Article Experimental Study of the Combustion Eﬃciency in Multi-Element Gas-Centered Swirl Coaxial Injectors
Seongphil Woo 1,* , Jungho Lee 1,2, Yeoungmin Han 2 and Youngbin Yoon 1,3 1 Department of Aerospace Engineering, Seoul National University, Seoul 08826, Korea; [email protected] (J.L.); [email protected] (Y.Y.) 2 Korea Aerospace Research Institute, Daejeon 34133, Korea; [email protected] 3 Institute of Advanced Aerospace Technology, Seoul National University, Seoul 08826, Korea * Correspondence: [email protected]
Received: 27 October 2020; Accepted: 16 November 2020; Published: 19 November 2020
Abstract: The eﬀects of the momentum-ﬂux ratio of propellant upon the combustion eﬃciency of a gas-centered-swirl-coaxial (GCSC) injector used in the combustion chamber of a full-scale 9-tonf staged-combustion-cycle engine were studied experimentally. In the combustion experiment, liquid oxygen was used as an oxidizer, and kerosene was used as fuel. The liquid oxygen and kerosene burned in the preburner drive the turbine of the turbopump under the oxidizer-rich hot-gas condition before ﬂowing into the GCSC injector of the combustion chamber. The oxidizer-rich hot gas is mixed with liquid kerosene passed through combustion chamber’s cooling channel at the injector outlet. This mixture has a dimensionless momentum-ﬂux ratio that depends upon the dispensing speed of the two ﬂuids. Combustion tests were performed under varying mixture ratios and combustion pressures for diﬀerent injector shapes and numbers of injectors, and the characteristic velocities and performance eﬃciencies of the combustion were compared. It was found that, for 61 gas-centered swirl-coaxial injectors, as the moment ﬂux ratio increased from 9 to 23, the combustion-characteristic velocity increased linearly and the performance eﬃciency increased from 0.904 to 0.938. In addition, excellent combustion eﬃciency was observed when the combustion chamber had a large number of injectors at the same momentum-ﬂux ratio.
1. Introduction Turbopump-driven liquid-rocket engines can generally be divided into gas-generator-cycle, expander-cycle, and staged-combustion-cycle types. In a gas-generator system, a portion of the propellant is burned within the gas generator to drive the turbine and is then discharged to the outside; here, the propellant supply pressure required for turbopump is relatively low, making it easy to develop and manufacture. However, such systems have the disadvantage of a propellant eﬃciency that is several percentage points lower than that in other systems, as the propellant burned inside the gas generator is dumped outside the turbine [1,2]. Expander-cycle systems are relatively simple, using a structure in which the vaporizing propellant is cooled in a cooling channel before entering the main combustion chamber that drives the turbine; however, it is diﬃcult to increase the combustion-chamber pressure due to the limited degree of propellant heating in the regenerative cooling channel. Staged-combustion-cycle conﬁgurations achieve eﬃciency by sending the turbine-driven gas back into the main combustion chamber for reburn, although the pressure required by the pump is high and the structure is complicated [2,3]. Staged-combustion-cycle engines are the most widely used, due to their ease of re-ignition and high speciﬁc impulse. This type of engine can be further
Energies 2020, 13, 6055; doi:10.3390/en13226055 www.mdpi.com/journal/energies Energies 2020, 13, 6055 2 of 14 divided into fuel-rich-combustion and oxidizer-rich-combustion subtypes. An example of fuel-rich combustion is the RS-25 engine used on the space shuttles, with liquid hydrogen and liquid oxygen as propellants . In oxidizer-rich combustion, kerosene and liquid oxygen are mainly used as propellants: examples include Russia’s RD-170, or the American AR-1; engines that use liquid methane and liquid oxygen as propellants, such as the Raptor or BE-4, are also being developed [4–6]. The gas-centered-swirl-coaxial (GCSC) injector used in oxidizer-rich staged-combustion-cycle engines has a structure that mixes the hot gas burned in the oxidizer-rich condition in the preburner with the fuel that cools the combustion chamber as it passes through the cooling channel; the mixture is then sprayed into the combustion chamber. Such injectors have mainly been developed by Russia in the past, but unfortunately, records of their characteristics are hard to ﬁnd. Recently, with the development of staged-combustion-cycle engines by the United States, China, and other countries, studies of GCSC injectors have been conducted. In particular, the characteristics of the propellant mixture have been determined through spray testing, numerical analysis, and ﬂow-instability modeling according to the shape of the injector. Jeon et al.  studied spray characteristics according to the gas–liquid = 2 2 momentum-ﬂux ratio (J ρgUg/ρlUl ) and recess length through cold-ﬂow tests. Kim et al.  analyzed the spray characteristics of the ﬂuid at atmospheric and high pressure in terms of the momentum-ﬂux ratio via spray visualization. This showed that, as the momentum-ﬂux ratio increases, the spray angle decreases until a certain limit is reached. Kang et al.  experimentally observed that the size of the sprayed droplet breaks up rapidly as the gas-ﬂow rate increases. Balance et al.  visualized the ﬂame-stabilization region under high-pressure conditions using gas–liquid propellant. Using numerical analysis, the propellant mixing, ﬂame formation, and ﬂow instability were studied under supercritical conditions for various shapes of 2D and 3D injectors [11,12]. Park et al. [13,14] investigated the instabilities and characteristics of the ﬂow of the gas and the liquid, respectively, for the GCSC injector; when a perturbation was applied to each ﬂuid, the momentum of the gas was transferred to the liquid as the momentum-ﬂux ratio increased in both cases and the gain value increased accordingly. The previous studies have focused mainly on the characteristics of injector mixing in single-injector or non-reacting ﬂows, but it is necessary to study how the momentum-ﬂux ratio relates to combustion eﬃciency in tests in which a large number of injectors are applied in actual combustion environments. In this paper, the eﬀects of the momentum-ﬂux ratio upon the combustion-performance of a full-scale 9-tonf staged-combustion-cycle engine combustion chamber were analyzed experimentally under variation of the shape of the injector, the mixture ratio, and the number of injectors. Among the various methods used to increase the performance of full-scale rocket engines, there are ways to increase the eﬃciency within the combustion chamber. This eﬃciency can be obtained by identifying the combustion’s characteristic velocity, which is the virtual thrust per unit mass with which the pressure in the combustion chamber acts on chamber’s neck. It can be obtained from the pressure of the combustion chamber, the combustion rate of the propellant, and the cross-sectional area of the nozzle as [1,2,15] . c∗ = PCCAt/m (1) To this end, a 9-tonf full-scale staged-combustion-cycle engine was manufactured, and combustion tests were performed according to the momentum-ﬂux ratio of the GCSC injector in an actual combustion environment with various combustion pressures and mixture ratios. The combustion’s characteristic velocity results were calculated and analyzed.
2. Material and Methods As shown in Figure1, The injector used in the experiment is a GCSC injector from which hot gaseous oxygen (GOx) is sprayed from the center and kerosene used to cool the combustion chamber is injected through the annular hole. The injector has three shapes depending on the velocity of GOx and its outer diameter; the A-type injector is designed such that the ﬂow velocity of GOx injected from the center will be 100 m/s; the B-type is designed to have a ﬂow rate of 75 m/s at the center, such that the Energies 2020, 13, x FOR PEER REVIEW 3 of 14
Energiesthat the2020 inner, 13, 6055diameter of the injector from which GOx is sprayed is larger than in the A-type 3case. of 14 TheEnergies C-type 2020, injector13, x FOR isPEER designed REVIEW to have GOx flow rate of 100 m/s, which is the same as the A-type;3 of 14 however, as the total number of injectors used in the full-scale combustor is smaller than the A-type, that the inner diameter of the injector from which GOx is sprayed is larger than in the A-type case. innerthe injector diameter diameter of the injector is designed from whichto be GOxlarger is th sprayedan that is of larger the thanA-type in theto equalize A-type case. the Theamount C-type of The C-type injector is designed to have GOx flow rate of 100 m/s, which is the same as the A-type; injectorpropellant is designedflow injected to have into the GOx combustion ﬂow rate ofchamber. 100 m/s, All which injectors is the have same a recess as the ratio A-type; of 1.0. however, ashowever, the total as number the total of injectorsnumber of used injectors in the used full-scale in the combustor full-scale iscombustor smaller than is smaller the A-type, than thethe injectorA-type, diameterthe injector is designed diameter to is be designed larger than to be that larger of the th A-typean that to of equalize the A-type the amountto equalize of propellant the amount ﬂow of injectedpropellant into flow the combustioninjected into chamber. the combustion All injectors chamber. have All a recess injectors ratio have of 1.0. a recess ratio of 1.0.
Figure 1. Schematic of the gas-centered-swirl-coaxial Injector.
Combustor heads were manufactured with two types of arrangements according to the injectors’ Figure 1. Schematic of the gas-centered-swirl-coaxial Injector. shape as presented in TableFigure 1. 1. FigureSchematic 2 shows of the gas-centered-swirl-coaxial the combustor heads for Injector. the A-type and the B-type configurationsCombustor comprise heads were 61 injectors. manufactured A hypergolic with two injector types of is arrangements placed at the accordingcenter of the to thefaceplate injectors’ of Combustor heads were manufactured with two types of arrangements according to the injectors’ shapethe combustor as presented head, in with Table 6 main1. Figure injectors2 shows in the the first combustor row and heads12 main for injectors the A-type in the and second. the B-type In the shape as presented in Table 1. Figure 2 shows the combustor heads for the A-type and the B-type conﬁgurationsthird row, 18 baffle comprise injectors 61 injectors.are arranged A hypergolicto suppress injector combustion is placed instability. at the The center 4th of row the comprises faceplate configurations comprise 61 injectors. A hypergolic injector is placed at the center of the faceplate of of6 sections the combustor of 3 main head, injectors, with 6 mainwith injectorsindividual in thebaffle ﬁrst injectors row and dividing 12 main each injectors section in theto prevent second. the combustor head, with 6 main injectors in the first row and 12 main injectors in the second. In the Incombustion the third instability. row, 18 ba ﬄThee injectorsC-type combustor are arranged head to has suppress a total of combustion 37 injectors: instability. a hypergolic The injector 4th row is third row, 18 baffle injectors are arranged to suppress combustion instability. The 4th row comprises comprisesagain centrally6 sections located,of 3 with main 6 injectors,main injectors with individual in the first barow,ﬄe 12 injectors baffle dividinginjectors eachin the section second to row prevent and 6 sections of 3 main injectors, with individual baffle injectors dividing each section to prevent combustion6 sections of instability.2 main injectors The C-type in the combustorthird row, with head individual has a total baffle of 37 injectors:injectors adividing hypergolic each injector section. is combustion instability. The C-type combustor head has a total of 37 injectors: a hypergolic injector is again centrally located, with 6 main injectors in the ﬁrst row, 12 baﬄe injectors in the second row and again centrally located, with 6 main injectors in the first row, 12 baffle injectors in the second row and 6 sections of 2 main injectors in theTable third 1. row, Injector with head individual configuration. baﬄe injectors dividing each section. 6 sections of 2 main injectors in the third row, with individual baffle injectors dividing each section. Injector Dg Do Dinj Linj Ug Recess Quantity Table 1. Injector head conﬁguration. Head [mm]Table 1. [mm]Injector head[mm] configuration. [mm] [m/s] Ratio
InjectorA-Type Head Quantity61 Dg [mm]7.4 Do [mm]10.4 Dinj [mm]0.8 Linj [mm]39.5 Ug [m100/s] Recess Ratio 1.0 Injector Dg Do Dinj Linj Ug Recess B-Type Quantity61 8.4 11.4 0.8 39.5 75 1.0 HeadA-Type 61[mm] 7.4 10.4[mm] 0.8[mm] 39.5[mm] 100[m/s] 1.0Ratio C-TypeB-Type 36 61 8.49.6 11.413.5 0.80.9 39.539.5 75100 1.0 1.0 A-Type 61 7.4 10.4 0.8 39.5 100 1.0 C-Type 36 9.6 13.5 0.9 39.5 100 1.0 B-Type 61 8.4 11.4 0.8 39.5 75 1.0 C-Type 36 9.6 13.5 0.9 39.5 100 1.0
(a) (b) (c)
Figure 2. Combustor Head: ( a)) A-type, A-type, ( b) B-type, ( c) C-type. (a) (b) (c) The combustor consists of a separate combustor head with injectors injectors and and a a combustion combustion chamber, chamber, as shown shown in in Figure Figure 3.3 .The TheFigure combustion combustion 2. Combustor chamber chamber Head: can (a) can beA-type, reused be reused (b) becauseB-type, because ( cit) C-type.is designed it is designed to have to a havebolting a bolting structure that allows separation and recombination of the combustor heads for various The combustor consists of a separate combustor head with injectors and a combustion chamber, as shown in Figure 3. The combustion chamber can be reused because it is designed to have a bolting EnergiesEnergies 20202020,, 1313,, x 6055 FOR PEER REVIEW 44 of of 14 14 structure that allows separation and recombination of the combustor heads for various injector shapes.injector Within shapes. the Within curved thepart curved of the combustor part of the head combustor through head which through the pre-burned which the GOx pre-burned flows, an oxidation-resistantGOx ﬂows, an oxidation-resistant coating was appliedcoating to protect was applied the wall’s to protect surface the from wall’s the hot surface gas flowing from the from hot thegas turbopump, ﬂowing from and the to turbopump, prevent ignition and to due prevent to particles ignition that due may to particles occur in that the may preburner. occur in The the preburner. The combustion chamber is designed with regeneration and ﬁlm cooling, and the ZrO2 combustion chamber is designed with regeneration and film cooling, and the ZrO thermal-barrier andthermal-barrier Ni-Cr coatings and are Ni-Cr applied coatings to protect are applied the comb to protectustion chamber’s the combustion inner chamber’swall from the inner combustion wall from flame.the combustion The high-pressure ﬂame. The kerosene high-pressure that is dischargedkerosene that from isdischarged the fuel pump from flows the fuel into pump the nozzle ﬂows end, into coolsthe nozzle the nozzle end, coolsinner the wall,nozzle and innerpasses wall, through and passesa film-cooling through ring a ﬁlm-cooling located at ringthe bottom located of at the the combustionbottom of the chamber combustion cylinder. chamber After cylinder. cooling After the wa coolingll of the the wallbottom of thecombustion-chamber bottom combustion-chamber cylinder, kerosenecylinder, passes kerosene through passes a through film-cooling a ﬁlm-cooling ring located ring at located the top at of the the top combustion-chamber of the combustion-chamber cylinder andcylinder then andenters then the enters fuel manifold the fuel manifoldof the combus of thetion-chamber combustion-chamber head before head being before sprayed being into sprayed the injector.into the injector.To measure To measure the temperature the temperature of kerose of kerosenene passing passing through through the the cooling cooling channel channel of of the the combustioncombustion chamber, chamber, temperature sensors werewere installed installed in in the the manifold manifold of of the the combustion-chamber combustion-chamber headhead fuel, fuel, two two film-cooling ﬁlm-cooling rings, and manifold atat the the end end of of the the nozzle. nozzle. This This combustor combustor was was designed designed toto test test the the performance characteristics characteristics of of the the GCSC GCSC injector; injector; thus, thus, the the length length of of the the combustion chamberchamber and and the the expansion expansion ratio ratio of of the the nozzle nozzle were were designed designed to to be be relatively relatively short short as as 2.3 2.3 to to fit ﬁt the the vertical-combustion-testvertical-combustion-test facility.
FigureFigure 3. 3. 9Tf9Tf combustion-chamber combustion-chamber test test model: model: ( (aa)) schematic schematic ( (bb)) actual actual shape. shape.
TheThe 9-tonf 9-tonf staged-combustion-cycle staged-combustion-cycle engine engine consis consiststs of of a a preburner, preburner, a a combustor, combustor, a a turbopump, turbopump, pneumatic-controlpneumatic-control valves,valves, and and piping, piping, as shown as shown in Figure in 4Figure. This engine4. This is designedengine is for designed the development for the developmentof preburner of and preburner combustor and combustor technology technology for staged-combustion-cycle for staged-combustion-cycle engines. engines.Unlike generalUnlike generalliquid-rocket liquid-rocket engines thatengines combine that acombine combustor a combustor and a turbopump, and a turbopump, the turbopump the and turbopump the combustor and the are combustorhere assembled are here as separate assembled structures as separate with structures a certain distance with a certain between distance them to between facilitate them disassembly to facilitate and disassemblyreplacement and of the replacement combustor. of In the addition, combustor. to protect In addition, the engine’s to protect components the engine’s from accidentscomponents that from may accidentsoccur during that the may ﬁring occur test, during major parts the suchfiring as test, the servomajor valves parts and such the as hypergolic-propellant the servo valves and ampule the hypergolic-propellantare placed on opposite ampule sides with are placed a thick on steel opposite plate between sides with them. a thick The steel combustion-pressure-control plate between them. The combustion-pressure-controland mixture ratio control valves and of mixture the combustor ratio control are installed valves at of the the exit combustor of the ﬁrst- are and installed second-stage at the exitfuel of pumps, the first- respectively. and second-stage Liquid oxygen fuel pumps, and kerosene respectively. propellants Liquid supplied oxygen and from kerosene the test facilitypropellants enter suppliedthe turbopump, from the where test facility they are enter highly the pressurized turbopump, and where supplied they are to the highly preburner. pressurizedPropellants and supplied burned toinside the preburner. the preburner Propellants become burned oxygen-rich inside hotthe gasespreburner and enterbecome the oxygen-rich gas injectors hot of gases the combustor. and enter theThe gas fuel injectors supplied of fromthe combustor. the ﬁrst-stage The fuel suppli pumped of from the turbopump the first-stage passes fuel throughpump of the the combustor’s turbopump passesmixture through ratio control the valvecombustor’s and enters mixture the nozzle ratio end cont ofrol the valve combustor and toenters cool thethe combustion-chambernozzle end of the combustorwall through to coolingcool the channels combustion-chamber and ﬁlm cooling. wall Thereafter, through the cooling fuel supplied channels to theand combustor film cooling. head Thereafter,is injected intothe fuel the supplied combustion to the chamber combustor through head theis injected annular into injector, the combustion along with chamber the pre-burned through thehigh-temperature annular injector, oxygen-rich along with gases. the pre-burned The engine high is started-temperature by supplying oxygen-rich high-pressure gases. The helium engine to the is started by supplying high-pressure helium to the starting turbine of the turbopump. The engine- Energies 2020, 13, 6055 5 of 14 Energies 2020, 13, x FOR PEER REVIEW 5 of 14 combustionstarting turbinetest is performed of the turbopump. by supplying The engine-combustion hypergolic propellant test is performed to the preburner by supplying and the hypergolic combustor in accordancepropellant with to the a preburner pre-set sequence. and the combustor in accordance with a pre-set sequence.
FigureFigure 4. 9Tf 4. 9Tfstaged-combustion-cycle-engine staged-combustion-cycle-engine model model forfor thethe ﬁring firing test: test: (a )(a left) left view view (b) ( frontb) front view. view.
Combustion tests were conducted for pressures between 81 bar and 105 bar (i.e., approximately 10 bar Combustion tests were conducted for pressures between 81 bar and 105 bar (i.e., approximately± of the design combustion pressure of 95 bar), and for mixture ratios between 2.2 and 2.75 ±10 bar of the design combustion pressure of 95 bar), and for mixture ratios between 2.2 and 2.75 (i.e., (i.e., approximately 12% of the design mixture ratio of 2.5), as shown in Figure5. The velocities of ± approximatelythe injected ±12% gas were of the calculated design basedmixture on theratio pressure, of 2.5), temperature,as shown in density,Figure and5. The ﬂow velocities rate of the of the injectedinjector, gas aswere obtained calculated from the based gas propertieson the pressure, of the preburner temperature, and the density, combustor and head; flow thisrate was of the done injector, to as obtainedallow analysis from the of the gas combustion properties e ﬃofciency the prebur accordingner and to the the momentum-ﬂux combustor head; ratio. this Also, was the done material to allow analysisproperties of the and combustion ﬂow rate of keroseneefficiency ﬂowing according into the to combustor the momentum-flux were measured 1000ratio. times Also, each the second material propertiesbased onand data flow from rate pressure of kerosene and temperature flowing sensorsinto the and combustor ﬂowmeters were installed measured in the engine 1000 andtimes the each secondtest based facility on in data order from to calculate pressure the speedand temperat of the liquidure propellant.sensors and From flowmeters the law of installed mass conservation, in the engine and the velocitytest facility of the in oxygen-rich order to calculate hot gas, U theg, and speed the axial of the liquid liquid velocity, propellant.Ul, were From calculated, the law and of the mass thickness of the liquid ﬁlm sprayed from the annular fuel injector, tl, was derived using the expression conservation, the velocity of the oxygen-rich hot gas, , and the axial liquid velocity, , were of Suyari and Lefebbre [1,16]: calculated, and the thickness of the liquid film sprayed. from the annular fuel injector, , was derived 4mg using the expression of Suyari and Lefebbre U[1,16]:g = ; (2) πρgdg 4 . = ml ; (2) Ul = ; (3) πρ t (d t ) l l l − l . 0.25 tl = 3.1= dlmlµl/ρl∆Pl ; . (4) (3) ( ) . =3.1( / ∆ ) . (4)
Figure 5. Operating points of the 9-Tf staged-combustion cycle engine. Energies 2020, 13, x FOR PEER REVIEW 5 of 14 combustion test is performed by supplying hypergolic propellant to the preburner and the combustor in accordance with a pre-set sequence.
Figure 4. 9Tf staged-combustion-cycle-engine model for the firing test: (a) left view (b) front view.
Combustion tests were conducted for pressures between 81 bar and 105 bar (i.e., approximately ±10 bar of the design combustion pressure of 95 bar), and for mixture ratios between 2.2 and 2.75 (i.e., approximately ±12% of the design mixture ratio of 2.5), as shown in Figure 5. The velocities of the injected gas were calculated based on the pressure, temperature, density, and flow rate of the injector, as obtained from the gas properties of the preburner and the combustor head; this was done to allow analysis of the combustion efficiency according to the momentum-flux ratio. Also, the material properties and flow rate of kerosene flowing into the combustor were measured 1000 times each second based on data from pressure and temperature sensors and flowmeters installed in the engine and the test facility in order to calculate the speed of the liquid propellant. From the law of mass conservation, the velocity of the oxygen-rich hot gas, , and the axial liquid velocity, , were calculated, and the thickness of the liquid film sprayed from the annular fuel injector, , was derived using the expression of Suyari and Lefebbre [1,16]:
4 = ; (2)