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AN EXAMINATION OF THE RD-120 ROCKETENGINE AND DETAILED MIXTURE RATIO TRADE STUDY ANALYSIS JOHANN SCHRELL THE GEORGIA INSTITUTE OF TECHNOLOGY AE6450 FALL 2009 DR. MITCHELL WALKER
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iiTABLE OF CONTENTSTABLE OF CONTENTS ________________________________________________________________ iiTABLE OF FIGURES AND TABLES ______________________________________________________ iiINTRODUCTION ____________________________________________________________________ 1ANALYSIS _________________________________________________________________________ 3 Specific Impulse (Isp) ___________________________________________________________________ 4 Flame Temperature ____________________________________________________________________ 4 Thrust Coefficient and Thrust ____________________________________________________________ 4 Characteristic Velocity C* _______________________________________________________________ 5 Delta-V Δv ____________________________________________________________________________ 5RESULTS AND ASSUMPTIONS________________________________________________________ 13CONCLUSIONS ____________________________________________________________________ 14REFERENCES ______________________________________________________________________ 14TABLE OF FIGURES AND TABLESFigure 1. Isp Results From CEA With Variation in O/F at Vacuum and Sea Level ______________ 6Figure 2. Flame Temperature With Variation in O/F _____________________________________ 7Figure 3. Thrust Coefficient From CEA With Variation in O/F ______________________________ 8Figure 4. Thrust With Variation in O/F at Vacuum and Sea Level___________________________ 9Figure 5. C* With Variation in O/F at Vacuum and Sea Level _____________________________ 10Figure 6. Delta-V With Variation in O/F at Vacuum and Sea Level ________________________ 11Figure 7. Delta-V/c with Variation in Mass Ratio _______________________________________ 12Table I. RD-120 Engine Dimensions ____________________________________________________ 2Table II. RD-120 Performance Parameters ______________________________________________ 3Table III. Results of Analysis Using Design Mixture Ratio _________________________________ 13
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1INTRODUCTIONThe RD-120 engine is a Russian liquid fuel rocket engine developed and used 1976-1982. Thedeveloping company is NPO Energomash, previously Gasdynamic Laboratory - ExperimentalDesign Bureau. The engine is designed to operate at altitude as an upper stage and incorporates alarge exit area ratio nozzle to accommodate this. The engine was used on the Zenit launch vehiclesecond stage. It has a non-vectorable nozzle. The propellant delivery is provided via one mainand two boost turbopumps. Ignition and turbopump operation is accomplished by using a singlepreburner. The engine was extremely reliable during its service. 177 RD-120s were built andtested 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 chamberpressure. The engine’s specific impulse (Isp) in vacuum is listed as 350 sec. Though the Isp of theKerosene/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 downto 85% nominal thrust. This allows for less fuel volume to be carried on the vehicle. Dimension ofthe engine are detailed below in Table I.
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2 Table I. RD-120 Engine Dimensions Engine Dimensions Characteristic Dimension (mm, kg)Length 3872Diameter 1954Dry Mass 1125Wet Mass 1285Chamber Diameter 320Characteristic Length, L* 1274Contraction Ratio 1.74 (unitless)Throat Diameter 183.5Exit Diameter 1895Exit Area Ratio 106.7 (unitless)Chamber Length 2992
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3ANALYSISA review of the performance of this engine is desired. To accommodate this, a trade study ofmixture ratios was run using the same operating pressure and real configuration’s dimensions tofind the optimum performance. This was then compared to the predicted and actual performanceof the real configuration of the RD-120. The real contraction ratio, throat diameter, exit arearatio, 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 equationfor this is shown below. C12 H 24 (l ) 18O2 (l ) 12CO2 ( g ) 12 H 2O( g ) Equation 1. RP-1 and LOX Reaction EquationThis 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 EquationAlong 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 therockets performance. The system’s level of control of the mixture ratio is a property of the feedsystem and is rated at ±10%. Since the real configuration’s mixture ratio is O/F=2.6, this meansthat the mixture ratio is really anywhere between 2.34≤O/F≤2.86. A list of the real configuration’sperformance 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
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4Specific Impulse (Isp)The vacuum and sea level Isp was calculated using NASA CEA with the methods previouslydescribed. The vacuum Isp is our primary data set of interest as the RD-120 was used as an upperstage in vacuum conditions. Values for the sea level performance are not available for the realengine to compare to. In Figure 1 it is seen that the optimum mixture ratio is 3 and produces anIsp of 376 sec. Also shown on this figure are the values for the stoichiometric and realconfiguration mixture ratios. It is seen that in this case the stoichiometric mixture ratio does notproduce the best performance. The real configuration mixture ratio also does not produce thelargest optimum specific impulse. However, it will be seen in further investigation why thismixture ratio was chosen by the designers.Flame TemperatureThe flame temperature was also calculated using the NASA CEA code and the results are plottedversus mixture ratio in Figure 2. The mixture ratio, 3, that produced the highest specific impulsealso has the highest flame temperature of 3835 K. This becomes a heat transfer managementproblem. In this case there are two options, increase the mixture ratio (fuel lean) or decrease themixture ratio (fuel rich). In either case there is a sacrifice that must be considered in trading heatmanagement with lower specific impulse. In the real case, 0.8% in Isp is lost but the temperatureis also dropped 1.1%. It is also seen that the stoichiometric mixture ratio produces a higher flametemperature and lower Isp than the real configuration mixture ratio. This is more evidence as towhy the designers decided to go fuel rich instead of fuel lean.Thrust Coefficient and ThrustThe thrust coefficient is an important parameter in nozzle design and represents the amplificationof thrust due to the supersonic expansion of the nozzle compared to if the nozzle exit area wereequal 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 ismainly due to the gas properties at those particular mixture ratios. A sharp change in pressureratio and specific heat ratio occurs that forces the thrust coefficient lower. This however does notaffect the thrust as seen in Figure 4. Also first notice that the thrust coefficient is not at its highestfor 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 alsoseen 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 shownbelow. F m v2 P2 A2 Equation 3. Thrust Equation
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5Characteristic Velocity C*The characteristic velocity was calculated using the effective exhaust velocity and thrustcoefficient obtained from NASA CEA according to the equation below. c C* CF Equation 4. Characteristic Velocity EquationThe optimum mixture ratio seen in Figure 5 producing the highest C* is 2.5 and the real mixtureratio is very close with a C*=1876 m/sec. The characteristic velocity is a function of the propellantcombustion and acts as a measure of how well the combustion chamber is designed. This meansthat since the C* is near its highest for the real mixture ratio, the chamber of the RD-120 wasproperly designed for optimum combustion at the design mixture ratio. Because some energy islost through the nozzle exhibiting incomplete combustion during real tests, it is beneficial to lookat C* efficiency. This is a ratio of the actual C* and the theoretical C* and is a measure of thecompletion of energy release in the propellants and creation of high temperature and highpressure gases. The C* efficiency for the real mixture ratio is 94% which is in the typical range of92-99.5%.Delta-V ΔvThe delta-V is the amount change in velocity the rocket engine would produce if it was to fly withjust its own mass plus fuel mass. It was calculated using the equation below and plotted in Figure6. mf mtb v c ln mf Equation 5. Delta-V Equation for Rocket OnlyIn this equation mf represents the dry mass of the engine and the added term represents the massof propellant consumed. The plot shows confirmation that the real mixture ratio provides one ofthe highest values of over 15 km/sec. On a vehicle this number would be lower due to thestructural mass and payload. The calculated inverse mass ratio for the RD-120 is 69.012 and it canbe seen in Figure 7 how much delta-V can be expected when taking the effective exhaust velocityinto consideration.
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6 400 Stoichiometric 380 Real O/F = 2.6 O/F = 3.429 Isp = 373 sec Isp = 371 sec 360 340Specific 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
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7 4500 Stoichiometric Real O/F = 2.6 O/F = 3.429 4000 T = 3816 K T = 3793 K 3500 3000Temperature (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
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8 2.1 Stoichiometric O/F = 3.429 CF = 1.99 2 Real O/F = 2.6 CF = 1.95 1.9Thrust 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
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9 1000000 Stoichiometric Real O/F = 2.6 O/F = 3.429 900000 F = 886 kN F = 881 kN 800000 700000 600000Thrust (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
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10 2000 1800 1600 Real O/F = 2.6 C* = 1876 m/sec Stoichiometric 1400Characteristic 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
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11 18000 16000 14000 Real O/F = 2.6 12000 Stoichiometric Δv = 15412 m/sec O/F = 3.429 Δv = 15387 m/secDelta-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
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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 = 4.234Delta-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
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13RESULTS AND ASSUMPTIONSThe 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 spaceIt is desired to compare the results of this analysis with the performance values provided inTable II. After performing the mixture ratio trade study there is agreement that the designmixture ratio O/F=2.6 is an appropriate value to use. Additional analysis was performed atthis mixture ratio using the mixture ratio control. This gives results for performanceparameters 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 CalculatedO/F 2.6 2.34 2.6 2.86Isp (vacuum), sec 350 366 373 376Flame Temp, K 3670 3700 3793 3815Thrust Coeff 1.95 1.91 1.95 1.98C*, m/sec 1760 1870 1876 1860Thrust, kN 834 870 886 892Exit Pressure, bar .127 .084 .103 .120In this table it is seen that the values are close but do not match those of the real engine. Inmost cases the values are larger than those of the real engine. This is due to the assumptionsthat are stated previously as well as a few other possibilities. One main possibility for losses isthe efficiency of which the propellants are injected into the combustions chamber and mixedprior 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 besmaller than that of the physical throat.
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14CONCLUSIONSNo real safety issues have been found with the RD-120 rocket engine. Liquid engines of thissize have some inherent risk. The engine requires a complex turbopump system to deliverthe 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 gasesflowing in very close proximity to the propellants in the feed system. In 1983 a fuel systemleak led to the destruction of Soyuz T-10-1 on the launch pad. The RD-120 was not used onthis vehicle however.From the analysis completed it appears there is only a little room for improvement viaoptimizing the mixture ratio. Some extra performance could be obtained with little increasein system risk it appears. The areas for improvement on a rocket engine of this size includematerials studies. A trade could be performed to see if only the throat region and chamberneed to be regeneratively cooled and the exit cone made of an ablative composite such ascarbon/carbon or carbon phenolic. This would save weight and reduce system complexity.The combustion chamber could have small dimensional changes with test to determine aconfiguration with more complete combustion, however the current design probably alreadyresulted in much testing for this as well as combustion stability. Because the RD-120 alreadyuses staged combustion, an extremely efficient system, it is already a very well designedengine 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.
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