air, which was attained

in the models.

Duration of the Transitional Regime. Three

ow regimes are possible in a dual-mode combustor,

depending on the relative heating of the test gas =

T 0;5 =T 0;in , where T 0;5 is the stagnation temperature at

the combustor exit (cross section 5 in Figs. 1 and 2)

and T 0;in is the stagnation temperature of the incoming

ow.

Regime 1. < switch (heat is added to a supersonic

ow as a whole). The value of switch is understood

as the relative heating of the test gas, such that heat

addition to a supersonic ow is changed to heat addition

to a subsonic ow.

Regime 2. switch < 6 max (heat is added to a

subsonic ow as a whole). The value of max is the max-

imum relative heating of the test gas, which is possible

for a given combustor geometry without stalling of the

air inow into the model. At = max , the transition

of a supersonic ow to a subsonic ow should occur in

the inlet throat (cross section 2 in Figs. 1 and 2).

Fig. 7. Static temperature of the incoming ow at which

ow reconstruction in the combustor occurs: heat addi-

tion to a supersonic ow (1) and to a subsonic ow (2).

Regime 3. > max (stalling of the air inow into

the model). In this regime, the normal inow of air into

the model cannot be provided.

Figure 7 shows the temperature at which ow re-

construction in the combustor occurs as a function of

the level of hydraulic losses in the combustor (in model-

duct section 2{4{5). The point B refers to the tempera-

ture of switching of ow regimes for h:ch: = 1: for val-

ues of T in below and above the point B, heat is added

to a subsonic ow and a supersonic ow, respectively.

If h:ch: < 1 but it lies on the right of the line CD, heat

is added to a subsonic ow for values of T in above the

line BC. This means that the line BC for h:ch: < 1,

as well as the point B for h:ch: = 1, corresponds to

the temperature of switching of heat-addition regimes

in the combustor.

If the values of h:ch: are on the left of the line CD,

heat addition to a subsonic ow in the combustor is

impossible, which follows from the ow-rate equations

written for cross sections 2a and 5 (2a is the cross section

behind the normal shock located in the inlet throat).

It is seen from curve 3 in Fig. 5 that the line

C F m = f(T in ) has an inection. The reason is that ow

reconstruction in the combustor takes a certain time.

Indeed, a supersonic ow is formed in the model duct

after inlet starting. Since the time of ow stabilization

in the inlet is 2 msec [8], we may assume that the

rst portions of the fuel enter the combustor when a

supersonic air ow has formed (or has almost formed).

Therefore, self-ignition and combustion of the rst por-

tions of the fuel occur in a supersonic air ow. The heat

released thereby decreases the Mach number of the su-

personic ow in the combustor and increases T and p,

i.e., improves conditions for self-ignition and combus-

tion of the subsequent portions of the fuel. This, in

turn, leads to an even greater decrease in the Mach

number and an increase in p and T in the combustor,

Eect of the Fuel Type on the Characteristics of Air-Breathing Engines

391

etc. This process continues until some stable regime is

established (or some quasi-steady regime, since the pa-

rameters p in , T in , q in , and T 0;in decrease). Depending

on the heat addition available, the total stagnation tem-

perature of the incoming ow, and the level of losses in

the combustor, this stable regime is formed either in a

supersonic ow as a whole or in a subsonic ow as a

whole. The calculations show that a subsonic ow is

formed in combustors of the models considered under

test conditions used.

Thus, during the operation regime of the wind tun-

nel, the test-gas ow in the model combustor should

change from a supersonic ow established during the

rst 2 msec of wind-tunnel operation, when there is

no fuel injection, to a subsonic ow. The experiments

showed that the time of this change determined by

thrust measurements is 20{28 msec.

In Fig. 5, the transitional process in the combustor

is determined by measuring forces applied to the model.

The time of the transitional process determined by pres-

sure measurements in the combustor is 20{25 msec,

which is in good agreement with that determined by

force measurements. Thus, the analysis of pressure in

the models also conrms the fact observed in force mea-

surements and numerical analysis: heat is added to a

subsonic ow as a whole after completion of the transi-

tional process.

A thermodynamic analysis of the change in heat-

removal regimes was performed independent of the fuel

type: the basic parameter was the relative heating of

the test gas in the model duct. The experiments with

the models conrmed this fact: the duration of the tran-

sitional period from the supersonic process of heat ad-

dition, which is established during the rst milliseconds

of wind-tunnel operation, to the subsonic process is the

same for models operating on hydrogen and the liquid

fuel.

addition, was too close to cross section 2. An increase

in the length of the initial section of the combustor in

all models allowed a \delay" in stalling of the inow

toward the region of large relative heatings of the test

gas.

In experiments with constant parameters of the in-

coming ow, the relative heating remained constant.

Therefore, if the combination of the length of the initial

section of the combustor, F ch , and the relative heat-

ing was such that the beginning of the pseudoshock

reached the external compression surface of the inlet,

then stalling of the air inow occurred in the beginning

of the test regime.

Thus, in testing models operating both on hydro-

gen and the liquid fuel, the mechanism of stalling of

the air inow into the model is the same. The mea-

sures taken to prevent inow stalling were also identical.

These measures included increasing the relative length

of the initial section of the combustor.

Eect of the Combustor on Inlet Operation.

If disturbances from the combustor, which are induced

by a pressure increase in the latter, are transferred up-

stream of cross section 2 to compression surfaces of the

inlet, the combustor aects inlet operation, since the

transfer of disturbances to compression surfaces usually

leads to stalling of the inow into the inlet. If the dis-

turbances from the combustor are not transferred up-

stream of cross section 2 to compression surfaces, then

combustor operation does not aect inlet operation.

The transition from supersonic to subsonic ow in

the duct for M > 2 occurs in the pseudoshock. The

pseudoshock has a certain length and may be located

in this or that place of the combustor, depending on the

relative heating of the test gas in the model duct. Such

a pattern of pressure increase is clearly demonstrated by

the curves of pressure distribution along the combustor.

The calculations of force characteristics and test-

gas parameters along the model duct involve usually

the normal shock rather than the pseudoshock; the nu-

merical parameters of the subsonic ow behind the nor-

mal shock are assumed to coincide with the parameters

behind the pseudoshock [9, 10]. The estimate of the

transition point from supersonic to subsonic ow, which

was obtained using a simplied scheme including the

normal shock, shows that stalling of the air inow into

the model occurred if the distance between the normal

shock and cross section 2 was smaller than the pseu-

doshock length. It follows from here that there should

be some extra length behind the inlet throat, and the

pseudoshock beginning should be located at the initial

section of the combustor and should not be shifted up-

stream of the corner point of the inlet.

Stalling of Air Inow into the Model. For

models being tested in a pulsed regime (decreasing tem-

perature of the incoming air), stalling of the air inow

into the model was observed. For dierent fuels and

equivalence ratios, stalling occurred at dierent times

counted from the beginning of the test regime in the

wind tunnel and dierent values of T 0;in . However, the

value of the complex

T 0;5 =T 0;in

immediately before stalling was signicantly smaller

than the limiting value possible before thermal chok-

ing for combustor congurations tested. The reason for

stalling was the fact that the zone, where the ow tran-

sition from supersonic to subsonic occurred due to heat

T 0;5 =T 0;in at the moment of

stalling of the air inow into the model remained con-

stant in all experiments ( is the increase in the test-gas

mass due to fuel addition). The value of

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