. 

NASA Technical Memorandum 103 107 

AIAA-90-1935 

On the Use of External Burning 

to Reduce Aerospace I Vehicle 

Transonic Drag 

Charles J. Trefny 

Lewis Research Center 

Cleveland, Ohio 

Prepared for the 

26th Joint Propulsion Conference 

cosponsored by the AIAA, ASME, SAE, and ASEE 

Orlando, Florida, July 16- 1 8, 1990 

(NACA-IP-103197) ON THf USE c?F EXTFRNAL rU90-2176.2 

BURNING TO REnUCr AERqSPACE VFHICLE 

TPANSnNTC WAG (NASA) 12 p C5LL 21F 

Uncl as 

G3/OJ 02 L30U37 
I QQ 

HYDROGEN COMBUSTION BOUNDARY 

INJECTION P =Po; 

M0>1 

FQ. 4. Use of external burning with flap deflection to eliminate 

nozzle drag. 

initial fuel distribution and subsequent rate of mixing 

between the fuel and the air. This, of course, assumes 

that the hydmgen-air chemical reactions proceed at a rate 

approaching equilibrium (i.e., the combustion process can 

be thought of as "mixing limited"). Related to this is the 

basic question of flame stability at the conditions of 

interest. Is there an altitude or Mach number limit for 

this process? If indeed a flameholder is required to 

initiate combustion at the low-pressure, low-temperature, 

high-velocity conditions existing outside the engine cowl, 

its drag must be a small fraction of the total drag 

reduction. Finally, assuming that the concept is workable, 

the potential performance in terms of fuel used per net 

drag force reduction should be estimated. 

to this area falls roughly into two categories, "base 

burning," or combustion in the wake of projectiles to 

reduce drag, and "external burning," loosely defied by 

this author as the fueling and combustion of an airstream 

adjacent to an aerodynamic surface so as to actively 

control the pressure distribution on that surface. Most 

past work falls into the former category; Murthy, et 

contains a bibliography with over 350 references. The 

present application tends more toward the external 

buming concept, which has been used to reduce dra to 

provide control forces, and even to produce thnrst.g'll 

Most of this work had been done, however, with pyrophoric 

fuels and free-stream Mach numbers higher than the 

present range of interest. The high heating value of 

hydrogen combined with constant-pressure (and therefore 

constant velocity) combustion in a transonic stream results 

in an interesting deviation from most past studies, since 

the Mach number in the burning stream may be reduced to 

a subsonic value solely by increasing the sonic velocity 

without the usual turning or shocks. The resulting highly 

complex flowfield is characterized by an embedded elliptic 

region that has tittle reason to "close" downstream of the 

aircraft because of the constant-velocity nature of the 

process. [n the absence of velocity shear it seems that 

heat dissipation may be the only mechanism available to 

Strahle12 addressed this phenomenon analytically with a 

two-dimensional, small-perturbation analysis and 

concluded that a positive pressure coefficient could be 

maintained on a flat plate regardless of the transition to 

subsonic flow, but experimental verification was needed 

because the downstream boundary conditions could not be 

treated properly. 

A review of the Literature revealed that work related 

P return the combustion pmducts to a supersonic condition. 

h order ,to make an initial assessment of the 

transonic drag reduction potential of external burning, a 

control-volume analysis was done to obtain a first-order 

estimate of performance and fuel flow requirements. An 

experimental program was then begun to resolve issues 

including flame stability and the validity of various 

assumptions used in the control-volume analysis. The 

balance of this paper presents and discusses the results of 

both the analysis and the experimental program. 

cons tan t -Pressure, Con trol-Volume Analysis 

A detailed analysis of external burning in a transonic 

flow would be a formidable task, characterized by mixing 

and finite-rate combustion of hydrogen and air, 

three-dimensional mixed supersonic-subsonic flowfields, 

and the interaction of at least three streams. External 

burning analysis methods of varying de ees of 

sophistication do appear in the literature,lp20 but none are directly applicable to the current problem. In the 

present application of external burning, expanding 

combustion products must "fill" a void left by the vehicle 

base and the engine exhaust. To accomplish this, an 

amount of fuel must be burned with an appropriate amount 

of air - possibly at a specified rate. The size of the 

airstream that must be fueled and burned will be 

determined by the amount of expansion or the stream area ratio provided by the mass addition and combustion. It 

would seem that to completely relieve base drag, the free 

stream must be prevented from expanding into the base 

area. Therefore, the combustion products must occupy 

an area equal to the projected base area plus the 

cross-sectional area of the fueled airstream. 

In order to quantify the amounts of fuel and air 

involved, as well as to assess the fuel injection problem, 

the control volume pictured in Fig. 5 was studied. The 

scenario just described corresponds to a flow deflection 

angle of zero and will be referred to as the "desiqn" 

condition. The lower control surface is fonned by the 

boundary between the fueled streamtube and the free 

stream. Strictly speaking, the upper control surface 

coincides with the main engine exhaust shear layer and any 

aft-facing cowl surface. The dynamics of the engine 

exhaust stream are neglected, however, so that the upper 

control surface is thought of simply as a solid body. Air at 

free-stream conditions flows hugh the inflow boundary, 

which is perpendicular to the free stream. Fuel is injected 

normal to the free stream downstream of the inflow plane, 

and combustion products flow out of the outflow plane, 

which is also perpendicular to the free stream. It is 

assumed that both inflow and outflow properties are 

uniform and that velocity is parallel to the free stream. 

At the design condition the entire control volume is 

assumed to be at the local static pressure in the free 

stream; thus, disturbances due to the injection of fuel, etc. are neglected. 

7 HYDROGEN INJECTION 

I FUEL PRESSURE, PI : 1 FUEL TEMPERATURE, T, : \ ORIFICE DIAMETER, d': 

I ORIFICE SPACING, S 

STATION 2 

FREE-STREAM CONDITIONS 

(Mo. Po. To) 

FQ. 5. Control-volume nomenclature. 

f. 2. 

PRECEDING PAGE BLANK NOT FILMED 
The design condition of zero free-stream deflection 

angle results in the following geometric relation between 

the height of the streamtube of air that must be fueled 

and the area ratio obtained by combustion: 

yo 1 

Combining the continuity and momentum equations results 

in the following simple expression for the control-volume 

area ratio due to mass addition and combustion: 

- A2 = R2 T2 (1 + 

0 0 

A mieWor*&y consequence of the constant-pressure 

assumption is that the ratio of inflow to outflow velocity 

given by the momentum equation is equal to 1 plus the 

fuel-air mass ratio. The stoichiometric fuel-air mas 

ratio for hydrogen and air is 0.02916, which results in an outflow velocity very nearly equal to the inflow or 

free-stream velocity. Because the ratio of inflow to 

outflow sound speed is about 3 for stoichiometric 

combustion, the outflow (or downstream) boundary 

condition for the transonic external burning process is 

nearly always subsonic. This has important implications in 

the interpretation of experimental results to be discussed 

later. The fuel-air ratio in Q. (2) is considered an 

independent variable for the time being, and properties at 

the outflow boundary are assumed to be the equilibrium 

combustion products at this Fuel-air ratio. The fuel flow 

required at the design condition is easily determined fmn 

the fuel-air (or equivalence) ratio and the size of the 

streamtube to be fueled (Q. (1)). In terms of free-stream 

conditions the fuel flow per unit base area is given by 

Finally, the measure of goodness for external burning 

is taken as the net drag force reduction per unit fuel flow 

and is referred to herein as the specific impulse. The fuel 

flow comes directly from Eq. (3); the net drag force 

reduction requires further definition. At the design 

condition of zem drag the control-volume pressure is 

equal to the free-stream static pressure, and this pressure 

acts over an area equal to the aft-facing projected area of 

the body. By defining an "effective base pressure'' as 

the area-weighted average pressure acting on the 

aft-facing base surfaces without external burning, the 

specific impulse is defined as 

(4) 

In terms of the free-stream conditions, the equivalence 

ratio, and the control-volume area ratio given by m. (2), 

the specific impulse becomes 

burning concept depends on the equivalence ratio assumed, 

the flight condition, and the seventy of the drag problem. 

In Fig. 6, Eqs. (1) to (5) are applied to a 1000-lb/ft2 

abs dynamic pressure trajectory from Mach 0.8 to 2.6. 

Figure 6(a) shows the variation of required air streamtube 

height along the trajectory for equivalence ratios of 0.5, I. 

and 2. Streamtube heights of approximately 10 percent of 

the base height are required for equivalence ratios greater 

than 1 but increase sharply as equivalence ratio is 

decreased below stoichiometric. There is little benefit in 

using equivalence ratios greater than 1. since only mas 

addition contributes to a further increase in 

control-volume area ratio and this is partially offset by 

decreasing equilibrium temperature. Streamtube height 

varies with free-stream Mach number because 

free-stream static temperature decreases as altitude 

increases; the colder the inflow, the larger the 

temperature ratio from combustion. Note that the curves 

become flat as the vehicle climbs into the tropopause at 

about Mach 1.75. The effect of different trajectories on 

the curves of Fig. 6(a) is slight and due only to the inflow 

temperature effect. 

The required hydrogen flow per unit base area appears in Fig. 6@). For equivalence ratios of 1 or less 

0.1 to 0.2 Iblsec per square foot of base area is indicated. 

The required fuel flow increases more dramatically for 

rich mixtures as the streamtube height fmrn Fig. 6(a) 

becomes almost constant. For trajectories other than the 

.10 

.05 

EQUIVALENCE 

RATIO, 

0 

$$1111II 0 

(a) Streamtube height. 

.4 r 

0- 

@) Fuel flow. 

.5 1.0 1.5 2.0 2.5 3.0 

FREE-STREAM MACH NUMBER. M, 

(c) Speclfii impulse. 

Fig. 6. Results for free-stream-pressure 

control volume. Dynamic pressure at 

station 0. Qo. 1000 Iwh2 abs. 

Examination of as. (I) to (5) reveals that, under the 

assumptions discussed, the performance of the external 

4 
one shown the fuel flow is approximately proportional to 

the dynamic pressure, so that higher altitude trajectories 

would require less fuel. 

A specific impulse parameter is plotted in Fig. 6(c) 

and is practically independent of trajectory. Low 

equivalence ratios give high performance but require that 

fuel be distributed over large streamtube cross-sectional 

areas. It appears, however, that performance does not 

suffer greatly with a stoichiometric system. Note that the 

absolute value of specific impulse depends on the severity 

of the drag problem; the lower the base pressure Pt, the 

higher the specific impulse. It should be remembered that 

these high performance numbers are based on reduction, with the potential for this high performance 

being generated by the main propulsion system. 

Fuel Distribution Considerations 

Given the required fuel flow and the height of the 

streamtube of air to be fueled, the mechanics of 

distributing the fuel can be examined. Normal injection 

from a row of sonic orifices is discussed here although 

many different variations, including the use of spraybas, 

are possible. It is assumed that the height of the 

streamtube to be fueled, given by Eq. (I), coincides with 

the jet penetration of the choked orifices. Many 

correlations describe the penetration of a highl 

that is particularly useful for this application is that of 

Povinelli, et aL28 This correlation describes the contour 

representing a 5-percent volume concentration in the 

centerline plane of the injector and takes the following 

form for the case of normal, sonic injection and a thin 

approaching boundary layer: 

underexpanded jet into a supersonic crossflo~.~1-3 d One 

0.281 0.483 Y -$ = l.l$&) 6 + 0.5) (6) 

where the effective backpressure Pet, is taken to be 

two-thirds of the total pressure downstream of a normal 

shock at the free-stream Mach number for a supersonic 

free stream and two-thirds of the free-stream total 

pressure for a subsonic free stream. Equating the 

streamtube height with the height of the 5-percent 

hydrogen volume concentration at some distance 

downstream Wd' of the injectors may at first seem 

tenuous, but it at least provides the proper variation of 

fueled streamtube height with changing free stream and 

fuel conditions. The value of Wd' chosen wiU depend on details of the orifice and flameholder geometry and will 

,.Il....- *e met5od to be cdibrated. From cxperii~efiti! 

results on an expansion ramp, to be discussed in a 

subsequent section, an Wd' of 30 seem to work 

reasonably well and is used henceforth. 

The fuel flow rate for a choked injector can be 

written in terms of the fuel conditions and the orifice 

diameter: 

Equations (6) and (7) can now be combined to yield two 

parameters that are functions only of the desired 

equivalence ratio and the flight condition: 

.517 10.16($)(& + 0.5r*281k) (g) f 

483 

eb 

N\ 

.8g3[ i)eb 0 0.483 

(9) ($ t o.5)O*281 

As long as Eqs. (8) and (9) are satisfied, the streamtube 

height Yo from Fig. 6(a) will be injected with the 

appropriate amount of fuel from Fig. 6(b) consistent with 

the desired equivalence ratio. The actual distribution of 

fuel over the streamtube cms section will not be uniform, 

of couse, and in this sense the equivalence ratio is of an 

average or "global" nature. 

The performance shown in Fig. 6 was based on flight 

at constant dynamic pressure and streamtube equivalence 

ratio. With Eqs. (8) and (9) it is now possible to determine 

what schedule of fuel pressure and temperature is required 

to achieve a constant equivalence ratio over a range of 

Mach numbers, given an injector orifice diameter and 

spacing ratio (varying the orifice diameter and spacing 

ratio seems unlikely). Figure 7 shows such a schedule for a 

1000-lb/ft2 abs dynamic pressure flightpath and an 

equivalence ratio of 1. A significant variation in pressure 

and temperature is required for the particular injector 

geometry shown, which was chosen so as to keep the fuel 

temperature between about 500 and 1000 "R. Modulation 

of fuel pressure would be relatively easy as long as 

sufficient pressure were available in the fuel system, but 

the fuel temperature variation required probably will not 

match what is available in a power-balanced cycle. Some 

supplementary method of heating or cooling the hydrogen 

would have to be devised and would add weight and 

complexity to a system that must be carried to orbit after 

a short period of transonic operation. Obviously, other 

fuel schedules could be devised where the fuel conditions 

are relatively constant and the equivalence ratio varies, 

but as stated previously, equivalence ratios of 

approximately 1 are desirable. It is apparent that an 

"analysis" model of the external burning system is needed 

where fuel conditions and geometry are the independent 

loo0 r 

200 

0 

/ FUEL TEMPERATURE 

ECUEDL'LE 7 

'\ / 

\ 

\ r- FUEL PRESSURE '. ,' SCHEDULE 

-------- 

.5 1.0 1.5 2.0 2.5 3.0 

FREE-STREAM MACH NUMBER, Mo 

Fig. 7. Constant-equivalenca-ratio 

(Q - 1) fuel injection schedule. 

Dynamic ressure at station 0, O,, 

1000 lbmg abs; ratio of orifice spacing to orifice diameter, S/d', 7.56; 

ratio of ortice diameter to projected 

base height. d'Nb. 0.00694. 

5 
variables and the resulting equivalence ratio, 

control-volume pressure, and specific impulse are 

predicted. This involves some additional modeling. 

Off-Design Performance Prediction 

For the case where the inflow static pressure is 

different from the control-volume pressure and the 

free-stream deflection angle is nonzero (Eq. (1) no longer 

holds), combining the momentum and continuity equations 

yields the following equation for the control-volume area 

ratio: 

- A2 

*O 

+ 

(1 - $)I 

Note that this equation is identical to Eq. (2) except for 

the denominator, which involves the ratio of 

control-volume to free-stream pressure. Ultimately, this 

pressure ratio will be determined on the basis of the 

free-stream Mach number and a flow deflection angle. 

But first, the relationship between the control-volume 

area ratio given by Eq. (10) and the flow deflection angle 

must be determined. This is accomplished by assuming 

that the expansion is three dimerlsional, from the 

rectangle defined by the jet penetration height and the 

width of the expansion surface at station 0 to an 

appropriate rectangle at station 2, such that the area ratio 

defined by Eq. (10) is satisfied and the three sides of the 

control volume in contact with the free stream are at 

equal angles with the free stream. Since the control 

surfaces are all considered to be planar, this results in a 

streamtube area distribution that is quadratic in the axial 

direction and implies a quadratic temperature distribution 

as well. Whether or not this is physically realistic is 

beyond the scope of this simple analysis, but at least the 

three-dimensional "relieving" effect is accounted for 

approxha t ely. 

Now, the pressure throughout the control volume is 

considered to be equal to the pressure in the free stream 

after a turn through the deflection angle, as discussed 

previously. For supersonic flow small deflection angles 

are assumed and the pressure-versus-deflection-angle 

relation from linear theory is used: 

6 

For subsonic flow the problem is not quite as clearcut, but an approximation can be obtained by assuming 

incompressible flow over a wedge for which the velocity 

potential and stream function are known. Briefly, the 

pressure distribution corrected for compressibility is used 

to obtain the area-weighted average pressure acting on 

the deflected control surfaces. The final result is an expression for the control-volume pressure in terms of the 

free-stream Mach number and the deflection angle: 

The off-design or general problem of predicting the 

control-volume pressure given fuel conditions, orifice 

geometry, expansion surface geometry, and flight 

conditions can now be solved. First, the equivalence ratio 

is estimated by ratioing the fuel flow through one choked 

orifice to the amount of air at free-stream conditiori 

passing through a rectangle of width equal to the orifice 

spacing and height equal to the jet penetration. 

Figure 8(a) shows the variation of @valence ratio with 

Mach number on the 1000-lb/ft2 abs trajectory for a 

constant fuel pressure and temperature that correspond to 

the Mach 1.4 design point of Fig. 7. Also shown for 

comparison is the constant equivalence ratio obtained by 

varying the fuel conditions as per the Fig. 7 schedule. The 

equivalence ratio increases continuously as the jet 

penetration (and airflow) decrease at constant fuel flow. 

The station 2 properties after equilibrium combustion can 

now be determined, and Eq. (10) is used to determine the 

deflection angle. Note that Eq. (10) contains the 

control-volume pressure ratio so that an iterative solution 

using Eq. (11) or (12) is required. In Fig. 8(b) the 

control-volume pressure ratio so obtained is plotted, along 

with the design pressure ratio of 1. Note that additional 

geometric parameters describing the expansion surface 

adjacent to the upper control surface must now be 

specified. For Mach numbers less than 1.4, negative 

deflection angles result and control-volume pressures are 

below ambient. For Mach numbers greater than 1.4, the 

control-volume area ratio is sufficient to cause a positive 

deflection angle in spite of decreasing jet penetration and 

FUEL CONDITIONS 

SCHEDULED 

- 4 l.6r ____ CONSTANT 

.6 

(a) Equivalence ratio. 

LU 

8% .a- 

(b) Control-volume pressure. Ratio of base 

wdth at station 0 to projected base height, 

w/vb - 2.5; ratio of length of expansion 

surface to projected base height. UY, = 5. 

0 

.5 1.0 1.5 2.0 2.5 3.0 

FREE-STREAM MACH NUMBER. M, 

(c) Specific impulse; wnb I 2.5; Wb I 5. 

Fig. 8. Performance of example gWmetIY. 

Dynamic resure at station O.Qo., 

to orifice diameter, Sd'. 7.56; ratio of 

orifice diameter to projected base 

1000 lbnt g abs; ratio of orifice spacing 

height, d"b. 0.00694. 

6 
open symbols fall into a band that exhibits the same trend 

as the premixed stability Curve, except for the point at 

elevated free-jet temperature, which was stable to a much 

lower fuel pressure. The stability parameter has no explicit temperature dependence, however; and the higher 

temperature simply redts in a bigher velocity and an apparently more severe condition with no allowance for 

changes in reaction rates. 

Since the DeZubay parameter seems to be adequate 

for ambient temperatures, it was used to construct 

Fig. 14, where lines of constant DeZubay parameter are overlaid on an altitude-versus-Mach-number plot for a 

flameholder dimension of 1 in. In Fig. 13 a DeZubay 

parameter value of about 1000 could be construed as a 

practical limit for a slightly fuel-rich design. This limit is 

reached at Mach 1.5 for a 500-lb/ft2 abs dypmic pressure 

trajectory, and at Mach 2.4 for 2000 lWft abs. These 

limits, of course, increase with increases in flameholder 

dimension, but another practical limit of 2-psia static 

pressure is also shown beyond which stable combustion is 

unlikely regardless of the flameholder size. The preceding 

results indicate that ignition and flame stability must be 

carefully considered in the design of the external burning 

system but will not preclude its successful operation. 

FREEJET TOTAL 

TEMPERATURE, 

"R 

TCO. 

SOLID SYMBOLS 

- REPRESENT 

80 

60 

200 400 600 800 1000 

DUUBAY STABILITY PARAMETER, 

Fig. 13. Stability results. 

VO D4.74 , Wsec x atm x 11 

0 

VELOCITY TIMES 

PRESSURE AT DYNAMIC PRESSURE 

STATION 0. AT STATION 0. 

00, 

vo Po""'; 8 Wsec x psia lbfit2 abs 

500 

1000 

2000 

.5 1.0 1.5 2.0 2.5 3.0 

FREE-STREAM MACH NUMBER, Mo 

Fig. 14. Variation of DeZubay parameter with 

flight condition. 

7- ti' 

PRECEDING PAGE BLANK NOT FILMED 

In subscale combustion tests the relative importance 

of chemical kinetics should be evaluated at least 

qualitatively. Ideally, the combustion process would be 

mixing limited and similar results would be obtained at 

larger scale as long as the appropriate similarity 

parameters were matched. In order to gain some insight 

into the effect of finite reaction rates on the plume at the 

present conditions and scale, two calibrated infrared 

images of the spraybar plume are compared in Fig. 15. 

The Reynolds and Mach numbers are the same for each; 

however, the free-stream pressure and temperature are 

different. In Fig. 15(a) the free-stream pressure and 

temperature are both roughly twice those in Fig. 15@) at 

comparable equivalence ratio and jet penetration. In order 

to see the potential this creates for a change in the 

reaction rate, the reaction time correlation of 

PergamentJ2 was extrapolated to the present conditions; 

it predicts a reaction length for Fig. IS@) that is a factor 

of 3 greater than that of Fig. 15(a]. If the flow were 

premixed and completely reaction rate limited, this would 

result in a substantial change in the plume temperature 

contours. On the other hand, if the flow were completely 

mixing limited, the plumes should appear similar. Because 

a difference in the plume characteristics is apparent, it 

may be concluded that chemical reaction rates do have an 

effect on the plume characteristics at this Reynolds 

number. This undesirable result is not of great 

consequence for the flame stability results, since the 

correlating parameters contain appmpriate length scales 

and the spraybars were probably not too far from full scale 

anyway. The issue will be with subscale tests of the entire 

external burning process, where reaction rates and model 

scale may have a significant effect on the resultant 

pressure distributions. Fortunately, it would seem that 

these problems, while making data interpretation difficult, 

wiU lead to conservative results. If external burning is 

successful in small scale, confidence in full-scale success 

is increased greatly. 

(a) Free-stream static pressure, Po, 12 psia; tree-stream total temperature, Tt,o 

960 "R; equivalence ratio. 4. 0.48; jet penetration height, Y,. 0.72 in. 

(b) Free-stream static pressure, PO, 6 psia; free-stream total temperature, Ttp 

540 "A; equivalence ratio. 0, 0.51; jet penetration height. Yp. 1.0 in. 

Fig. 15. Infrared images of plume at Reynolds number of 4.8 million per 

foot. Black areas are <lo00 OR; white areas are 3400 OR. 

9 
ORJGINAL PAGE 

BLACK AND WIT€ PHOTOGW-.H 

\ 

i 

Fig. 16. Expansion ramp model with upper sidewalls mounted in free jet. 

Expansion Ramp Tests 

Apparatus and procedure. - The expansion ramp tests 

were intended to demonstrate drag reduction on a simple 

expansion ramp geometry while providing calibration and 

verification information for analysis methods. No changes 

were made to the facility, and the expansion ramp models 

were mounted in the free jet in much the same way as the 

spraybar models. The basic configuration, shown in 

Fig. 16, consisted of a 3- by 6-in. flat plate with a sharp 

leading edge, followed by an 11.2- by 6-in. expansion 

ramp. Two similar expansion ramps were tested, the only 

difference between the two being the spacing and 

diameter of the fuel injection orifices. The models were 

constructed of a single piece of 3/4-in.-thick stainless 

steel with a zirconium-oxide coating sprayed on the 

expansion surface. Upper sidewalls extending 2 in. above 

the leading edge are shown, but the models were tested 

primarily with lower sidewalls that were flush with the 

upper surface and extended 2 in. below the model at the 

trailing edge. The lower sidewalls were intended to keep 

high-pressure air generated by compression beneath the 

model from spilling around and affecting pressure 

distributions on the top surface. Fuel was injected normal 

to the free-jet axis through a row of choked orifices in a 

plane 112 in. upstream of the expansion comer. A 

flameholder was used to ensure combustion at the desired 

location and consisted of a 1/4-in.-wide by 1/8-in.-high 

strip of stainless steel spanning the entire 6-in. width of 

the model with a 114-in. gap in the center to allow for 

thermal expansion. The trailing edge of the flameholder 

was coincident with the expansion comer. Details of the 

fuel injection and flameholder arrangement are given in 

Fig. 17. The configuration pictured had twenty-six 

0.025-in-diameter orifices equally spaced across the 6-in. 

width to provide a range of equivalence ratios from 0.4 to 

1.2. The other had eight 0.044-in. injectors designed for 

somewhat lower equivalence ratios from 0.2 to 0.7. 

Figure 18 depicts the location of instrumentation 

with respect to the expansion ramp models. The only 

change from the spraybar tests is the addition of static 

pressure and temperature instrumentation on the upper 

surface of the models. A single, centerline row of 18 

static pressure taps and an off-centerline row of 5 

thermocouples were used. 

The expansion ramp tests were all run at the ambient 

free-jet total temperature of 540 "R. Because the gas 

generator was not available during these tests, only 

ambient-temperature hydrogen fuel was used. 

Fig. 17. Expansion ramp fuel injection and flameholding region. 

INFRARED CAMERA 

FIELD OF VIEW 7, 

1 

I 

I STATIC 

I PRESSURE 

I 

FREEJET ! NOZZLEEXIT 

I 

! 

Results. 

(26 injectors) 

flameholder. 

- The higher-equivalence-ratio model 

was tested first without any type of 

Model ignition could only be accomplished at a subsonic free-jet Mach number of about 0.6. As the 

free-jet supply pressure was increased and the design 

Mach number of 1.26 was reached, color video and infrared 

images indicated that the leading edge of the flame was 

anchored at a point about halfway down the ramp surface 

probably at the boundary layer separation point, having 

little effect on the pressure distribution. With the 

flameholder installed the flame remained anchored near 

the flameholder trailing edge, but supersonic model 

ignition using the translating spark ignitor was still not 

possible. The inability to ignite the supersonic stream 

should not be taken as a general result, however, since the 

arc location is a critical parameter that was not varied. 

The low-equivalence-ratio model was ignited in subsonic 

flow as well but would not sustain combustion in 

supersonic flow even with the flameholder instailed. For 

this reason, only results for the high-equivalence-ratio, 

26-orifice model are presented. 

Model centerline static pressure distributions with 

the lower sidewalls installed are shown in Fig. 19 for a 

range of fuel pressures. Fuel-off and fuel-on (not burning) 

pressure distributions are also shown for comparison. The 

pressure gradient on the forward portion of the model was 

caused by a detached bow shock on the leading edge that 

could not be made thin enough for the Mach 1.26 free 

stream. The effect of the flameholder is also apparent as a large overpressure. The no-burning pr.e-ssure 

distributions exhibit the expected overexpansion at the 

3-in. station, which corresponded to the flameholder 

trailing edge and model "knea" Boundary layer separation 

and three-dimensional relief then caused a rapid 

recompression to free-stream static pressure. Combustion 

affected the pressure all the way upstream to the leading 

edge and eliminated much of the large overexpansion at 

- 

10 
the model knee, creating a region of relatively constant 

pressure over the ramp surface. The combustion pressure 

remained below the free-stream value, however, and 

showed little variation with fuel pressure in contrast to 

the control-volume prediction, even though both the jet 

penetration and the estimated equivalence ratio were increasing markedly. 

of 30 in Q. (6). This resulted in a good correlation 

between the measured maximum plume temperature and 

the estimated equivalence ratio for all altitudes and fuel 

pressures. The maximum plume temperatures were, 

however, somewhat lower than the theoretical equilibrium 

temperature at the correlated equivalence ratio. In order 

to provide a better (or at least more conservative) model 

of the station 2 conditions, a combustion efficiency was 

added to the control-volume pmcedure so that for any 

altitude and fuel pressure the calculated temperature at 

station 2 approximately matched the measured maximum 

plume temperature. This is the basis for the 

"control-volume predictions" in Figs. 19 to 23. 

The estimated jet penetration was based on an Wd' 

FUEL ESTIMATED JET ESTIMATED 

PRESSURE PENETRATION, EQUIVALENCE 

in. 

mno. 0 PI. YP * 

wa 

A OFF 0 0 

0 180 .25 .59 

v 350 .35 .a 

250 30 .69 

0 250(NOBURN) 30 .69 

.4 r 

2 .2 

$6 -0 0 

EE 0 

Eg -.4 

3- 0 cno 

wo -.2 CONTROL. 1 

Zii VOLUME 1 iu 

to 

z 

w 0 -.6 

PREDICTION 

2 .2 

$6 -0 0 

EE 0 

Eg -.4 

3- 0 cno 

wo -.2 CONTROL. 1 

Zii VOLUME 1 iu 

to 

z 

PREDICTION 

8 -.6 lr -. 8 

0 2 46 8 10 12 14 

DISTANCE FROM LEADING EDGE, in. 

Fig. 19. Effect of increasing fuel pressure on ramp surface 

static pressure distributions. Free-stream static pressure, 

Po, 8 psia. 

TEST ESTIMATED JET ESTIMATED 

CEU PENETRATION. EQUIVALENCE 

PRESSURE. Y, . RATIO. r 

psia in. 0 

0 12 0.29 CONTROL- 08 35 .83 VOLUME 

04 .a 1.17 PREDICTON7 

.4 

-.6 

I 

0 

0 

0 

I I I I I I I 

2 

TEST ESTIMATED JET ESTIMATED 

CELL PENETRATION. EQUIVALENCE 

PRESSURE. Y" . RATIO. 

ps'a in. 0 

0 12 0.29 0.68 

.35 .83 

.48 1.17 

us 

04 CONTROL0 0 0 VOLUME 

t t t PREDICnoN 

:- I PROJECTED 

RAMP 

HEIGHT 

I I 

-6 I U 

0 1000 2000 3000 4000 5000 

CORRECTED TOTAL TEMPERATURE, Ti 

Fig. 21. Effect of increasing altitude on plume total 

temperature profiles. Fuel pressure, Pf, 350 psia. 

(a) Free-stream static pressure, Po , 12 psia (5500 ft) 

(b) Free-stream static pressure, Po, 8 psia (16 000 n). 

(c) Free-stream static pressure, Po , 4 psia (32 000 ft). 

Fig. 22. Infrared images of plume at various altitudes. Fuel pressure, P,. 

350 psia. 

0 2 46 8 10 12 14 

DISTANCE FROM LEADING EDGE, in. 

Fig. 20. Effect of increasing altitude on ramp surface static 

pressure distributions. Fuel pressure, P,, 350 psia. 

11 
SIDEWALLS 

0 UPPER 

0 LOWER CONTROLVOLUME 

PREDICTION 7, 

'e 

I I I I I I I I 

0 2 46 8 10 12 14 

DISTANCE FROM LEADING EDGE, in. 

Fig. 23. Effect of sidewalls on ramp pressure distribution. 

Fuel pressure, P, - 250 psia; estimated jet penetration, 

Y,, 0.30; estimated equivalence ratio, 0. 0.70. 

The effect of increasing the altitude at a constant 

fuel pressure, shown in Fig. 20, was much the same as the 

effect of increasing the fuel pressure; both equivalence 

ratio and jet penetration increased without affecting the 

combustion pressure. Very little variation in combustion 

pressure was noted over the entire range of altitudes and 

fuel pressures despite a factor-of-2 variation in both the 

jet penetration and the estimated equivalence ratio. 

Plume total temperature profiles corresponding to 

the three conditions of the previous figure are shown in 

Fig. 21. The plume temperature and size both increased 

with altitude, as jet penetration and equivalence ratio 

increased. Obviously, the control-volume predictions show 

good agreement with the maximum plume temperatures, 

since the procedure was calibrated by using these data. 

The fact that measured plume temperatures did reach the 

theoretical maximum for hydrogen and air is encouraging. 

Calibrated infrared images given in Fig. 22 for the 

conditions of the previous figure show large changes in 

both the temperature and extent of the plume, with the 

plume apparently "filling" the base region. 

The effect of upper sidewalls on the model 

centerline static pressure distribution is shown in Fig. 23. 

These sidewalls (pictured in Fig. 16) extended 2 in. above 

the model at the leading edge and had the expected effect 

of limiting three-dimensional relief. Since pressure was 

below the free-stream static without sidewalls, the effect 

of the sidewalls was to slightly lower the ramp pressure. 

A two-dimensional expansion assumption was used in the 

control-volume analysis to model this effect. 

Lower than predicted ramp pressures in all cases 

could be due to a number of factors including the inherent 

assumption in the control-volume analysis that the control 

volume acts as a solid body to the free stream. The 

effects of nonunjform inflow and outflow and flameholder 

drag were neglected and could lead to discrepancies. 

Another source of uncertainty lay with the experimental 

apparatus itself as discussed in the next section. 

Factors lnfluencing Test Results 

The results obtained to date with the expansion ramp 

model are somewhat curious in nature given the vigorous 

combustion demonstrated and the lack of agreement with 

the control-volume analysis. Figure 24 depicts 

phenomena, currently being investigated in follow-on 

tests, that may possibly have influenced the static 

pressure distribution on the expansion ramp. One of the 

unique features of the external burning flowfield is the 

constant-pressure, constant-velocity plume, which 

resulted in a subsonic condition downstream of the model. 

WAVE REFLECTIONS FROM 

RELEASE - 

EFFECTS 

BENEATH 

AM 

Fig. 24. Factors influencing test results. 

Further, only a small velocity gradient existed between 

the subsonic plume and the supersonic free stream so that 

the subsonic condition could persist for large distances 

downstream of the model. Since upstream communication 

was possible within this subsonic core, reflected 

disturbances, the facility exhaust collector, expansion 

around the model base from beneath, or anything else 

causing a presjure perturbation downstream of the model 

could influence the ramp static pressure. 

The slight overexpansion still present at the model 

knee with external burning may be due to a delay in the 

onset of heat release. Mixing and reaction kinetics 

probably both played a role here. Moving the fuel 

injection and flameholding farther upstream and increasing 

the free-stream temperature will help to alleviate this 

pmblern. 

Finally, the necessity of using a flameholder could 

lead to unexpected results, since flameholder drag was 

neglected in the control-volume analysis. Testmg 

different size flameholders may give some insight into this 

effect; however, the use of a flameholder gives rise to a 

scaling issue. The 1/8-in.-high flameholder used extended 

a significant distance into the fueled stream. If this were 

scaled geometrically to give similar aerodynamic 

characteristics to a large test article, a prohibitive drag 

would result. Although from a flameholding standpoint it 

is not necessary to scale the flameholder geometrically, 

the mechanics of flame spreading fmrn the pilot region to 

the outer reaches of the fueled stream will only be similar 

if the ratio of flameholder height to jet penetration is held 

constant. 

Summary and Conclusions 

External burning, used in conjunction with a variable 

cowl flap to prevent exhaust flow overexpansion, is a 

promising transonic drag reduction concept. Results of a 

simple control-volume analysis indicate that transonic 

drag can be eliminated with hydrogen flow rates of 0.1 to 

0.2 Ib/sec per square foot of base area at 1000-lb/ft2 abs 

dynamic pressure, with fuel flows being roughly 

proportional to the dynamic pressure. The specific 

impulse performance of the external burning scheme in 

terms of drag force reduction was 1000 to 3000 sec and 

was proportional to the severity of the drag force without 

burning. Normal sonic orifices can be used to inject the 

fuel the required distance into the free stream, which is 

approximately 10 percent of the base height. 

Experimental results indicate that hydrogen and air 

will burn at altitude in transonic flow. A flame stability 

correlation parameter published for a premixed 

hydrogen-air stream worked adequately if a suitable 

definition of equivalence ratio was used for the 

non-premixed stream. Flame stability Limits may be 

encountered at high altitude, at high Mach number, or both. The effect of finite rate chemistry and the use of 

flameholders make scaling of small-scale test results 

difficult. The external burning process was used to 

12 
i 

increase pressures on a small expansion ramp at Mach 1.26 

to altitudes of 32 000 ft, but measured performance was not as high as predicted by the control-volume analysis. 

Ramp pressure showed little variation with fuel pressure 

and altitude despite large changes in the temperature and 

size of the plume; plume temperatures equal to the 

theoretical maximum for hydrogen and air were recorded 

just downstream of the expansion ramp. A number of reasons for these discrepancies, including anomalous 

facility effects, were discussed. The nearly 

constant-velocity nature of the external burning process 

presents a unique challenge to the experimentalist in 

providing a disturbance-free test medium. Some form of 

atmospheric or flight test may be required to completely 

resolve the magnitude of the external burning benefit. 

Also, finite chemical reaction times at these conditions 

and the use of an unscaled flameholder may necessitate 

testing at large scale, depending on the degree of 

confidence desired in the full-scale result. 

References 

1. 

2. 

3. 

4. 

5. 

6. 

7. 

8. 

9. 

10. 

11. 

Keyworth, G.A., II, "National Aeronautical R&D 

Goals," Aeronautical Policy Review Board, Executive 

Office of the President, Office of Science and 

Technology Policy, Washington, DC, 20500, March 

1985. 

Murthy, S.N.B., ed., et al., Aerodynamics of Base 

Combustion, MIT Press, Cambridge, 1976, pp. 143-210. 

Dorsch, R.G., Serafini, J.S. and Fletcher, E.A., 

"Exploratory Investigation of Aerodynamic Effects of 

External Combustion of Aluminum Borohydride in 

Airstream Adjacent to Flat Plate in Mach2.46 

Tunnel," NACA RM E57E16, 1957. 

Dorsch, R.G., Serafini, J.S. and Fletcher, E.A., 

"Experimental Investigation of Aerodynamic Effects 

of External Combustion in Airstream Below 

Two-Dimensional Supersonic Wing at Mach 2.5 and 

3.0," NASA memo 1-ll-59E, 1959. 

Serafini, J.S., Dorsch, R.G. and Fletcher, E.A., 

"Exploratory Investigation of Static- and BasePressure Increases Resulting From Combustion of 

Aluminum Borohydride Adjacent to Body of 

Revolution in Supersonic Wind Tunnel,'' NACA RM 

E57E1.5, 1957. 

Schadow, K.C., "Experimental Investigation of 

Combined External BurningIBase Burning," m, 

=E, ASME 17th Joint Propulsion Conference, 

Hardford, CT, 1980. 

Hubbartt, J.E. and Strahle, W.C., "ExternaVBase 

Burning for Base Drag Reduction at Mach 3," AIAA 

Journal, Vol. 19, Nov. 1981, pp. 1502-1504. 

Caswell, G.J., "Base Drag Reduction by External 

Burning in the Inviscid Stream," M.S. Thesis, Naval 

Postgraduate School, Monterey, CA, Sept. 1973. 

Reding, J.P. and Jecmen, D.M., "Effects of External 

Burning on Spike-Induced Separated Flow," AIM 

Paper 82-1360, Aug. 1982. 

Billig, F.S., "External Burning in Supersonic Streams," 

Johns Hopidns University Applied Physics Laboratory, 

TG912, May 1967. 

Dugger, G.L. and Monchick. L.. "External Burning 

Ramjets, Preliminary Feasibility 

Study," Johns Hopkins University Applied Physics 

Laboratory, TG892, Mar. 1967. 

12. Strahle, W.C., "Some Outer Limits in the Theory of 

Transonic Flow With External Burning," Combustion 

Science and Technology, Vol. 1, Apr. 1970, 

13. Pinkel. I.I., Serafini, J.S. and Gregg, J.L., "Pressure 

Distribution and Aerodynamic Coefficients Associated 

With Heat Addition to Supersonic Air Stream 

Adjacent to Two-Dimensional Supersonic Wing," 

NACA RM E51K26, 1952. 

14. Schetz, J.A., Billig, F.S. and Favin, S., "Approximate 

Analysis of Base Drag Reduction by Base and/or 

External Burning for Axsymmetric, Supersonic 

Bodies," AIAA Paper 80-1258, June 1980. 

15. Mehta, G.K. and Strahle, W.C., "Anaiysls of Axiaiiy 

Symmetric External Burning Propulsion for Bluff-Base 

Bodies," AIM Paper 78-26, Jan. 1978. 

pp. 357-363. 

16. Broadbent, E.G., "Flows With Heat Addition and 

Associated Pressure Fields," Fluid Dyn amic 

Transactions, Vol. 6, Pt. 1, 1972, pp. 80-113. 

17. Smithey, W. and Fuhs, A.E., "Base Pressure 

Calculations With External Burning," in 

Aerodynainics of Base Combustion, MIT Press, 

Cambridge, 1976, pp. 407-424. 

18. Maririo, A., "Theoretical Performance With External 

Burning on Surface," General Applied Science Labs, 

Inc., Westbury, New York, GASL TR-506, Feb. 1965. 

19. Oswatitsch, K., "Thrust and Drag With Heat Addition 

to a Supersonic Flow," R.A.E. Library Translation 

No. 1161, January 1967 (Summary report presented to 

the 3rd Euromech Symposium, Aachen, Feb. 1966). 

20. Harvey, D.W. and Davis, J.C., "External Burning 

21. Adamson, T.C., Jr., and Nicholls, J.A., "On the 

Propulsion Analysis," ALAA Paper 81-1477, July 1981. 

Structure of Jets From Highly Underexpanded Nozzles 

Into Still Air," Jounial of the AedSpace Sciences, 

Jan. 1959, pp. 16-24. 

22. Zukoslci, E.E. and Spaid, F.W., "Secondary Injection of 

Gases Into a Supersonic Flow," AIAA Journal, Vol. 2, 

Oct. 1964, pp. 1689-1696. 

23. Crist, S., Sherman, P.M. and Glass, D.R., "Study of 

the Highly Underexpanded %nic Jet," AIAA Journal, 

Vol. 4, Jan. 1966, pp. 68-71. 

24. Schetz, J.A. and Billig, F.S., "Penetration of Gaseous 

Jets Injected Into a Supersonic Stream," 

J. Spacecraft, Vol. 3, Nov. 1966, pp. 1658-1665. 

25. Schetz, J.A., Hawk, P.F. and Lehman, H., 

"Structure of Highly Underexpanded Transverse Jets 

in a Supersonic Stream," AIAA Journal, Vol. 5, 

May 1967, pp. 882-884. 

Temperature and Molecular Weight Effects in Jet 

Interaction," AIAA Journal, Vol. 8, Feb. 1970, 

26. Chrans, L.J. and Collins, D.J., "Stagnation 

pp. 287-293. 

13 
27. Hersch, M., PovineUi, L.A. and Povinelli, F.P., 

"Optical Study of Sonic and Supersonic Jet 

Penetration From a Flat Plate lnto a Mach 2 

Airstream," NASA TN D-5717, Mar. 1970. 

28. Povinelli, F.P. and PovinelLi, L.A., "Correlation of 

Secondary Sonic and Supersonic Gaseous Jet 

Penetration Into Supersonic Clossflows," NASA TN 

D-6370, June 1971. 

29. Bag, F.S., Orth, R.C. and Lasky, M., "A Unified 

Analysisof Gaseous Jet Penetration," AIAA Journal, 

Vol. 9, June 1971, pp. 1048-1058. 

30. Orth, R.C.. Schetz, J.A. and Billig, F.S., "The 

interaction and Penetration of Caseous Jets in 

Supersonic Flow," NASA CR-1386, 1969. 

31. DeZubay, E.A., "A Study of Flame Stability Based on Reaction Rate Theory." ASME Paper 54-SA-27, 

May 1954. 

32. Huber, P.W., Schexnayder, C J- aad McChton, C.R., 

"Criterion for Self-Ignition of Supersonic 

Hydrogen-Air Mixtures," NASA TP-1457, 1979. 

14 
National Aeronautics and Space Administration 

Unclassified 

Report Documentation Page 

Unclassified 16 A03 

2. Government Accession No. NASA TM-103107 1. Report No. 

AIAA-90-1935 

4 Title and Subtitle 

On the Use of External Burning to Reduce 

Aerospace Vehicle Transonic Drag 

7. Author(s) 

Charles J. Trefny 

9. Performing Organization Name and Address 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135-3191 

12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, D.C. 20546-0001 

3. Recipient's Catalog No. 

5. Report Date 

6. Performing Organization Code 

8. Performing Organization Report No. 

E-543 1 

10. Work Unit No. 

763-01 -2 1 

11. Contract or Grant No. 

13. Type of Report and Period Covered 

Technical Memorandum 

14. Sponsoring Agency Code 

15. Supplementary Notes 

Prepared for the 26th Joint Propulsion Conference cosponsored by the AIAA, ASME, SAE, and ASEE, Orlando, 

Florida, July 16-18, 1990. 

16. Abstract 

The external combustion of hydrogen to reduce the transonic drag of aerospace vehicles is currently being investigated. A preliminary analysis based on a constant-pressure control volume is discussed. Results indicate that the 

specific impulse of the external burning process rivals that of a turbojet and depends on the severity of the initial 

base drag as well as on the flight Mach number and the equivalence ratio. A test program was conducted to 

investigate hydrogen-air flame stability at the conditions of interest and to demonstrate drag reduction on a simple 

expansion ramp. Initial test results are presented and compared with the control-volume analysis. The expansion 

ramp surface pressure coefficient showed little variation with fuel pressure and altitude-in disagreement with the 

analysis. Flame stability results were encouraging and indicate that stable combustion is possible over an adequate 

range of conditions. Facility interference and chemical kinetics phenomena that make interpretation of subscale 

ground test data difficult are discussed. 

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

External burning 

Transonic drag reduction 

19. Security Classif. (of this report) 1 20. Security Classif. ( 

18. Distribution Statement 

Unclassified - Unlimited 

Subject Category 07 

this page) I 21. NO. of pages I 22. Price'