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na, 2919 NACA RM E52H27 tema RESEARCH MEMORANDUM INVESTIGATION OF RAM-JET AFTERBURNING AS A MEANS OF VARTING EFFECTIVE EXHAUST NOZZLE AREA By Eugene Perchonok and Fred A. Wilcox Lewis Flight Propulsion Laboratory Cleveland, Ohio CLABSIFIRD DOCTACENT rasa material contains information attecting th National Deterse cftho United Salra wítiia the mennins ot the espictage Java, Tito 18, U.6.C., Gocm. TOS and 794, the transmission ar revelation cf wiioh ln amp mamor to insuthorized person 1ã probíbitea by law, NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON 3 “Ig, FEÍ 15 November 17, 1952 ctz TECH LIBRARY KAFB, NM em 4, na 0143328 NATTONAL ADVISORY COMMITTEE FOR ABRONAUTICS NACA RM E52H27 RESEARCH MEMORANDUM INVESTIGATION OF RAM-JET AFTIERBURNING AS A MEANS OF YARTING EFFECTIVE EXHAUST NOZZLE AREA By Eugons Perchonok anã Fred A. Wilcox SUMMARY A discussion of the flow mechanism in an afterburner extension to a rem-jet engine is presented. Anajysis indicated that with an after- burner, ram-jet performance equal to that of the engine equipped with a variable-geometry converging exit nozzie coulã result. The use of an afterburner requires, under some conditions, heat addition to a super- sonic stream. , An experiments! evaluation of the operation of a 16-inch ram jet equipped with an afterburner was conducted at Mach numbers of 1.8 and 2.0 in the Lewis 8- by 6-foot supersonic wind tumnel. The contraction ratio of the primary nózzle was 0.59 anã of the secondary nozzle, 1.0. The experimental data confirmed the proposed method of afterburner oper- ation. Although the combustion efficiencies of the primary and after- turner sections were lower than the general level of combustion effi- ciency of conventional rem jets, It was possible et a Mach number of 2.0 and with essentially no change in the diffuser-exit Mach number to increase the net interns) thrust from 0.36 to 0.74 merely by increasing the total fuel flow. INTRODUCTTON, The flight paths of several proposed ram-jet-powered guided-missile configurations are such that engine operation during the boost phase is desirable. In addition it is usually adventageous to provide a thrust margin at the cruise Mach number for meneuvering purposes. With fixed- geometry nozzle configurations, such requirements cause off-design inlet diffuser operation which, in turn, is accompenied by total-pressure- recovery losses and drag increases. A veriable-geometry outlet ie one method of dealing vith this prob- lem; however, the inherent mechanical complications make other means 2613 NACA RM E52H27 €< total-temperature ratio Subscripts: A ahesã of normal shock a afterburner Pp primary combustor t total or over-all º free stream 1 engine inrlet 3 dlffuser exit and combustion-chember. inlet 4 primary nozzle inlet 5 primary nozzie throat [] primery nozzle outlet and afterburner inlet 7 engine exit AFTERBURNER OPERATION The afterburning ram jet is illustrated schematicslly in figure 1. It consists simply of s primary combustor, station 3 to station é, and a converging-diverging nozzle, station 4 to station 6, followed by a constant-area afterburner, station 6 to station 7, and a second nozzle, station 7, which in figure 1 has a contraction ratio of 1. Both nozzle- throst areas depend on the engine-performance requirements; however, the +hroat ares of the second nozzle must be larger than that of the primary nozzle. As an aid in understanding afterburner operation, the typical vari- atiop of the thrust coefficient as well as the Mach numbers at stations tions 3 eng 5 to 7 is given in figure 2 as a Function of Ta, the after- burner total temperature ratio. For efficient cruise operation, all heat addition ie completed before the gases resch the throat of the primary nozzle, station 5. The velocity at the throat is sonic, and 1? a suffi- cient pressure ratio is available, the velocity of the gases in the afterburner section and et the engine outlet is supersonic. Fuso 4 ni NACA RM E52H27 In order to increase the thrust above the cruise value without changing the diffuser exit Mach number Mas heat is addeã to the gases as they pass through the afterburner. The primary-combustor fuel-air ratio and consequently the primary-combustor total-temperature ratio Tous are mainteined constant. A small amount of afterburner heat addition will decelerate the supersonic flow in the afterburner to a limiting Mach number of 1 at the throat of the second nozzle, etation 7. When further heat is added, & normal shock forms in the region between sta- tion 5 anã station 6 and the value of M., which depends on the primary- nozzle expansion ratio, drops from « supérsonic to a subsonic value, Because o? Bubsonic heat addition, the flow ig then reaccelerated to My of 1. With a constant-area afterburner, no change in thrust can occur until Mo is 1 ang the normsl shock forms in the diverging section of the primsry nozzle. As larger amounts of heat are added. subsonically in the afterburner, the normal shock that forms initially at stetion 6 is driven toward the throat of the primary nozzle, station 5, and the thrust increases. With sufficient heat addition in the afterburner, Mç “becomes legs than unity and further thrust increases are accompanied, às with'the conventional engine, by a reduction in My and off-design inlet operation. A one-dimensional analysis and a detailed discussion of the operation of a constant-area afterburner of the type described are given in the appendix. . APPARATUS AND PROCEDURE . The experimental. verification of the operation of a ram-jet after- burner was made with & 18-inch ram-jet engine. The investigation was conducted at stream Mach numbers of 1,8 and. 2.0 in the Lewis 8- by 6-foot supersonic wind tunnel with the same engine as used in reference 3. The coordinates of? this engine are given in table I. As illustrateã in fig- ure 3, the engine was strut-mounted: and a water-cooled pressure rake was used in determining the exit momentum. In this preliminary investigation, advantage was taken of the high afterburner-inlet temperature and no afterburner fleme holder was ugseê. The primary combustor (fig. 4) consisteã of a can-type flame holder identical with that used in reference 3 and. was lgnited by the flame of a vortex-type pilot burner, located in the domstream enê of the center- body. The pressure-drop variation across the can is given in figure 5, Two primary-combustor fuel-injector configurations were required, one to obtain data below and the other above a Ma of approximately 0.24 (rig. 4). Above a Ms of approximately 0.24, a Bingle-stage injector consisting, of eight epray nozzles mounted in an annular manifolà was used. to introduce fuel into the can just downstream of the pilot. Below a Es of approximately 0.24, four nozzles were added to the original menifola anã, in adáition, a second annular manifolã of 12 nozzles vas used. Hed ET a 6 NAS NACA BM E52H27 The total-temperature ratio across the engine “+ is basesã on pres- sure measurements made at the afterburner outlet. The total-temperaturê ratio across the primary combustor Tp is computed from the relation (see reference 5) Ms 4Tp=EK (1) The value of the constant K was determined from the case of zero after- burner fuel flow anã a temperature rise in the afterburner sufficient to just decelerate the flow to & M; of 1. (It was assumed that Tg was 1.35 and Yr; was 1.28 in making this computetion.) Thus, with K kiowa, only Me was required to evaluate T . When afterburner fuel was injected ahead of the primary nozzle,“equation (1) was adjusted for the eúded mass. By the nature of the caleulation, Tt. includes any effect que to vaporization or combustion of afterburner"fuel oceurring before the throat of the primary nozzie. PRESTA The total temperature ratio across the afterburner Ta is computed from ta se ato (2) Once the component temperature rise ratios tp anã Ta vwere known, the combustion efficiencies of both the primary combustor and the afterburner could be computed. . Txe experimental data presented herein are in terms of an internal- thrust coefficient. If it ls desired to compute the propulsive thrust, - the external drag coefficients at critical flow were 0.165 and 0.175 at stream Mach numbers of 2.0 anã 1.8, respectively (reference 4). RESULTS AND DISCUSSION , The performance trends observed experimentally generally agreed with the postulated afterburner flow mechanism; however, wlth the actual after- burner, some deviations from the theoretical trends were observed and will be discussed. As predicted, large changes in thrust were observed at essentislly constant diffuser-exit Mach numbers M by merely increasing the after burner fuel flow. The net internal-thrust variation for a typical tran- eition region (operation between cruise and maximum thrust), Mg, 0.219t0.006 and Mo» 2.0, is given in figure 8 as a function of the total-temperature ratio across the engine. Also show are the corre- sponding component total-temperature ratios across the primary burner and the afterburner, Tp end T;. At the My considered, the cruise fuel-air ratio was 0.057 (table II). During'transition the primary £/a was reduced to 0.058 to give the same value of Ms. CGE 2613 NACA RM E52H27 CER 7 The net thrust coefficient was increaseã from & cruise value o? 0.36 to O.74 with afterburning. Since Tp remained essentially con- etent, the rise in T, vas due to the heat added to the gases as they passed through the afterburner. Apparently some of the afterburner fuel. burned in the primery combustor, because in order to obtain the seme value of Ty, a lesner primary fuel-air ratio £/a was required with the afterburner on than at cruise without afterburning (see table II), and in addition a very slight increase in Tp was generally observed as the afterburner fuel flow was increased, The E at which transition occurred Was a function o? Tp ená, therefore, the primary £/a. The effect of the primary £/e on the transition at Mo of 1.8 and 2.0 is shom in figure 9. Although- the crítical diffuser-exit Mach number at both streem conditions is approximately 0.20, it was felt that the minimum transition Mach numbers between 0.21 anã 0.22 were sufficient to demonstrate afterburner operation. According to the theory (see equation (1)), the transition M et a given primary £/a should remain constant. Experimentally, however, & slight decrease in Mz was observed as the thrust was raised by an increase in the afterburner fuel flow. This reduction in M, is due in pert to the change in mass flow through the primary nozzle throat caused by the injection of afterburner fuel and in part to the burning of afterburner fuel in the primary combustor. The meximum thrusts obtained with the sfterburner at a My of 2.0 have teen compared in figure 9(b) with the theoretical end experimentally observeê thrusts 0f the basic engine with the primary nozzle removed. The total engine length was held constant. The data indicate that the maximum thruste observed with the afterburnrer-equipped engine fell below the equivalent fixed-gecmetry engine having an exit-nozzle-area ratio 'of 1. Although it is possible that the afterburner-equipped engine had not yet reached ite maximum thrust limit (the maximum fuel. flow was timiteã by the capacity of the pumpíng system), the flow losses caused by the primary nozzle will probably reduce the maximum thrust of this engine below that available from the equivalent fixed-geometry engine. It is also probable that the cruise thrusts obtained with the afterburner inoperative and fuel sprayed only through the primary injector will be slightly below the thrusts of the equivalent fixed-geometry cruise engine because of afterburner flow losses. A Getaileã examination of the operation obtained in the primary as well as the afterburner combustion chambers is of interest. The data ata M, of 2.0 have therefore been reduced and sre tabulated in table II. Althougê temperature-rise ratios across the engine T, es great as 3.9 were realized, the corresponding combustion efficiencies were quite low. “E 2613 a! TD Sn 5 NACA RM ES2H27 É UN SE] os 9 that is, T;/T6 =1 (or My = Mg > 1). Agreement was observed, however, between the theoretical and experimental total-pressure ratios across the entire engine for the M; = 1 cruise data. The slight difference between the theoretical and the experimental curves probably represents the flow and burning losses through the primery nozzle, neither of wnich was included in the derivetion of the thsoretical curve. A similer differ- ence between theory and experiment appears for the maximum-thrust data between a Mz of 0.24 to 0.27. For a value of Mz less then 0.24, it is not clear from the data whether or not, for the maximm thrusts “recorded, the normal shock had been driven to the throat of the primary nozzle. Because of the large pressure losses accompanying supersonic heat addition (see appendix) in the afterburner during cruise operation with = 1, the totel-pressure ratio across the engine at cruise is lees than the total-pressure ratio across the engine under maximum-thrust operation. Since cruise operation should be more efficient than high- thrust operation, this result appears, at first, incongruous. Because | ideally the large afterburner total-presgure losses at cruise occur with- out a loss of momentum, they do not cause a reduction in either the cruise thrust or efficiency of the basic engine. In order to avoid the unintentional heat addition in the afterburner during cruise operation, the primary nozzle wes translated 10 inches downstream of its original position; thus an over-all afterburner length of 165 inches resulted. With this modification it was possible to reduce the heat addition in the afterburner anê thus maintain supersonic veloe- ities at the afterburner exit. The Mach number across the discharge jet varied from sonic at the center to values as high as 1.75 near the periphery, but the &verage values varied from 1.4 to 1.7. The 0.59 primary nozzle contraction ratio resulte in over-expanded flow in the diverging section of the nozzle during ideal cruise operation. Because such over-expansion leads to & reduction in the cruise thrust, the selection of the primary-nozzle expansion ratio and the afterburner nozzle area may be dictated in part by Flight conditions as well as thrust variation requirements. Although these experimental results are preliminary, they are nevet- theless sufficient to confirm the mode of operation proposed for .a rsm-jet afterburner end to demonstrate the practicability of obtaining large thrust changes with a fixed-geometry engine and at an essentially constant diffuser-exit Mach number merely by increasing the fuel flow. No air-flow or combustion-stability problems were encountered in changing from cruise to high-thrust operation. The thrust varieê continvously in the transition region and could be maintained, if desired, et any intermediate value between cruise and maximum thrust. Nato 2») 10 Nildo NACA RM ES2H27 Graphite proved a satisfactory primary nozzle material, and after several hours of operation there was no evidence of surface pltting or erosion of any kind. (The photograph of figure 6 was taken after approxi- “o mately 4 hr of operation.) o Experience indicated that a single fuel system serving both the pri- mery combustor and the afterbumner may be adequate. However, more effi- cient combustion than obtained in this investigation is desirable. Tn addition to the possibility of causing a reduction in the cruise thrust, lt may be desirable to avoid the added weight and drag of the afterburner section during cruise operation. Since the maximum thrust for many applications will be requireã only during boost end acceleration - to cruise periods, it may be possible to jettison the afterburner section during the cruise phase of the flight. An alternate method of operation may be to translate the primary nozzle into. the afterburner section and thus teke advantage of afterburner length for improved primary combustor performance during eruise operation. Although the discussion has been concerneã with the application of this system to a ram-jet engine, the same device may be applied to a turbojet engine requiring thrust variation and variable-exit-area control. This latter application may be as a second-stage afterburner or as a single-stage afterburner downstream o? a choked nozzle. CONCLUDING REMARKS The preliminary investigation reported herein regarding the appli- eation of an afterburner to a ram-jet engine indicates that thruste equal to those of the equivalent basic engine designed for either cruise or maximum thrust or. any thrust value between these limits can be obtained merely by & variation in fuel flow. This large thrust variation (over 100 percent) was accomplished with only a negligible variation in diffuser- exit Mach number anã therefore essentially no change in inlet flow or nacelle drag. The afterburner thus provides a fixed-geometry ram-jet engine with performance characteristics similar to those previously obtained only with & ram-jet engine equipped with a continuously variable exit nozzle. The experimental observations confirm the theory proposed as em explanation for the operation of the afterburner. In addition, the data demonstrate that supersonic heat addition (burning) in a constant-area duet is not only physicaily possible but generally follows the relations derived from simple one-dimensional analysis. o = Lewis Flight Propulsion Leboratory — National Advisory Committee for Aeronautics =." Cleveland, Ohio Neres... ' 12 . Dies NACA RM E52H27 These expressions are generally given for the case of a subsonic initial Mech number (M,< 1), references 5 end 7, but they also spply for the supergonte case (M > 1). The solution of equations (AL), (42), and (43) for M >1 and both Y end Ya equal to 1.5 is plotted in figures 11(a), 11(b), and 11(c), respectively. The simultaneous solution of equations (42) and (A3) (fig. 11(ã)) yielãs the variation in total-pressure retio with inttial Mach number amd heat-addition parameter. The limiting condition (flow decelerated to sonic velocity) is shom as & dashed line. Super- sonic flow is decelerated by constant-sarea heat aúdition and in the limit, the final Mach number reaches 1. The higher the initial Mach number, the larger the amount of heat the stream can absorb before deceleration to a of 1. In contrasgt to subsonic heat addition, a small amount of heat sdãition to a supersonic stream causes large total-pressure losses. Another mechanism whereby a nonvisçous, compressible fluid may be decelerateãd in a constant-ares duct from a supersonic stream Mach number M >1 to Ma = 2 is illustrated in the following sketch: Normal shock mma Station 12 2 M >1 M, =1 If a normal shock forms at station 1, the heat addition that begins at station 1' takes place at an initial Mach number M: « 1; sufficient heat is added to accelerate this subsonie stream to a Mo oí 1. Since tne derivation of equations (AL) to (A3) diã not require isentropic flow tetween stations 1 anã 2, these same equations also describe this letter process. It may therefore be concluded that ín a constant-area duct, if the inlet flow is eupersonic, the same heat-addition parameter is required for s final Mach number Mo of 1, whether or not a normal shock is involved in the process. (A constant valbe of Yy Is asgumed.) The values of the total- and static-pressure retios and the hest-addition parameter between stations 1 and 2 may either be obtained from figure 11 or from the normal shock and subsonic heat-addition relations. In either case, ldentical velues result. Cruise operation. - In ideel cruise operation (fig. 12(a)), ell primary-compustor heat addition is completeê by station 5 and no heat is added in the afterburner. Thus, Mg equals M; and if flow losses are Taz h 2613 13 Firlasims NACA RM E52H27 TIA ignoreã, the afterburner exerts no influence on the thrust delivered by the basic engine. The primary nozzle contraction ratio determines the amount of heat addition required in the primary combustion chamber in order to obtain the design velue of Mz. The primary nozzle expansion retio selecteã met avoid overexpansion of the flow and the attendant thrust losses. Transition operation. - Thrust increases with no change in the aiffuser-exit Mach number may be obtained by adding heat to the gases as they flow through the afterburner section; however, before the meximum- thrust condition can be reached, the afterburner passes through a region of transition operation. This transition region oceure in two stages. Initially, heat is added to a supersonic stream. &=1 Mg > My >1 and because momentum is conserved, no thrust change 1s observed. This condition is shown schematically in figure 12(b). In deriving the pressure, temperature, and Mach number relations thet describe afterburner operation (fig. 13), nonviscous, one-dimensional, compressible flow was assumed. It was further assumed that all after- turner heat addition occurreã between stations 6 and 7 and that Te/7 = 1 The ratio o? the specific heats y was helã constant at 1.3. The theoreticel retation between the heat-addition parameter By T p2 27 07 and My (where My >1) 1s given in figure 1s(a) for several Rg Tê primary-nozzle-expansion ratios. At any given primary-nozzle-expansion ratio As/Ag the addition of heat to the supersontc stresm causes a flow deceleration, and im the limiting case, the stream cap be decelerated to a final Mach number M, of 1 (fig. 12(c)). During cruise operation of an actual engine, burning is usually not yet completed at the throat of the primary nozzle, end some heat addition may occur in the afterburner. If this is the case, the engine wlll actually be operating in the first stage of the transition region. At the condition of sufficient afterburner heat addition to obtain a M of 1, the process can be either heat addition to a supersonic stream or, &s previously indíceted, a normal shock at station 6 followed by sub- sonic heat addition. The values cf the heat-addition parameter and the Rlol? NACA RM E52H27 15 REFERENCES Aron.: Booster Rem Jet Study. CAE Rep. No. 375, Continental Aviation and Engineering Corp. (Detroit, Michigan), Nov. 29, 1948. (Contract W-33-038-ac-13371.) Minton, S. J.: Study of Self-Acceleration Potentialities of Ramjet Engines. Rep. No. 5145, Proj. 20-108, Marquardt Aircraft Co. May 22, 1951. (Contract W-53-038-se-19050, E.0. No. 506-150.) Perchonok, Eugene, Wilcox, Fred, and Pennirgton, Donald: Effect of Angle of Attack and Exit Nozzle Design on the Performance of & 16-Inch Rem Jet at Mach Numbers from 1.5 to 2.0. NACA RM ES1G26, 1951. Nussãorfer, T., Wilcox, F., and Ferchonok, E.: Investigation at Zero Angle of Attack of & 16-Inch Ram-Jet Engine in 8- by 6-Foot Super- sonie Wind Tunnel. NACA RM ESOLO4, 1951. Perchonok, Eugene, Sterbentz, William H., and Moore, Stanley E.: Indirect Methods for Obtaining Ram-Jet Exhaust-Gas Temperature Applied to Fuel-Metering Control. NACA RM ETE27, 1948. Hicks, Bruce L.: Addition of Heat to a Compressible Fluiã in Motion. NACA ACR ESA29, 1945. - Hawthorne, W. R., and Cohen, H.: Pressure Losses and Velocity Changes Due to Heat Release and Mixing in Frictionless, Compressible Flow. Rep. No. E.3997, British R.A.E., Jean. 1944. poruaensiEd NACA RM ES2H27 16” TABLE I - 16-INCE RAM-JET ENGINE COORDINATES A B (6) po | - - Station Location A | B Cc Miscellaneous (in.) Gn) (mo) (in) -5.05| Tip cof spike | O 4.0 0.48 -3.0 0.94 -2.0 l.4l -1,0 1.88 EO] Lip of inlet | 2,34 5.05 Lip redius 0.032 2.0 2.i.e 5.13 5.57 2.0 3.10 5.30 D.54 3.0 3.56 5.45 5.89 4.0 3.58 5.59 5.83 6.0 3.94 5.83 6.07 8.0 4.21 6.03 6.28 10.0. 4.40 6.20 6,45 12.0 4.52 6,36 6.81 14,0 4,58 6.48 6.74 16.0 4.60 6,58 6.82 18.0 4.58 6.61 6.85 30.0 4,44 | Straight | Straight 46.0 4.02 taper taper 59.0 3.08 7.75 8.13 63.0 2.43 7.45 68.4 |Enã of (o) 7.58 center body ei Pilot air 1.5 inlets 93 Pilot maxi- ,4.0 Cylindrical mum diameter section 107 Station 3 3.3 8.00 148 Primary noz- zle intet 187 Engine exit 8.00 8.13 ctoz Station o 1 2 5 4 5 6 7 e S ] eS < im E 17 Diffuser Primary combustor | Primary Afterburner “o nozzle Ra Figure 1. - Schematic Giagram of afterburner-equipped ram-jet engine. LEEESE WE VOYN 2613 NACA BM E52H27 Mg 1.0 t Supereonie Supersonic 1.0 I [ Pi£êuser design Ma E Decreasing High trust Cruise Inereas ing ] l | | I Í 1.0 Naa —— Increasing Tt, — >» Figure 2. - Fredicted effect of afterburner heat addition on thrust coefficient Cr,n and internal Mach numbers My, M5, Mg, anê My. o E Sremnecemeram 19 
