This document analyzes the performance of the Russian RD-120 rocket engine. It examines how performance metrics like specific impulse, thrust, and characteristic velocity vary with different fuel-to-oxidizer mixture ratios. The analysis finds that while a ratio of 3 is optimal, producing the highest specific impulse, the actual engine design uses a ratio of 2.6. This ratio provides performance close to optimal while lowering combustion temperatures for improved durability. The analysis confirms several of the engine's reported performance specifications using its actual mixture ratio.

1. AN EXAMINATION OF THE RD-120 ROCKET
ENGINE AND DETAILED MIXTURE RATIO TRADE
STUDY ANALYSIS
JOHANN SCHRELL
THE GEORGIA INSTITUTE OF TECHNOLOGY
AE6450 FALL 2009
DR. MITCHELL WALKER

2. ii
TABLE OF CONTENTS
TABLE OF CONTENTS ________________________________________________________________ ii
TABLE OF FIGURES AND TABLES ______________________________________________________ ii
INTRODUCTION ____________________________________________________________________ 1
ANALYSIS _________________________________________________________________________ 3
Specific Impulse (Isp) ___________________________________________________________________ 4
Flame Temperature ____________________________________________________________________ 4
Thrust Coefficient and Thrust ____________________________________________________________ 4
Characteristic Velocity C* _______________________________________________________________ 5
Delta-V Δv ____________________________________________________________________________ 5
RESULTS AND ASSUMPTIONS________________________________________________________ 13
CONCLUSIONS ____________________________________________________________________ 14
REFERENCES ______________________________________________________________________ 14
TABLE OF FIGURES AND TABLES
Figure 1. Isp Results From CEA With Variation in O/F at Vacuum and Sea Level ______________ 6
Figure 2. Flame Temperature With Variation in O/F _____________________________________ 7
Figure 3. Thrust Coefficient From CEA With Variation in O/F ______________________________ 8
Figure 4. Thrust With Variation in O/F at Vacuum and Sea Level___________________________ 9
Figure 5. C* With Variation in O/F at Vacuum and Sea Level _____________________________ 10
Figure 6. Delta-V With Variation in O/F at Vacuum and Sea Level ________________________ 11
Figure 7. Delta-V/c with Variation in Mass Ratio _______________________________________ 12
Table I. RD-120 Engine Dimensions ____________________________________________________ 2
Table II. RD-120 Performance Parameters ______________________________________________ 3
Table III. Results of Analysis Using Design Mixture Ratio _________________________________ 13

3. 1
INTRODUCTION
The RD-120 engine is a Russian liquid fuel rocket engine developed and used 1976-1982. The
developing company is NPO Energomash, previously Gasdynamic Laboratory - Experimental
Design Bureau. The engine is designed to operate at altitude as an upper stage and incorporates a
large exit area ratio nozzle to accommodate this. The engine was used on the Zenit launch vehicle
second stage. It has a non-vectorable nozzle. The propellant delivery is provided via one main
and two boost turbopumps. Ignition and turbopump operation is accomplished by using a single
preburner. The engine was extremely reliable during its service. 177 RD-120s were built and
tested 560 times for a total common operating time of 139186 sec.
The RD-120 produces over 833,565.25 N of thrust operating at about 162.7962 bar of chamber
pressure. The engine’s specific impulse (Isp) in vacuum is listed as 350 sec. Though the Isp of the
Kerosene/LOX combination is lower of similar engines running on LH/LOX, its Isp density is greater.
The RD-120 operates at a nominal propellant flow rate of 242.9 kg/sec and is able to throttle down
to 85% nominal thrust. This allows for less fuel volume to be carried on the vehicle. Dimension of
the engine are detailed below in Table I.

4. 2
Table I. RD-120 Engine Dimensions
Engine Dimensions
Characteristic Dimension (mm, kg)
Length 3872
Diameter 1954
Dry Mass 1125
Wet Mass 1285
Chamber Diameter 320
Characteristic Length, L* 1274
Contraction Ratio 1.74 (unitless)
Throat Diameter 183.5
Exit Diameter 1895
Exit Area Ratio 106.7 (unitless)
Chamber Length 2992

5. 3
ANALYSIS
A review of the performance of this engine is desired. To accommodate this, a trade study of
mixture ratios was run using the same operating pressure and real configuration’s dimensions to
find the optimum performance. This was then compared to the predicted and actual performance
of the real configuration of the RD-120. The real contraction ratio, throat diameter, exit area
ratio, and dry mass were used in analysis. The real mass flow rate was also used.
One mixture ratio that is desired to be analyzed is the stoichiometric mixture ratio. The equation
for this is shown below.
C12 H 24 (l ) 18O2 (l ) 12CO2 ( g ) 12 H 2O( g )
Equation 1. RP-1 and LOX Reaction Equation
This equation yields a mixture ratio of 3.429 using the equation below.
O
N ox MWox
F
N fuel MW fuel
Equation 2. Molecular Mixture Ratio Equation
Along with the stoichiometric mixture ratio, the real configuration mixture ratio will be analyzed.
Other mixture ratios from 0.5 to 5.0 will be analyzed as well for a good characterization of the
rockets performance. The system’s level of control of the mixture ratio is a property of the feed
system and is rated at ±10%. Since the real configuration’s mixture ratio is O/F=2.6, this means
that the mixture ratio is really anywhere between 2.34≤O/F≤2.86. A list of the real configuration’s
performance parameters is shown in Table II.
Table II. RD-120 Performance Parameters
Engine Performance Parameters
Parameter Value
Mixture Ratio Control, % ±10
Throttling, % 85
Thrust (vacuum), N 833,565.25
Burn Time, sec 315
Specific Impulse (vacuum), sec 350
Propellant Mass Flow, kg/sec 242.9
Mixture Ratio, O/F 2.6
Combustion Flame Temperature, K 3670
Chamber Pressure, bar 162.7962
Nozzle Exit Pressure, bar 0.127491
Thrust Coefficient (vacuum) 1.95

6. 4
Specific Impulse (Isp)
The vacuum and sea level Isp was calculated using NASA CEA with the methods previously
described. The vacuum Isp is our primary data set of interest as the RD-120 was used as an upper
stage in vacuum conditions. Values for the sea level performance are not available for the real
engine to compare to. In Figure 1 it is seen that the optimum mixture ratio is 3 and produces an
Isp of 376 sec. Also shown on this figure are the values for the stoichiometric and real
configuration mixture ratios. It is seen that in this case the stoichiometric mixture ratio does not
produce the best performance. The real configuration mixture ratio also does not produce the
largest optimum specific impulse. However, it will be seen in further investigation why this
mixture ratio was chosen by the designers.
Flame Temperature
The flame temperature was also calculated using the NASA CEA code and the results are plotted
versus mixture ratio in Figure 2. The mixture ratio, 3, that produced the highest specific impulse
also has the highest flame temperature of 3835 K. This becomes a heat transfer management
problem. In this case there are two options, increase the mixture ratio (fuel lean) or decrease the
mixture ratio (fuel rich). In either case there is a sacrifice that must be considered in trading heat
management with lower specific impulse. In the real case, 0.8% in Isp is lost but the temperature
is also dropped 1.1%. It is also seen that the stoichiometric mixture ratio produces a higher flame
temperature and lower Isp than the real configuration mixture ratio. This is more evidence as to
why the designers decided to go fuel rich instead of fuel lean.
Thrust Coefficient and Thrust
The thrust coefficient is an important parameter in nozzle design and represents the amplification
of thrust due to the supersonic expansion of the nozzle compared to if the nozzle exit area were
equal to the throat area. In Figure 3 the thrust coefficient has been plotted with the mixture ratio.
It seen that this curve can have a rather strange shape. This dip in the value prior to rising again is
mainly due to the gas properties at those particular mixture ratios. A sharp change in pressure
ratio and specific heat ratio occurs that forces the thrust coefficient lower. This however does not
affect the thrust as seen in Figure 4. Also first notice that the thrust coefficient is not at its highest
for the real mixture ratio. This means that some improvement in the nozzle sizing could be made.
Once again, the best performing mixture ratio is 3, showing the highest thrust. However, it is also
seen again that the real mixture ratio performed better than the stoichiometric mixture ratio,
producing over 886 kN of thrust. The thrust was calculated using the ideal thrust equation shown
below.
F m v2 P2 A2
Equation 3. Thrust Equation

7. 5
Characteristic Velocity C*
The characteristic velocity was calculated using the effective exhaust velocity and thrust
coefficient obtained from NASA CEA according to the equation below.
c
C*
CF
Equation 4. Characteristic Velocity Equation
The optimum mixture ratio seen in Figure 5 producing the highest C* is 2.5 and the real mixture
ratio is very close with a C*=1876 m/sec. The characteristic velocity is a function of the propellant
combustion and acts as a measure of how well the combustion chamber is designed. This means
that since the C* is near its highest for the real mixture ratio, the chamber of the RD-120 was
properly designed for optimum combustion at the design mixture ratio. Because some energy is
lost through the nozzle exhibiting incomplete combustion during real tests, it is beneficial to look
at C* efficiency. This is a ratio of the actual C* and the theoretical C* and is a measure of the
completion of energy release in the propellants and creation of high temperature and high
pressure gases. The C* efficiency for the real mixture ratio is 94% which is in the typical range of
92-99.5%.
Delta-V Δv
The delta-V is the amount change in velocity the rocket engine would produce if it was to fly with
just its own mass plus fuel mass. It was calculated using the equation below and plotted in Figure
6.
mf mtb
v c ln
mf
Equation 5. Delta-V Equation for Rocket Only
In this equation mf represents the dry mass of the engine and the added term represents the mass
of propellant consumed. The plot shows confirmation that the real mixture ratio provides one of
the highest values of over 15 km/sec. On a vehicle this number would be lower due to the
structural mass and payload. The calculated inverse mass ratio for the RD-120 is 69.012 and it can
be seen in Figure 7 how much delta-V can be expected when taking the effective exhaust velocity
into consideration.

8. 6
400
Stoichiometric
380 Real O/F = 2.6 O/F = 3.429
Isp = 373 sec Isp = 371 sec
360
340
Specific Impulse Isp (sec)
320
Vacuum
300
Sea Level
280
260
240
220
200
0 1 2 3 4 5 6
Mixture Ratio O/F
Figure 1. Isp Results From CEA With Variation in O/F at Vacuum and Sea Level

9. 7
4500
Stoichiometric
Real O/F = 2.6 O/F = 3.429
4000 T = 3816 K
T = 3793 K
3500
3000
Temperature (K)
2500
2000
1500
1000
500
0
0 1 2 3 4 5 6
Mixture Ratio (O/F)
Figure 2. Flame Temperature With Variation in O/F

10. 8
2.1
Stoichiometric
O/F = 3.429
CF = 1.99
2
Real O/F = 2.6
CF = 1.95
1.9
Thrust Coefficient Cf
1.8
1.7
1.6
1.5
0 1 2 3 4 5 6
Mixture Ratio O/F
Figure 3. Thrust Coefficient From CEA With Variation in O/F

11. 9
1000000
Stoichiometric
Real O/F = 2.6 O/F = 3.429
900000
F = 886 kN F = 881 kN
800000
700000
600000
Thrust (N)
Vacuum
500000
Sea Level
400000
300000
200000
100000
0
0 1 2 3 4 5 6
Mixture Ratio O/F
Figure 4. Thrust With Variation in O/F at Vacuum and Sea Level

12. 10
2000
1800
1600
Real O/F = 2.6
C* = 1876 m/sec Stoichiometric
1400
Characteristic Velocity C* (m/sec)
O/F = 3.429
C* = 1822 m/sec
1200
Vacuum
1000
Sea Level
800
600
400
200
0
0 1 2 3 4 5 6
Mixture Ratio O/F
Figure 5. C* With Variation in O/F at Vacuum and Sea Level

13. 11
18000
16000
14000
Real O/F = 2.6
12000 Stoichiometric
Δv = 15412 m/sec
O/F = 3.429
Δv = 15387 m/sec
Delta-V ( m/sec)
10000
Vacuum
Sea Level
8000
6000
4000
2000
0
0 1 2 3 4 5 6
Mixture Ratio O/F
Figure 6. Delta-V With Variation in O/F at Vacuum and Sea Level

14. 12
5
4.5
4
3.5 Mass Ratio for Max Mass Flow
Rate and 315 sec Burn Time,
1/MR = 69.012
3 Delta-V/c = 4.234
Delta-V/c
2.5
2
1.5
1
0.5
0
0 10 20 30 40 50 60 70 80 90 100
1/MR
Figure 7. Delta-V/c with Variation in Mass Ratio

15. 13
RESULTS AND ASSUMPTIONS
The analysis performed here makes some assumptions and they are listed below.
Combustion is adiabatic and isentropic
There are no heat losses to the engine materials
The ambient pressure is 0 bar
Propellant mass flow rate is constant and at 100% capacity
The given thrust and thrust coefficient are experimental values
Chamber pressure is constant and at 100% capacity
Rocket is operating in gravity free, drag free space
It is desired to compare the results of this analysis with the performance values provided in
Table II. After performing the mixture ratio trade study there is agreement that the design
mixture ratio O/F=2.6 is an appropriate value to use. Additional analysis was performed at
this mixture ratio using the mixture ratio control. This gives results for performance
parameters in the domain of 2.34≤O/F≤2.86. These results are summarized in Table III.
Table III. Results of Analysis Using Design Mixture Ratio
Results of Analysis
Parameter Real Calculated
O/F 2.6 2.34 2.6 2.86
Isp (vacuum), sec 350 366 373 376
Flame Temp, K 3670 3700 3793 3815
Thrust Coeff 1.95 1.91 1.95 1.98
C*, m/sec 1760 1870 1876 1860
Thrust, kN 834 870 886 892
Exit Pressure, bar .127 .084 .103 .120
In this table it is seen that the values are close but do not match those of the real engine. In
most cases the values are larger than those of the real engine. This is due to the assumptions
that are stated previously as well as a few other possibilities. One main possibility for losses is
the efficiency of which the propellants are injected into the combustions chamber and mixed
prior to combusting. There is no such thing as perfect mixing of propellant in rockets. Also,
due to boundary layer effect, the effective aerodynamic throat of the real rocket in use will be
smaller than that of the physical throat.

16. 14
CONCLUSIONS
No real safety issues have been found with the RD-120 rocket engine. Liquid engines of this
size have some inherent risk. The engine requires a complex turbopump system to deliver
the propellants at high flow rates. This involves a lot of plumbing that is subject to leaks.
Also, this rocket uses staged combustion so it is likely there are hot preburner exhaust gases
flowing in very close proximity to the propellants in the feed system. In 1983 a fuel system
leak led to the destruction of Soyuz T-10-1 on the launch pad. The RD-120 was not used on
this vehicle however.
From the analysis completed it appears there is only a little room for improvement via
optimizing the mixture ratio. Some extra performance could be obtained with little increase
in system risk it appears. The areas for improvement on a rocket engine of this size include
materials studies. A trade could be performed to see if only the throat region and chamber
need to be regeneratively cooled and the exit cone made of an ablative composite such as
carbon/carbon or carbon phenolic. This would save weight and reduce system complexity.
The combustion chamber could have small dimensional changes with test to determine a
configuration with more complete combustion, however the current design probably already
resulted in much testing for this as well as combustion stability. Because the RD-120 already
uses staged combustion, an extremely efficient system, it is already a very well designed
engine with similar characteristics as the LH 2/LOX fueled Space Shuttle Main Engines.
REFERENCES
1. “Combustion Chemistry”. http://www.innovatia.com/Design_Center/rktprop2.htm.
2. “RD-120”. http://www.npoenergomash.ru/eng/engines/rd120/.
3. Sutton, George P., Biblarz, Oscar. Rocket Propulsion Elements. John Wiley & Sons.
2001. 7th Ed.