NASA Technical Memorandum 107285

AIAA-96-2582

A Performance Comparison of

Two Small Rocket Nozzles

Lynn A. Arrington

NYMA, Inc.

Brook Park, Ohio

Brian D. Reed

Lewis Research Center

Cleveland, Ohio

Angel Rivera, Jr.

University of Texas

San Antonio, Texas

Prepared for the

32nd Joint Propulsion Conference

cosponsored by AIAA, ASME, SAE, and ASEE

Lake Buena Vista, Florida, July 1-3, 1996

National Aeronautics and

Space Administration

A Performance Comparison of Two Small Rocket Nozzles

Lynn A. Arrington

NYMA, Inc.

Brookpark, Ohio 44142

Brian D. Reed

National Aeronautics and Space Adminstration

Lewis Research Center

Cleveland, Ohio 44135

and

Angel Rivera, Jr.

Undergraduate, Senior

University of Texas

San Antonio, Texas

Abstract

An experimental study was conducted on two small rockets (110N thrust class) to directly compare a standard conical nozzle with

a bell nozzle optimized for maximum thrust using the Rao method. In large rockets, with throat Reynolds numbers of greater than

1x 105, bell nozzles outperform conical nozzles. In rockets with throat Reynolds numbers below 1× 105, however, test results have

been ambiguous. An experimental program was conducted to test two small nozzles at two different fuel film cooling percentages

and three different chamber pressures. Test results showed that for the throat Reynolds number range from 2x 104 to 4x 104, the bell

nozzle outperformed the conical nozzle. Thrust coefficients for the bell nozzle were approximately 4 to 12 percent higher than those

obtained with the conical nozzle. As expected, testing showed that lowering the fuel film cooling increased performance for both

nozzle types.

Introduction

Future space exploration will require increasing payload

fraction and decreasing system masses. Therefore, maximizing

rocket engine performance is an important design goal. Maximizing the thrust for a rocket engine can be achieved by

optimizing the nozzle contour. Optimizing the performance of

finite length nozzles is typically accomplished using an inviscid core flow and a boundary layer displacement. G.V.R. Rao

developed a method which optimizes a rocket nozzle contour

for a given length or expansion ratio such that maximum thrust

is achieved. 1 Rao's method was based on the assumption of

inviscid isentropic flow.

This method has been used in many studies of different classes

of rocket engines with a variety of results. Rocket engines are

generally classified by their throat Reynolds number based on

uniform flow and properties at the throat. For engines with

Reynolds numbers larger than lxl05, the Rao optimization is

quite effective, because the ratio of boundary layer flow to the

total flow is small. In a study by Farley, 2 three large Rao

optimized bell nozzles were compared to a 15 degree conical

nozzle. The thrust produced by the optimized nozzles was greater

than that obtained with the conical nozzle. In fact, bell contour

nozzles have been used routinely for many years in large liquid

rocket engines. 3,4,5 Conical nozzles are typically used only when

fabrication and design costs outweighed performance.

In very small thrusters with Reynolds numbers less than

I xl 04, viscous boundary layer effects are large in comparison

to the total flow and a bell contour is not nearly as effective, as

discussed in apair of experimental studies. 6,7 Inthese studies,

a variety of small nozzles were tested to determine the effect of

contour, propellant, expansion ratio, and Reynolds number on

performance. Normally, factors in performance loss are divergence, boundary layer effects, and heat transfer effects to the

walls. By using ambient temperature gas, the heat transfer

effect was eliminated in both studies. The nozzles were conical,

trumpet, and Rao optimized bell contours. One paper concluded that the difference in contours had no effect on either

divergence or viscous losses. 6 The other concluded the bell

contour had the lowest performance of the nozzles tested. 7 The

reasons given in both were that the viscous effects of boundary

layer were so large that an inviscid isentropic assumptions were

not valid.

For rockets that are sized between the very small and large

ones the usefullnes of the Rao optimization code is less clear.

Investigations on a gaseous hydrogen/oxygen engine were
performed by M.A. Appel with different nozzle contours in the

134 to 445 N thrust class. 8 A rhenium 80 percent bell nozzle, a

regeneratively cooled nozzle, and a stainless steel 18 degree half

angle conical nozzle were each tested using the same injector in

this effort. The stainless steel nozzle was built to verify injector

and thrust stand operation prior to testing the other two nozzles.

The regeneratively cooled nozzle had different dimensions for a

different aspect of the study, but the conical and bell nozzles had

the same length, exit area and expansion ratio. Unexpectedly, the

test results showed that the conical nozzle provided a higher

specific impulse than the bell nozzle. The author postulated that

the conical nozzle operated with a higher kinetic efficiency and

suggested this as an area for further investigation. Kushida 9

examined a 22 N monopropellant hydrazine thruster with a 100:1

area ratio conical nozzle, a300:1 Rao optimized bell, andthe same

bell nozzle cut off at expansion ratios of 200:1 and 100:1. Results

showed that the 300:1 bell nozzle outperformed the conical

chamber but the conical was the better performer than the lower

area ratio bell nozzles. A direct comparison was difficult, however, because several dimensions were changed in addition to the

nozzle contours. Both the length and expansion ratio were greater

for the optimized nozzle, which generally improves performance

in a rocket engine. Based on the ambiguous results reported to

date, the validity of the Rao nozzle optimization method is

uncertain for thrusters with Reynolds numbers between lxl04

and lxl05.

In the study reported herein, testing was performed to directly evaluate the applicability of Rao optimization methods to

rockets with Reynolds numbers below 1x 105. The performance

of two 110 N chamber/nozzles with throat Reynolds numbers

between 2x 104 and 4><104 was measured to examine the effect

of nozzle contour on performance in this intermediate Reynolds

number range. Both nozzles were designed with the same

chamber profile, throat diameter, and expansion ratio. Because

the chamber and nozzle are one continuous structure, each

piece of hardware is simply identified by the nozzle shape, ie.

bell or conical, from here on. The conical nozzle had a 19 degree

expansion half angle. This rocket was originally tested in

1993.10 The bell nozzle was designed via the Rao code. As part

of the effort, a one dimensional equilibium code, ODE, was

used to generate theoretical performance predictions to calculate the characteristic velocity efficiencies presented later in

this study. 11

Test Hardware

Both nozzles tested were fabricated via the same techniques

using oxygen free high conductivity (OFHC) copper with water

cooling passages in the walls. The inner housings were milled

first to cut water cooling passages into the outer surface. The

chamber profiles were identical from the injector face to the

throat. The contours of the nozzles are shown in figures 1(a) and

(b), and further details of the bell nozzle contour are given in

table I. The divergence angle of the conical nozzle was about

19 degrees and the divergence angle of the bell nozzle was

about 9 degrees at the exit plane. The design specified throat

area was 1.267 cm 2. Post fabrication measurements of the

chamber, throat, and exit gave dimensions that were very

similar as shown in table II. The resulting dimensional differences based on area were 0.10 percent at the chamber, 2.2 percent at the throat, and 0.14 percent at the nozzle exit.

The outer housing of each nozzle was machined in two pieces

separated along the thruster axis. The sections fit around the

inner housing as a shell to insure the interior contour was one

continuous profile from chamber inlet to nozzle exit. The outer

housing halves were attached to each other with bolts around

the inner housing. All seams were filled with high temperature

solder. Both chambers were instrumented with thermocouples

and pressure taps in several quadrants along the contour.

The same injector, 12 shown schematically in figure 2, was

used in all testing. The injector was designed to produce a core

flow of combustion products with an annulus of fuel film

cooling (FFC). In the core, oxygen was injected through a

platelet stack into the combustion chamber just upstream of a

spark plug used for ignition. The hydrogen flow was divided by

means of a fuel splitting washer, with pan injected downstream

of the spark plug into the core and the rest used to fuel film cool

the chamber wall. The washer was specifically designed to give

a predetermined fuel film cooling split from the total fuel flow.

The film cooling was directed along the chamber wails by an

annular sleeve. By changing the washer, the amount of film

cooling could be varied between 55 to 85 percent. A 61 percent

FFC percentage was selected as the baseline for this study since

that was nominal design operating point for this injector. Tests

at 75 percent FFC were also run to investigate the sensitivity to

FFC changes.

Test Facility

All tests were conducted in a small rocket test facility, 13in

which an equivalent vacuum of 36.6 km altitude is achieved

using air ejectors. Gaseous hydrogen and gaseous oxygen were

used for propellants. A schematic is shown in figure 3. The tank

is 1.8 m long between the flanges and 0.9 m in diameter. Inside

the tank, the rockets are oriented horizontally to fire through a

diffuser into an exhaust quenching chamber. The rocket

exhaust is pulled through the pair of ejectors and vented to

atmosphere by mufflers. The ejectors are supplied with

900 kPa air at 4.5 kg/sec to provide the pumping for the altitude

tank.

Thrust is measured by mounting the rocket on a thrust stand.

The stand and rocket are supported with flexure plates. Both

float free on the flexures and apply a load to a strain gage load

cell through a ball joint. To ensure accurate thrust calibration,

the ball joint must make contact with the load cell at all times.

The thruster is positioned to apply a small load on the load cell
evenwhenthethrusterisnotfiring, to ensure contact with the

load cell is maintained. The small load is accounted for in thrust

calibration and is subtracted from the rocket thrust. The propellant and pressure transducers lines are mounted perpendicular to the rocket axis with stainless steel tubing to provide

repeatable thrust measurements. The thrust stand is calibrated

by applying loads through a piston assembly to compare the

thrust measurement load cell against another previously calibrated reference load cell. Pretest, posttest, and applied load

tests are taken at altitude conditions with pressurized propellant

lines. All thrust measurements are corrected for background

pressure to determine vacuum thrust.

Rao Optimization Code

Rao discussed the need for nozzle contour optimization. 14 In

order to maximize thrust, a thruster is designed to produce

uniform parallel exit flow, with the exit plane pressure equal to

the ambient pressure. In a conical nozzle, flow divergence

losses are significant, particularly as cone angles increase. Rao

suggested that contouring the nozzle wall would turn the flow

closer to the thruster axis and thus reduce flow divergence. He

developed an optimization method for designing nozzles which

assumes inviscid isentropic flow. 15The method uses the calculus of variations for an ideal gas with constant gamma expansion to calculate the optimum nozzle contour. Supersonic flow

properties across the nozzle throat are input to start the solution.

Transonic flow in the throat region is calculated as a function

of the upstream radius of curvature at the throat and the specific

heat ratio for axisymmetric flow. The method iterates to an

optimum nozzle contour defined to meet for one of three initial

conditions: a given length, a given expansion ratio, or a fixed

envelope (exit radius and length). The Rao code was used in this

study to obtained an optimized bell nozzle contour with the

dimensions shown in table I using the fixed envelope condition.

Test Procedures

The same tests were conducted on each nozzle using gaseous

hydrogen and oxygen propellants. Tests were conducted at

three different chamber pressures, 500, 370, 255 kPa, and two

FFC percentages, 61 and 75 percent, over a mixture ratio range

of 4 to 8. The design condition for the rocket hardware was

500 kPa with 61 percent FFC. The higher FFC percentage and

the lower chamber pressures were chosen to allow testing

through a throat Reynolds number range of 2x 104 to 4><104. The

throat Reynolds number (Re) range was calculated using the

equation

4rh

Re = m

_tgDt

where m is the mass flow rate of the propellants, !1 is the

absolute viscosity of the combustion products, and D t is the

diameter of the nozzle throat. The viscosity is determined from

theoretical predictions generated by the Gordon-McBride

program.11 The two percent difference in throat areas sometimes resulted in slightly lower chamber pressure in the conical

nozzle. A typical test began by pumping the test tank to altitude

and pressurizing the propellant lines with fuel. A thrust stand

calibration was then performed. After the calibration, a constant load was applied to the thrust stand, which provided data

for determining the uncertainties. Once the applied load was

removed and inlet pressure conditions at the injector were

adjusted, the rocket engine was fired. By modifying the propellant inlet pressures, mixture ratios from 4 to 8 were tested. Once

tests were completed for the day, another constant load test and

another load cell calibration were conducted.

Measurement Uncertainties

Measurement uncertainties were calculated using the

JANNAF recommended procedure.16 The uncertainties were

calculated from the precision (random) and bias (constant)

errors associated with the measurement instruments and data

acquisition system. Precision errors propagated through the

performance parameters of specific impulse, characteristic

velocity, and thrust coefficient are shown in table III. Initial

testing with the bell nozzle for the 500 and 370 kPa chamber

pressure cases at 61 percent FFC had a very small pretest load

(less than 0.5 N) on the thrust stand prior to ignition. Since the

thrust stand could float free of the load cell at this load, a larger

load was applied prior to ignition to ensure pretest contact

between the thruster and the load cell. The low pretest load

affected the repeatability of the thrust stand zero. In an effort to

ensure that the thrust measurement was not affected by the

small pretest load, the pretest load was increased and a few

selected points in that test series were repeated. A field check

at that time showed performance results consistent with the

original tests at the smaller pretest load. However, in subsequent data reductions, the tests with the small pretest load had

significantly larger uncertainties. The range of uncertainties for

both cases are shown in table HI, the cases with the smaller

pretest load being indicated by parenthesis. The remainder of

the test data uncertainties were also presented in the same table.

In general, the greatest uncertainties were at the lower chamber

pressures.

Results and Discussion

Injector/Chamber Performance

To show that the injector and chamber behaved identically

for the two nozzles, the characteristic velocity efficiencies are
plottedinfigures 4 and 5 as a function of mixture ratio for tests

run at 61 and 75 percent FFC, respectively. The efficiencies

were determined by dividing the experimental characteristic

velocity by the theoretical equilibrium characteristic velocity

predictions. 11 As expected, the plots show no measurable

differences in this parameter between the bell and conical

nozzles as the injector and chamber profiles are the same. The

slight difference in throat area resulted in slightly higher

chamber pressures (approximately I percent), in the bell nozzle

for the same inlet conditions. Clearly the difference did not

have a measurable impact on the performance of the injector

and chamber, since the characteristic velocities measured with

both nozzles fell within the same uncertainty band. The overall

efficiency is slightly higher for the 61 percent FFC cases. This

was an expected result because more hydrogen is reacted in the

core combustion flow at this condition. Similiar trends are

shown in the plots of characteristic velocity versus mixture

ratio (figs. 6 and 7).

Nozzle Performance

Vacuum specific impulse versus mixture ratio and thrust

coefficient versus mixture ratio are typically used to illustrate

nozzle performance. The vacuum specific impulse versus mixture ratio for the 61 and 75 percent FFC cases are plotted in

figures 8 and 9, respectively, for each of the three chamber

pressures tested. The data were fit to a second order polynomial

curve and the measurement uncertainties added. At 61 percent

FFC (fig. 8), the plots of nozzle data show that the bell nozzle

specific impulse are at least 5 percent higher at 500 kPa,

6 percent higher at 370 kPa, and 12 percent higher at 255 kPa

than that of the conical for the 61 percent FFC cases. In

75 percent FFC cases (fig. 9) the bell nozzle specific impulse is

at least 5 percent higher at 500 kPa, 10 percent higher at

370 kPa, and 14 percent higher at 255 kPa. These results are

statistically significant and show clearly that the bell nozzle out

performed the conical nozzle. The higher performance obtained

with lower FFC percentage was expected as discussed above.

Thrust coefficients are plotted for the 61 and 75 percent FFC

cases in figures 10 and 11, respectively, again for each chamber

pressure case. As seen in the specific impulse plots, the bell

nozzle had the higher performance for both FFC percentages.

The thrust coefficients at 61 percent FFC are generally 6 percent higher at 500 kPa, 7 percent higher at 370 kPa, and

11 percent higher at 255 kPa for the bell nozzle than for the

conical. The thrust coefficients at 75 percent FFC are generally

4 percent higher at 500 kPa, 8 percent higher at 370 kPa, and

11 percent higher at 255 kPa for the bell nozzle than for the

conical. Again clearly, the bell nozzle outperforms the conical

nozzle for this Reynolds number range of between 20,000 to

40,000.

Conclusions

Testing was conducted to directly compare the performance

of a bell nozzle designed by the Rao optimization method and

a conical nozzle of the same size in the 2x104 to 4x104 throat

Reynolds number range. Previous studies had indicated that an

optimized bell nozzle may not outperform a conical nozzle at

these Reynolds numbers. To eliminate the effects of injector

performance, the same injector and chamber profile were used.

Tests were conducted over a mixture ratio range of 4 to 8 for

three different chamber pressures and two fuel film cooling

levels. Characteristic velocity measurements indicated that the

injector and chamber gave the same performance with both

nozzles. Plots of vacuum specific impulse and thrust coefficients indicated a higher performance by the bell nozzle. Also,

higher performance at the lower FFC percentages for both

nozzles. For the throat Reynolds number range of this design,

the test results confirm that the Rao optimized bell nozzle yields

higher performance than a simple conical nozzle design.

References

IRam, G.V.R., "Exhaust Nozzle Contour for Optimum Thrust," Jet Propulsion,

Vol. 28, No. 6, 1958.

2Farley, J.M. and Campbell, C.E., "Performance of Several Method of Characteristics Exhaust Nozzles," NASA TN D293, October 1960.

3"Liquid Rocket Engine Nozzles," NASA SP-8120, July 1976.

4Dunn, C., Kaith, 1., O'Shaughnessy, G., Rankel, R., and Rutherford, P.,

"Design of a High Performance Bell Rocket Nozzle for a High Pressure

and Thrust NTR," AIAA 91-3627, September 1991.

5Smith, T.A., "Boundary Layer Development as a Function of Chamber

Pressure in the NASA Lewis 1030:1 Area Ratio Rocket Nozzle,"

AIAA-88-3301, 1988.

6Grisnik, S.P., Smith, T.A., and Saltz, L.E., "Experimental Study of Low

Reynolds Number Nozzles," AIAA Paper 87_992, May 1987.

7Whalen, M.V., "Low Reynolds Number Nozzle Flow Study," NASA

TM-100130, July 1987.

8Appel, M.A., Schoenman, L., and Berkman, D.K., "Oxygen/Hydrogen Thrusters for the Space Station Auxiliary Propulsion Systems," JANNAF

Propulsion Meeting, Vol. 1, 1984, pp. 309-350.

9Kushida, R.O., Hermel J., Apfel, S., and Zydowicz, M., "Performance of High

Area Ratio Nozzles for a Small Rocket Thrusters," AIAA Paper 86-1573,

June 1986.

t°Momen, S.H., Myers, R.M., Benko, S.E., Arrington, L.A., and Reed B.D., "A

Laboratory Model of a Hydrogen/Oxygen Engine for Combustion and

Nozzle Studies," NASA TM- 106281, AIAA Paper 93-1925, June 1993.

llGordon, S. and McBride, B.J. "Computer Program for Calculation of

Complex Chemical Equilibrium Compositions, Rocket Performance,

Incident and Reflected Shocks, and Chapman-Jouguet Detonations,"

NASA Lewis Research Center, NASA SP 273, 1971.

12Robinson, P.J., "Space Station Auxiliary Thrust chamber Technology,"

Final Report 2210-90-FR, NASA CR-185296, July 1990.

13Arrington, L.A. and Schneider, S., "Low Thrust Rocket Test Facility,"

NASA TM-103206, AIAA Paper 90-2503, July 1990.

14 Rao, G.V.R., "Recent Developments in Rocket Nozzle Configurations,"

ARS Journal, November 1961, pp. 1488-1494.
15Nickerson, G.R., Dang, A.L.,Dunn, SS.,"TheRaoMethodOptimumNozzle

ContourProgram," GeorgeC.Marshall SpaceFlightCenter,

NAS8-36863, February 15,1988.

16"Handbook for Estimating the Uncertainty in Measurements Made with

Liquid Propellant Rocket Engine Systems," Chemical Propulsion Informarion Agency Document Number 180, April 1969.

TABLE I.--WALL CONTOUR POINTS FOR

RAO NOZZLE BEGINNING AT THROAT

TO NOZZLE EXIT

Axial Radial Axial Radial

distance, distance, distance, distance,

cm cm cm cm

0.000

0.157

0.175

0.203

0.231

0.262

0.295

0.330

0.366

0.404

0.444

0.488

0.533

0.579

0.632

0.686

0.744

0.805

0.871

0.940

1.016

1.095

1.179

0.638

0.681

0.691

0.709

0.726

0.744

0.765

0.785

0.808

0.833

0.856

0.884

0.912

0,942

0.973

1.006

1.041

1.080

1.120

1.161

1.204

1.252

1.298

1.267

1.359

1.450

1.544

1.623

1.727

1.910

2.217

2.621

2.891

3.152

3.495

3.843

4.280

4.740

5.314

5.928

6.302

6.891

7.534

8.037

8.434

8.989

1.349

1.400

1.450

1.501

1.542

1.595

1.687

1.836

2.017

2.129

2.233

2.362

2.487

21631

2.774

21936

3.096

3.185

3.315

3.444

3.538

3.607

3.693

TABLE II._ROSS SECTIONAL AREAS OF THE

BELL AND CONICAL NOZZLES

Location Cross sectional area,

cm:, cm

Bell nozzle,

radius

Conical nozzle,

radius

Chamber 5.045 (1.268) 5.050 (1.267)

Throat 1.248 (0.637) 1.275 (0.630)

Nozzle exit 42.65 (3.687) 42.70 (3.685)

TABLE III.--PERFORMANCE PARAMETER UNCERTAINTY PERCENTAGES

FOR THE BELL AND CONICAL NOZZLES AS A FUNCTION

OF CHAMBER PRESSURE

Uncertainty, :1:percent Bell nozzle Conical nozzle

Specific 500 kPa

impulse

370 kPa

255 kPa

Characteristic 500 kPa

velocity

370 kPa

255 kPa

Thrust 500kPa

coefficient

370 kPa

255 kPa

61% FFC 75% FFC 61% FFC 75% FFC

2.6 2.6 1.1 1.1

(4.0)

3.3 2.9 1.4 1.5

(6.8)

5.7 4.6 2.0 1.9

1.2 1.1 1.2 1.1

1.6 1.4 1.5 1.5

2.4 2.1 2.0 2.0

2.6 2.6 0.4 0.5

(4.1)

3.3 2.9 0.7 0.7

(7.0)

6.0 4.5 1.1 0.9
Hydrogen inlet,

 oo,.

center " _/

line _

z

(_

Water inlet

F2 places,

180 dog apart

Water outlet, 2 places,

180 deg apart

F Throat center line

18.9 °

ref

Hydrogen inlet,

2 places, 180 deg apart

(b)

F Water inlet 4 places,

_r_' 90 deg apart

Nozzle ! 1

center ....

line _

Water outlet, 4 places,

90 deg apart----_, i

L

Throat

I center line _ I J//

t/ J_

Figure 1 ._(a) Schematic of conical nozzle. (b) Schematic of bell contoured nozzle.
Sparkplug Fuelsplittingwater

/ Sleeve

Typical

combustion

chamber

H2 injection

H2 film

cooling

injection

Fuel manifold

02 injection

Figure 2.--Details for Aerojet SN 03 injector.

GO2

GH2

GN2

H20

H20

H20

Water cooled

pipe

Not to scale

Spray

cart

Figure 3.--Schematic of test rig.

Air

Ejector #2

Ejector #1

Exterior wall

-7
1.0

0.9

0

¢-

O

0.8

0.5 4

o Bell, 500 kPa

u Bell, 370 kPa

A Bell, 255 kPa

• Conical, 500 kPa

• Conical, 370 kPa

Conical, 255 kPa

O

o

>=

L)

._ 0.7

¢J

a=

J_

o

0.6

Mixture ratio

Figure 4.---Characteristic velocity efficiency versus mixture ratio for

61 percent FFC for both bell and conical nozzles.

1.0

0.9

L)

f-

(.1

,0.8

>

L)

"_ 0.7

"C

r0.6

0.5

o o ¢b

:=_. ,.=.. _, ,,% ._, ._. ,_,

AA t

o Bell, 500 kPa

o Bell, 370 kPa

A Bell, 255 kPa

• Conical, 480 kPa

• Conical, 365 kPa

• Conical, 250 kPa

4 5 6 7 8

Mixture ratio

Figure 5._haracteristic velocity efficiency versus mixture ratio for

75 percent FFC for both bell and conical nozzles.
24OO

2300

2200

2100

03

2000

"_ 1900

O

> 1800

"c 1700

1600

1500

1400

1300

1200

_A° _Q

o Bell, 500 kPa

o Bell, 370 kPa

= Bell, 255 kPa

• Conical, 500 kPa

• Conical, 370 kPa

• Conical, 255 kPa

ooo

4 5 6 7

Mixture ratio

Figure 6._haracteristic velocity versus mixture ratio for 61 percent

FFC for both bell and conical nozzles.

2400

23OO

2200

o 2100

2000

_5 1900

0

> 1800

om

•c 1700

1600

0 1500

1400

1300

1200

o

=n

IBBO

o Bell, 500 kPa

o Bell, 370 kPa

" Bell, 255 kPa

• Conical, 480 kPa

• Conical, 365 kPa

• Conical, 250 kPa

o

'_ A _lbE]•

"" A AA%• I_

4 5 6 7 8

Mixture ratio

Figure 7._Characteristic velocity versus mixture ratio for 75 percent

FFC for both bell and conical nozzles.

9
480

46O

440

420

i°°

38O

36O

34O

320

30O

280

260

240

A Bell, 255 kPa

o Bell, 370 kPa

4 5 -6 _ 8 9

Mixture ratio

Figure 8._Specific impulse versus mixture ratio for 61 percent FFC for

both bell and conical nozzles, with uncertainty bars.

48O

460

440

420

400

W

_o

360

.B

_- 340

_320

3O0

280

260

240

o Bell, 370 kPa

A Bell, 255 kPa

o Bell, 500 kPa

• Conical, 480 kPa

• Conical, 365 kPa

• Conical, 250 kPa

, , ., ,

4 5 6 7 8 9

Mixture ratio

Figure 9._Specific impulse versus mixture ratio for 75 percent FFC for

both bell and conical nozzles, with uncertainty bars.

]0
2.4

"5

0

¢J

2

r"

p2.3

2.2

2.1

2.0

1.9

1.8

1.7

1.6

1.5

1.4

1.3

1.2

¢-_"= _ T ..... _ = k= _ w= -- _r

A Bell, 255 kPa

o Bell, 370 kPa

o Bell, 500 kPa

• Conical, 500 kPa

• Conical, 370 kPa

• Conical, 255 kPa

4 5 6 "/ 8 9

Mixture ratio

Figure lO.--Thrust coefficient versus mixture ratio for 61 percent FFC for

both bell and conical nozzles, with uncertainty bars.

2.4

2.3

2.2

2.1

2.0

"E

"_ 1.9

,_

8 1.8

_ 1.7

r1.6

1.5

1.4

1.3

II |, l Ij[[ ,.

a Bell, 255 kPa

a Bell, 370 kPa

o Bell, 500 kPa

• Conical, 480 kPa

• Conical, 365 kPa

• Conical, 250 kPa

1.2 - , - • - •

4 5 6 7 8

Mixture ratio

Figure 11 .--Thrust coefficient versus mixture ratio for 75 percent FFC for

both bell and conical nozzles, with uncertainty bars.

11
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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

August 1996 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

A Performance Comparison of Two Small Rocket Nozzles

8. AUTHOR(S)

Lynn A. Arrington, Brian D. Reed, and Angel Rivera, Jr.

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

WU-242-70--02

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-10360

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM-107285

AIAA-96-2582

11. SUPPLEMENTARY NOTES

Prepared for the 32nd Joint Propulsion Conference cosponsored by AIAA, ASME, SAE, and ASEE, Lake Buena Vista,

Florida, July 1-3, 1996. Lynn A. Arrington, NYMA, Inc., 2001 Aerospace Parkway, Brook Park, Ohio 44142 (work

performed under NAS3-27186); Brian D. Reed, NASA Lewis Research Center; Angel Rivera, Jr., Undergraduate Senior,

University of Texas, San Antonio, Texas. Responsible person, Brian D. Reed, organization code 5330, (216) 977-7489.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 18

This publication is available from the NASA Center for AeroSpace Information, (301) 621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

An experimental study was conducted on two small rockets (110N thrust class) to directly compare a standard conical

nozzle with a bell nozzle optimized for maximum thrust using the Rao method. In large rockets, with throat Reynolds

numbers of greater than 1x 105, bell nozzles outperform conical nozzles. In rockets with throat Reynolds numbers below

lxl05, however, test results have been ambiguous. An experimental program was conducted to test two small nozzles at

two different fuel film cooling percentages and three different chamber pressures. Test results showed that for the throat

Reynolds number range from 2x104 to 4x104, the bell nozzle outperformed the conical nozzle. Thrust coefficients for the

bell nozzle were approximately 4-12% higher than those obtained with the conical nozzle. As expected, testing showed

that lowering the fuel film cooling increased performance for both nozzle types.

14. SUBJECT TERMS

Nozzle; Ran; Conical; Low thrust; Reynolds number

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

13

16. PRICE CODE

A03

20. LIMITATION OF ABSTRACT

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298-102

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