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Stability rating tests for the length-optimization of baffles in a liquid propellant combustion chamber using a pulse gun
Aerospace Science and Technology 12 (2008) 214–222 www.elsevier.com/locate/aescte
Stability rating tests for the length-optimization of baffles in a liquid propellant combustion chamber using a pulse gun Hong Jip Kim a,∗ , Seonghyeon Seo a , Kwang Jin Lee a , Yeoung Min Han a , Soo Yong Lee a , Young Sung Ko b a Korea Aerospace Research Institute, PO Box 113, Yuseong, Daejeon 305-600, Republic of Korea b Department of Aerospace Engineering, Chungnam National University, Daejeon 305-764, Republic of Korea
Received 2 January 2006; received in revised form 4 June 2007; accepted 4 June 2007 Available online 14 June 2007
Abstract Stability rating tests using a pulse gun, one type of artificial disturbance device, have been performed to optimize and to limit the axial baffle length of a combustor. To determine whether the combustor is stable when it is exposed to strong external perturbations or pressure oscillations, the decay time and amplitude ratio have been analyzed to quantify the stabilization capacity of the combustor. A baffle not big enough to cover the flame zone considered as the collision plane of the F-O-O-F impinging injector was not able to suppress external perturbations sufficiently. The combustion stability boundary limit of the present combustor can be determined with respect to the axial baffle length. © 2007 Elsevier Masson SAS. All rights reserved. Keywords: Stability rating test; Combustion instabilities; Baffle; F-O-O-F impinging injector; Pulse gun; Decay time; Decay rate; Amplitude ratio; Dynamic stability boundary
1. Introduction Combustion instabilities, which are often induced by the interaction of the combustion zone and the acoustic field in a combustion chamber, have gained great attention in most liquid rocket development programs. These undesirable phenomena can cause deterioration of performance and severe physical damage to rocket combustor hardware, consequently resulting in a mission failure [1–6]. To overcome combustion instability problems, the modification of the characteristics of acoustic or combustion fields is quite necessary. A passive control method has been widely used as a means to modify acoustic fields. For instance, an ablative-cooled, 90-mm baffle with a 1-hub/6-blade configuration has been successfully adopted in flight tests [7,8]. When using a baffle as a passive control device, problems such as additional cooling and thrust loss need to be solved. Especially in the case of internal cooling systems, the amount of additional coolant could change the performance characteristics * Corresponding author. Tel.: +82 42 860 2814; fax: +82 42 860 2602.
E-mail address: [email protected] (H.J. Kim). 1270-9638/$ – see front matter © 2007 Elsevier Masson SAS. All rights reserved. doi:10.1016/j.ast.2007.06.003
by varying the O/F ratio of an overall engine system. Moreover, the axial baffle length with respect to the chamber diameter is recommended as roughly 20% [4], but it is necessary to decrease the axial baffle length to minimize the above-mentioned problems. In this respect, the verification and optimization of the axial baffle length for combustion stability is very important. Although a rocket engine operates in a steady state, intrinsic pressure fluctuations exist in the combustion chamber. On the whole, rockets having ±5% pressure fluctuations with respect to the mean chamber pressure are assumed to be statically stable [1,2]. Since there are always inevitable operational changes due to the variation of ambient conditions such as gravitational acceleration especially in a pressurized system, a combustion chamber which is considered as statically stable can become unstable. Therefore, the necessity of stability rating tests as a final verification is essential for the elucidation of operational changes affecting combustion stability [1–4,6]. Combustion response characteristics in a rocket engine can be explained by sub-critical bifurcation phenomena which can be illustrated in the operating conditions such as chamber pres-
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Nomenclature A amplitude An,max amplitude ratio of dynamic pressure defined in Fig. 4 AR amplitude ratio defined in Fig. 4 D diameter of the chamber d diameter of the injector orifice f0 frequency of unstable mode LB axial baffle length or height O/F oxidizer-fuel mixture ratio p pressure t time
Greek letters α ττ
decay rate recommended reference decay time
Subscripts ch cr F max O
chamber critical value or state fuel maximum oxidizer
tem return to normal operation after any transients. The second method has the advantage of requiring a relatively small number of tests. To introduce artificial external perturbations, several methods are widely used such as explosive bomb, pulse guns, gas flow, etc. [1,2]. The pulse gun is advantageous because there is no disturbing effect on the combusting flow before the explosion and the explosion timing can be electronically controlled. In the present study, stability rating tests have been performed by adopting a pulse gun device which has been developed to be able to generate two separate artificial perturbations. The objectives of this study are the optimization of axial baffle length and the indigenous establishment of stability rating test procedures. 2. Experiments Fig. 1. The amplitude of dynamic pressure in a combustion chamber with respect to operation parameters (a sub-critical bifurcation diagram) [3].
sure, manifold pressure and O/F ratio, etc., versus the amplitude of pressure fluctuations, shown in Fig. 1 [3], where the physically reachable state is shown as a solid line and the unreachable state as a dotted line, representing a hysteresis character. According to characteristics of engines and operating conditions, there are three different regimes shown in Fig. 1. An absolutely stable combustor would be a perfect solution and an absolutely unstable one useless. In reality, most cases show potentially stable characteristics; below the threshold amplitude (Acr ) stabilization can be attained. The span of this threshold value corresponds to the dynamic stability margin of the object combustor. To assess the stability margin of an engine, a spontaneous instability method and an artificial initiation method can be used [1,2,6]. The spontaneous instability method is based on conducting a large number of combustion tests and flights to determine statistical stability. From previous experiences with the F-1 liquid rocket engine development [6], it is obvious that the process of designing a rocket system statistically is a very time-consuming trial-and-error method. On the contrary, the artificial initiation method is concerned only with the response to transients in the system operation and requires that the sys-
The schematic diagram and the geometrical dimensions of the present combustor, characterized as a pressurized liquid rocket system burning liquid oxygen (LOx) and kerosene [7,8], are shown in Fig. 2(a). The present combustor utilizes 228 splittriplet(F-O-O-F) impinging injectors in a H-type array, which can be found in Fig. 2(b). Each injector has two orifices for fuel with the diameter (dF ) of 1.6 mm, and two orifices for oxidizer with dO = 2.2 mm. The impinging angle between fuel and LOx stream is 20◦ which has been selected to prevent thermal damage on the injector faceplate. The chamber diameter is 420 mm and the nozzle diameter is 310 mm. The axial length to the nozzle throat is 452 mm and the nozzle expansion ratio is about 5.04. Nominal conditions are a chamber pressure of 13.8 bar, mass flow rates of 18 and 42 kg/s for fuel and LOx, respectively, resulting in a typical mixture ratio of 2.34. The injector for the ignition is located in the center of the faceplate, and triethylaluminum (TEA), a hypergolic fluid, is injected by pressurized kerosene flow at the beginning of a test. The installed 1-hub/5-blade baffles, shown in Fig. 2, are made of stainless steel and are arranged in a position of minimal interference with injectors in the faceplate. Axial baffle lengths, LB are selected as 50, 70 and 85 mm, which correspond to about 12, 16.7 and 20% of the chamber diameter, respectively. To introduce artificial disturbances into the combustion chamber and to monitor consequent pressure fluctuations, the ports for a
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(a)
(b) Fig. 2. Schematic diagram of the present combustor and injector faceplate with an installed baffle, and the location of a pulse gun port and two dynamic pressure sensors (FCC1, FCC2). (a) Geometrical configuration of the present combustor. (b) An injector faceplate with baffle.
pulse gun and two dynamic pressure sensors, FCC1 and FCC2 (PCB piezoelectronics, model 123A24), are located at the axial distance of 100 mm from the faceplate. These three ports are located at 50◦ , 240◦ and 275◦ respectively along the chamber wall with respect to the reference line, which can be seen in Fig. 2. The sampling time of these sensors is 3.9062 × 10−5 s and sampling frequency is 25,800 Hz, resulting in the Nyquist frequency of 12,800 Hz for preventing aliasing. And the effective frequency range of the present transducers is about 1–20,000 Hz and the filtering range is 30–10,000 Hz, which is below the Nyquist frequency. To reduce the cost of the development program and experimental risks, multiple generations of external perturbations by a
pulse gun device will be favorable. In this study, a pulse gun device having two ports of explosive charges shown in Fig. 3 has been adopted. To prevent abnormal explosion at an unwanted instant due to overheating caused by hot gases in the combustion chamber, the nitrogen purging system was introduced [9]. Shock waves generated from this pulse gun device are vertically introduced into the combusting flow. In general, pressure fluctuations by external perturbations, p(t) show following traces with an exponential decay shown in Fig. 4 when a stable combustion response is attained, p(t) = pmax e−αt sin(2πf0 t)
(1)
where pmax is the initial maximum amplitude by an external perturbation, α the decay rate, and f0 the frequency of a con-
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cerning unstable mode. Various parameters for evaluating the stabilization capacity of rocket engines in relation to external perturbations are introduced with schematics shown in Fig. 4 [3, 10,11]. One of these various parameters is An,max , the ratio of the maximum amplitude of an artificial disturbance to the mean pressure fluctuation in a steady state. Another of these parameters is the maximum amplitude of an artificial disturbance with respect to mean chamber pressure, AR, which can be a good indicator of stabilization capacity. In this study, these two amplitude-ratio parameters have been investigated to elucidate dynamic stability characteristics. Generally, An,max and AR should be larger than the threshold value (Acr in Fig. 1) to properly evaluate a stability margin. But if the practical value of An,max and AR do not exceed this threshold value, it would be favorable to make these amplitude-ratio parameters as large as possible. Roughly An,max is recommended as O(10–102 ) and if combustion stabilization is established, the maximum of An,max can be regarded as the maximum dynamic stability capacity of the combustor. The other parameter is the well-known decay time estimated by transient pressure fluctuations. In this study, the decay time is quantified as the elapsed time when the amplitude of pressure fluctuations becomes 1/e times of pmax in Eq. (1). It has been
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conservatively recommended that the decay time be estimated reversely in the time domain [10,11]. In the case of Russian RP-1 engines operating at a high pressure range, a reference decay time of ττ = 15–25 ms is recommended [11]. Alterna1/2 tively, criterion such as ττ 1250/f0 has been established by JANNAF standards [12], and by adapting this for the first tangential mode, ττ is approximately 30 ms in our case. In the present study, ττ has been conservatively selected as 15–25 ms. 3. Results Combustion instability phenomena have occurred in the initial development stage of the present combustor. These instabilities were identified as typical spontaneous instability phenomena due to the first tangential mode having a 1680 Hz resonant frequency and its harmonics (without any higher acoustic modes), which can be easily observed in Fig. 5 [7,8]. To eliminate these phenomena, a baffle with a 1-hub/5-blade configuration has been adopted in this study and the stability capacity of the axial baffle length, LB , has been investigated. First,
(a)
Fig. 3. The schematic of two pulse guns, a cavity and an adaptor for the chamber installation.
(b)
Fig. 4. A conceptual pressure trace and some important parameters which can be evaluated from the experimental data.
Fig. 5. Time history of dynamic pressure and FFT signals in the initial development stage of the present combustor showing the characteristics of combustion instabilities [7,8]. (a) Time history. (b) FFT of (a).
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artificial external disturbances were introduced in the baffled combustion chamber of LB = 85 mm (LB /Dch ≈ 20%) under nominal conditions (pch = 13.8 bar, O/F = 2.34). The unperturbed pressure oscillation is about ±3% with respect to the steady chamber pressure, and this condition can be assumed to be statically stable. Two pulses of 1.0 and 1.8 g explosive charges were generated 1.0 s and 0.5 s respectively before normal shutdown. Consequent raw pressure fluctuations, pressure fluctuation filtered through 1 k–2 kHz bandpass, and the acceleration data are shown in Fig. 6. The sensitivity to acceleration of the present transducers is about 0.002 psi/g. The maximum acceleration was less than 700 g during combustion tests which
can be found in Fig. 6(c). Therefore, the maximum pressure amplitude induced by acceleration is about 1.4 psi or 0.097 bar. Consequently, the pressure induced by acceleration via present dynamic pressure sensors can be negligible. An,max caused by these two perturbations was 150 and 85, and the AR was 168 and 73%, respectively. These values were much higher than the inherent noise level. No particular unstable combustion modes have been observed. The corresponding decay time for each pulse is smaller than 2 ms, and it can be concluded that a baffle of LB = 85 mm can sufficiently suppress relatively large external perturbations, showing dynamically stable characteristics under the An,max and AR of O(102 ). Next, a stability rating test was done for the case of LB = 50 mm (LB /Dch ≈ 12%) under the same conditions as the previous case. A small LB implies that the baffle cannot compart active combustion zone sufficiently so the combustion field could be prone to instability. Therefore, to minimize the possible damage to the combustion chamber and the test facility, decreased explosive charges of 0.6 and 1.4 g have been used. In
(a)
(a)
(b)
(c)
(b)
Fig. 6. Various dynamic pressure characteristics and acceleration data of an 85 mm baffled combustion chamber under nominal conditions. (a) Raw signal. (b) Filtered (1 k–2 kHz) signal. (c) Acceleration data.
Fig. 7. Filtered dynamic pressure characteristics and FFT of a 50 mm baffled combustion chamber under nominal conditions. (a) Filtered (1 k–2 kHz) time history. (b) FFT waterfall of (a).
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addition, the pulse guns were triggered 0.6 s and 0.3 s before normal shutdown. Due to electrical cut-off in the second pulse gun, only the first pulse gun generated shock waves. The time history of the pressure fluctuations is shown in Fig. 7(a). Frequency analysis shows that there exists an unstable mode in the range of 1.2 k–1.6 kHz, which can be seen in Fig. 7(b). The phase difference of two dynamic pressures measured by FCC1 and FCC2 sensors is about 155◦ , and this mode was identified as the first tangential (1T) mode coupled with combustion. On the contrary, there was a time gap, 1.8 ms shown in Fig. 7(a) between the local pressure maximums in the two dynamic pressure measurements. This characteristic seems to result from a certain tangential flow or a flame oscillation in the combustion chamber. Similar behavior was observed in the single injector’s and sub-scale combustion test previously [13,14]. The flames exhibited oscillating lift-off shapes with asymmetric configurations, despite being under atmospheric conditions. By considering these characteristics, flames can be thought not to have stable lift-off heights and these phenomena may create pressure fluctuation characteristics as shown in Fig. 7(a). Low frequency pressure fluctuations of about 170 Hz resonance were also observed, shown in the waterfall of Fig. 7(b). These low frequency oscillations were observed throughout the test although unstable combustion responses attenuated and a stable operation was achieved with less than ±5% oscillation. Also, similar oscillations were monitored during the tests of a different combustor. This is assumed to be due to the coupling with the feeding system of our test facility. These unstable responses can be explained by combustion characteristics. One of the possible reasons is the hydrodynamic character of the adopted impinging injectors [13–15]. As in Fig. 8(a), the F-O-O-F impinging injector has two collision planes, at the axial distances about 30 mm and 50 mm from the faceplate which can be evaluated geometrically by considering the mass flow rates of kerosene and LOx. And the axial profile of the heat release rate can be shown schematically like Fig. 8(a) [13]. After injection, mixing, atomization and vaporization processes, the zone having the maximum reaction rate is assumed to exist near the second collision plane which is axially 40–50 mm away from the faceplate and plays an active role in interacting with acoustic modes. In this respect, a baffle of LB = 50 mm cannot completely separate the active flame zone from these calculated planes. Moreover, it is known that there is the longitudinalization of transversal acoustic modes in the baffle tip region [16,17], so a stiff pressure gradient can occur near the baffle tip, observed easily in Fig. 8(b), which shows the acoustic pressure contours of the first tangential mode. This can cause strong coupling between a stiff acoustic pressure gradient and an active combustion zone, resulting in a high susceptibility to a certain combustion instability phenomenon. In this regard, the 50 mm baffle cannot damp out pressure fluctuations sufficiently and pressure oscillations remained for a relatively longer duration than in the previous case when a baffle of LB = 85 mm was used.
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(a)
(b) Fig. 8. Hydrodynamic characteristics of a F-O-O-F impinging injector [13] and acoustic field illustrating a possible acoustic-flame coupling [16,17]. (a) Hydrodynamic characteristics of a present injector. (b) Acoustic pressure contours of the first tangential mode showing the longitudinalization near baffle tip.
Although pressure oscillations of high amplitude were observed during a certain time period, the decay time was about 6 ms shorter than generally recommended values of 15–25 ms [1,3,6,12]. We can now judge that the simple comparison of decay time is not a necessary and sufficient condition for combustion stability assessment. Instead, it is necessary to consider overall combustion characteristics. To further investigate these characteristics, an off-design condition prone to combustion instability was chosen. The possibility of the occurrence of combustion instability increases generally with pressure. According to the laminar flamelet model [15,18,19], flames near an extinction regime have a large amplification effect on acoustic pressure. In the case of a kerosene/LOx system, a low O/F ratio condition with respect to the nominal operating condition corresponds to a fuel-rich near-extinction regime. Therefore, a high pressure and low O/F ratio condition can be assumed to be the most prone to combustion instability. Although not mentioned herein, other off-
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design conditions were tested and confirmed to be relatively more stable than the above-mentioned off-design conditions. The combustion chamber of LB = 50 mm (LB /Dch ≈ 12%) under this condition (pch = 15.9 bar, O/F =2.00) has been tested with the same explosive charge masses and triggering times, and the results are shown in Fig. 9 only for the first external perturbation. Highly unstable pressure fluctuations were observed and sustained with a large amplitude for 60 ms, a relatively long time period compared with the previous results under nominal conditions. However, after 60 ms, the unstable oscillating pressure suddenly attenuated and disappeared, and returned to a stable condition. Responses to the second perturbation have shown qualitatively similar characteristics with those of the first one, having a relatively short decay time. In Fig. 9(b), the unstable pressure response characteristics occurred in about 170 Hz and 1.2 k–1.4 kHz frequency ranges. The former is thought to be due to the coupling with the feeding system, as explained earlier. The latter frequency range is
(a)
not exactly the same as in the previous test, the reason of which may be due to the dynamic characteristics of the impinging injectors adopted in this study. On the whole, impinging injectors are known to have relatively broad-band frequency characteristics [20]. Also, a temperature decrease due to a low O/F ratio could cause this frequency difference. According to the equilibrium analysis [21], the averaged sonic velocity under nominal condition is about 1256 m/s and that under low O/F ratio condition is about 1228 m/s. So the frequency shift by the decrease of sonic velocity can be estimated to be about 50 Hz and this characteristic would play considerable role in decrease of resonant frequency. Consequently, the combustion chamber with a baffle of LB = 50 mm is thought to be unstable. Although the combustion response using a 50 mm baffle is not fully unstable, it cannot be considered as stable. To support this tentative conclusion, a chamber of LB = 70 mm (LB /Dch ≈ 16.7%) was also tested under high pressure and a low O/F ratio. As in the case of LB = 85 mm, there were not any particular unstable modes observed, which can be seen in Fig. 10. Decay times were within 1.5 ms for both of the external perturbations. Therefore, both combustors, LB = 70 and 85 mm, are thought to have a large stability margin within the specified operating window. All test results are summarized in the operational working window, which are shown in Fig. 11. In the pre-specified working window, the present combustor has a large stability boundary with LB 70 mm. However, in the case of LB = 50 mm, this combustor exists in the vicinity of marginal stability under the nominal and off-design conditions. Characteristics of decay time, decay rate, An,max , and AR with respect to LB are shown in Fig. 12. For the cases of LB = 70 and 85 mm (LB /Dch ≈ 16.7, 20%), pressure fluctuations decayed within 1.6 ms and the amounts of decay rate were larger than 0.5, on the average. The decay time of LB = 85 mm is slightly larger than that of LB = 70 mm due to the different sizes of explosive charges used in the pulse gun device. Therefore, a clear quantitative comparison cannot be made. However,
(b) Fig. 9. Filtered dynamic pressure characteristics of a 50 mm baffled combustion chamber under off-design conditions. (a) Filtered (1 k–2 kHz) time history. (b) FFT waterfall of (a).
Fig. 10. Filtered (1 k–2 kHz) dynamic pressure characteristics of a 70 mm baffled combustion chamber under off-design conditions.
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(a)
Fig. 11. The operating conditions in stability rating tests.
combustion stabilization has been achieved and the one stable system having decay time of 1.25 ms (LB = 70 mm) is not always more stable than the other stable system having decay time of 2.5 ms (LB = 85 mm). They are thought to be all stable, irrespective of the amount of decay time and decay rate. Consequently, the present combustor under both cases is regarded as being quantitatively in the stable regime in the parameter space of decay time, decay rate and amplitude ratios. The test results showing stable combustion responses do not significantly deviate from each other regardless of the operating conditions such as the nominal and the off-design conditions. On the other hand, the results of LB = 50 mm (LB / Dch ≈ 12%) have long decay time, low decay rate and are scattered. Considering these characteristics and the data of the amplitude of pressure fluctuations An,max and AR, the combustor of LB = 50 mm is thought to be in the unstable regime in terms of practical application. In addition, An,max and AR can reflect the responses of a combustion zone to external perturbations and are much more dependent on the characteristics in combustion chambers than on the size of explosive charge used in the pulse gun [8,9,11, 14]. In other words, An,max and AR would be large if combustion characteristics were very sensitive to external perturbations. Therefore, large values of An,max and AR show that a combustion chamber can be moved to an unstable regime by any undesired changes to the operating conditions. The case of LB = 50 mm, which shows unstable characteristics, reveals considerably large and scattered values compared with the LB = 70 and 85 mm test cases. As mentioned before, the effect of an axial baffle length on the thrust or specific impulse of the present combustor has been analyzed and confirmed under unperturbed conditions to validate the optimizing process of an axial baffle length. Specific impulse, Isp with respect to length is shown in Fig. 13. Intuitively, the specific impulse of the present combustor decreases with axial baffle length. Under nominal conditions, the decrease of specific impulse is relatively steep, compared with off-design conditions. In other words, a long-length baffle for stability assurance may cause considerable performance loss
(b)
(c) Fig. 12. Decay time, decay rate and the amplitude ratios with respect to three different axial baffle lengths, LB . (a) Decay time. (b) Decay rate. (c) An,max and AR.
for the present combustor. Therefore, this length-optimization process would be very important for the determination step of design confirmation. Consequently, the stability boundary of the combustion chamber in this study is assumed to be between 50 and 70 mm in axial baffle length, LB . Quantitative optimization with only a few experimental results would be insufficient and difficult; however, by synthesizing the response characteristics of a decay time and two amplitude ratios practically, the actual stability boundary exists close to LB = 50 mm.
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Fig. 13. The specific impulse, Isp with respect to axial baffle lengths.
4. Conclusion Stability rating tests on a liquid propellant rocket combustor were carried out for the evaluation of stability capacity with respect to axial baffle length. A pulse gun device capable of generating two external perturbations was adopted to determine the dynamic stability boundary of an axial baffle length. Generally, the axial baffle length with respect to the chamber diameter is recommended to be about 20%. However, it is necessary to decrease the axial length to minimize the thrust loss and cooling problems caused by installing a baffle in a combustion chamber. A decay time and two amplitude ratios of a steady-state to the induced external perturbation have been used to elucidate the stability boundary of the combustion chamber. In the case of LB = 50 mm (LB /Dch ≈ 12%), unstable combustion responses caused by external perturbations were observed. These characteristics can be explained by comparing the axial baffle lengths and the axial distances of the second geometrical collision planes of the F-O-O-F impinging injectors in the viewpoint of heat release rate. On the other hand, pressure fluctuations of large amplitude were damped out within the recommended reference decay time in the cases of larger baffle lengths, showing satisfactory stabilization capacity under a specified operating window. Therefore, the stability boundary with respect to axial baffle length is assumed to exist around 50 mm in the present combustor. Finally, the optimization process of an axial baffle length and the indigenous establishment of stability rating test procedures have been set up. Acknowledgements This work has been financially supported by Korean Ministry of Science and Technology. The authors would like to express sincere thanks to the efforts of all members in Combustion Chamber Department at Korea Aerospace Research Institute for conducting these hot firing tests. References [1] D.T. Harrje, F.H. Reardon, (Eds.), Liquid Propellant Rocket Combustion Instability, NASA SP-194, 1972.
[2] E. Laudien, R. Pongratz, R. Pierro, D. Preclik, Experimental procedures aiding the design of acoustic cavities, in: V. Yang, W.E. Anderson (Eds.), Liquid Rocket Engine Combustion Instability, in: Progress in Astronautics and Aeronautics, vol. 169, AIAA, Washington, DC, 1995, pp. 377–399. [3] A.F. Agarkov, K.P. Denisov, M.L. Dranovsky, I.A. Zavorohin, V.N. Ivanov, V.P. Pikalov, A.A. Shibanov, Injector flame stabilization effects on combustion instability, in: V. Yang, W.E. Anderson (Eds.), Liquid Rocket Engine Combustion Instability, in: Progress in Astronautics and Aeronautics, vol. 169, AIAA, Washington, DC, 1995, pp. 281–305. [4] NASA, Liquid Rocket Engine Combustion Stabilization Devices, NASA SP-8113, 1974. [5] J.J. Fang, Longitudinal mode combustion instabilities of a high-pressure fuel-rich LOX/RP-1 preburner, in: Proceedings of the 22nd JANNAF Combustion Meeting, vol. 1, 1985, pp. 429–441. [6] J.C. Oefelein, V. Yang, Comprehensive review of liquid-propellant combustion instabilities in F-1 engines, Journal of Propulsion and Power 9 (5) (1993) 657–677. [7] C.H. Sohn, W.S. Seol, S.Y. Lee, Y.M. Kim, D.S. Lee, Application of combustion stabilization devices to liquid rocket engine, Journal of the Korean Society for Aeronautical and Space Sciences 31 (6) (2003) 79–87. [8] Y.M. Han, Y.S. Ko, S.H. Kim, K.J. Lee, Y.G. Chung, W.S. Seol, S.Y. Lee, Combustion stability rating test using a pulse gun device in a liquid rocket engine, in: The 4th Symposium on Space Launch Technology, May, 2003, pp. 33–37. [9] K.J. Lee, H.J. Kim, S. Seo, I.Y. Moon, H.M. Kim, S.Y. Lee, Y.S. Ko, Experiments on development of a pulse-gun pressure-wave inducing device for stability rating test technique, Journal of the Korean Society for Aeronautical and Space Sciences 33 (3) (2005) 99–106. [10] C.H. Sohn, S.K. Kim, Y.W. Moon, Y.M. Kim, Technologies for Suppression/Protection of Combustion Instabilities in Rocket Engines for Launch Vehicle, Ministry of Science and Technology, Korea, 2003. [11] Denisov, et al., KARI combustion chamber firing tests, Report under Contract No. KARI-01-032, Phase 4, NIICHIMMASH, 2002. [12] M.D. Klem, R.S. Fry, Guidelines for combustion stability specifications and verification procedures for liquid propellant rocket engines, CPIA Publication 655, 1997. [13] C.H. Sohn, W.S. Seol, A.A. Shibanov, V.P. Pikalov, Combustion stability boundaries of the subscale rocket chamber with split-triplet impinging jet injectors, Journal of Propulsion and Power 23 (1) (2007) 131–139. [14] S. Seo, K.J. Lee, Experimental study of combustion stability assessment of injector, Journal of the Korean Society for Aeronautical and Space Sciences 32 (4) (2004) 61–66. [15] J.S. Kim, F.A. Williams, Acoustic-instability boundaries in liquidpropellant rockets – theoretical explanation of empirical correlation, Journal of Propulsion and Power 12 (3) (1996) 621–624. [16] J.M. Wicker, M.W. Yoon, V. Yang, Linear and non-linear pressure oscillations in baffled combustion chambers, Journal of Sound and Vibration 184 (1995) 141–171. [17] S.-K. Kim, H.J. Kim, W.S. Seol, C.H. Sohn, Acoustic stability analysis of liquid propellant rocket combustion chambers, AIAA Paper 2004-4142, 2004. [18] C.H. Sohn, S.H. Chung, J.S. Kim, F.A. Williams, Acoustic response of droplet flames to pressure oscillations, AIAA Journal 34 (9) (1996) 1847– 1854. [19] H.J. Kim, C.H. Sohn, S.H. Chung, J.S. Kim, Nonlinear acoustic-pressure responses of oxygen droplet flames burning in gaseous hydrogen, Journal of Mechanical Science and Technology 15 (4) (2001) 510–521 (previously, KSME Int’l Journal). [20] Y.M. Han, S.H. Kim, I.Y. Moon, H.J. Kim, J.G. Kim, W.S. Seol, S.Y. Lee, S.T. Kwon, C.J. Lee, Combustion performance tests of fuel-rich gas generator for liquid rocket engine using an impinging injector, Journal of the Korean Society of Propulsion Engineers 8 (2) (2004) 10–17. [21] S.-K. Kim, Y.W. Moon, T.S. Park, Development of chemical equilibrium cfd code for performance prediction and optimum design of LRE thrust chamber, Journal of the Korean Society of Propulsion Engineers 9 (1) (2005) 1–8.
Stability rating tests for the length-optimization of baffles in a liquid propellant combustion chamber using a pulse gun Hong Jip Kim a,∗ , Seonghyeon Seo a , Kwang Jin Lee a , Yeoung Min Han a , Soo Yong Lee a , Young Sung Ko b a Korea Aerospace Research Institute, PO Box 113, Yuseong, Daejeon 305-600, Republic of Korea b Department of Aerospace Engineering, Chungnam National University, Daejeon 305-764, Republic of Korea
Received 2 January 2006; received in revised form 4 June 2007; accepted 4 June 2007 Available online 14 June 2007
Abstract Stability rating tests using a pulse gun, one type of artificial disturbance device, have been performed to optimize and to limit the axial baffle length of a combustor. To determine whether the combustor is stable when it is exposed to strong external perturbations or pressure oscillations, the decay time and amplitude ratio have been analyzed to quantify the stabilization capacity of the combustor. A baffle not big enough to cover the flame zone considered as the collision plane of the F-O-O-F impinging injector was not able to suppress external perturbations sufficiently. The combustion stability boundary limit of the present combustor can be determined with respect to the axial baffle length. © 2007 Elsevier Masson SAS. All rights reserved. Keywords: Stability rating test; Combustion instabilities; Baffle; F-O-O-F impinging injector; Pulse gun; Decay time; Decay rate; Amplitude ratio; Dynamic stability boundary
1. Introduction Combustion instabilities, which are often induced by the interaction of the combustion zone and the acoustic field in a combustion chamber, have gained great attention in most liquid rocket development programs. These undesirable phenomena can cause deterioration of performance and severe physical damage to rocket combustor hardware, consequently resulting in a mission failure [1–6]. To overcome combustion instability problems, the modification of the characteristics of acoustic or combustion fields is quite necessary. A passive control method has been widely used as a means to modify acoustic fields. For instance, an ablative-cooled, 90-mm baffle with a 1-hub/6-blade configuration has been successfully adopted in flight tests [7,8]. When using a baffle as a passive control device, problems such as additional cooling and thrust loss need to be solved. Especially in the case of internal cooling systems, the amount of additional coolant could change the performance characteristics * Corresponding author. Tel.: +82 42 860 2814; fax: +82 42 860 2602.
E-mail address: [email protected] (H.J. Kim). 1270-9638/$ – see front matter © 2007 Elsevier Masson SAS. All rights reserved. doi:10.1016/j.ast.2007.06.003
by varying the O/F ratio of an overall engine system. Moreover, the axial baffle length with respect to the chamber diameter is recommended as roughly 20% [4], but it is necessary to decrease the axial baffle length to minimize the above-mentioned problems. In this respect, the verification and optimization of the axial baffle length for combustion stability is very important. Although a rocket engine operates in a steady state, intrinsic pressure fluctuations exist in the combustion chamber. On the whole, rockets having ±5% pressure fluctuations with respect to the mean chamber pressure are assumed to be statically stable [1,2]. Since there are always inevitable operational changes due to the variation of ambient conditions such as gravitational acceleration especially in a pressurized system, a combustion chamber which is considered as statically stable can become unstable. Therefore, the necessity of stability rating tests as a final verification is essential for the elucidation of operational changes affecting combustion stability [1–4,6]. Combustion response characteristics in a rocket engine can be explained by sub-critical bifurcation phenomena which can be illustrated in the operating conditions such as chamber pres-
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Nomenclature A amplitude An,max amplitude ratio of dynamic pressure defined in Fig. 4 AR amplitude ratio defined in Fig. 4 D diameter of the chamber d diameter of the injector orifice f0 frequency of unstable mode LB axial baffle length or height O/F oxidizer-fuel mixture ratio p pressure t time
Greek letters α ττ
decay rate recommended reference decay time
Subscripts ch cr F max O
chamber critical value or state fuel maximum oxidizer
tem return to normal operation after any transients. The second method has the advantage of requiring a relatively small number of tests. To introduce artificial external perturbations, several methods are widely used such as explosive bomb, pulse guns, gas flow, etc. [1,2]. The pulse gun is advantageous because there is no disturbing effect on the combusting flow before the explosion and the explosion timing can be electronically controlled. In the present study, stability rating tests have been performed by adopting a pulse gun device which has been developed to be able to generate two separate artificial perturbations. The objectives of this study are the optimization of axial baffle length and the indigenous establishment of stability rating test procedures. 2. Experiments Fig. 1. The amplitude of dynamic pressure in a combustion chamber with respect to operation parameters (a sub-critical bifurcation diagram) [3].
sure, manifold pressure and O/F ratio, etc., versus the amplitude of pressure fluctuations, shown in Fig. 1 [3], where the physically reachable state is shown as a solid line and the unreachable state as a dotted line, representing a hysteresis character. According to characteristics of engines and operating conditions, there are three different regimes shown in Fig. 1. An absolutely stable combustor would be a perfect solution and an absolutely unstable one useless. In reality, most cases show potentially stable characteristics; below the threshold amplitude (Acr ) stabilization can be attained. The span of this threshold value corresponds to the dynamic stability margin of the object combustor. To assess the stability margin of an engine, a spontaneous instability method and an artificial initiation method can be used [1,2,6]. The spontaneous instability method is based on conducting a large number of combustion tests and flights to determine statistical stability. From previous experiences with the F-1 liquid rocket engine development [6], it is obvious that the process of designing a rocket system statistically is a very time-consuming trial-and-error method. On the contrary, the artificial initiation method is concerned only with the response to transients in the system operation and requires that the sys-
The schematic diagram and the geometrical dimensions of the present combustor, characterized as a pressurized liquid rocket system burning liquid oxygen (LOx) and kerosene [7,8], are shown in Fig. 2(a). The present combustor utilizes 228 splittriplet(F-O-O-F) impinging injectors in a H-type array, which can be found in Fig. 2(b). Each injector has two orifices for fuel with the diameter (dF ) of 1.6 mm, and two orifices for oxidizer with dO = 2.2 mm. The impinging angle between fuel and LOx stream is 20◦ which has been selected to prevent thermal damage on the injector faceplate. The chamber diameter is 420 mm and the nozzle diameter is 310 mm. The axial length to the nozzle throat is 452 mm and the nozzle expansion ratio is about 5.04. Nominal conditions are a chamber pressure of 13.8 bar, mass flow rates of 18 and 42 kg/s for fuel and LOx, respectively, resulting in a typical mixture ratio of 2.34. The injector for the ignition is located in the center of the faceplate, and triethylaluminum (TEA), a hypergolic fluid, is injected by pressurized kerosene flow at the beginning of a test. The installed 1-hub/5-blade baffles, shown in Fig. 2, are made of stainless steel and are arranged in a position of minimal interference with injectors in the faceplate. Axial baffle lengths, LB are selected as 50, 70 and 85 mm, which correspond to about 12, 16.7 and 20% of the chamber diameter, respectively. To introduce artificial disturbances into the combustion chamber and to monitor consequent pressure fluctuations, the ports for a
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(a)
(b) Fig. 2. Schematic diagram of the present combustor and injector faceplate with an installed baffle, and the location of a pulse gun port and two dynamic pressure sensors (FCC1, FCC2). (a) Geometrical configuration of the present combustor. (b) An injector faceplate with baffle.
pulse gun and two dynamic pressure sensors, FCC1 and FCC2 (PCB piezoelectronics, model 123A24), are located at the axial distance of 100 mm from the faceplate. These three ports are located at 50◦ , 240◦ and 275◦ respectively along the chamber wall with respect to the reference line, which can be seen in Fig. 2. The sampling time of these sensors is 3.9062 × 10−5 s and sampling frequency is 25,800 Hz, resulting in the Nyquist frequency of 12,800 Hz for preventing aliasing. And the effective frequency range of the present transducers is about 1–20,000 Hz and the filtering range is 30–10,000 Hz, which is below the Nyquist frequency. To reduce the cost of the development program and experimental risks, multiple generations of external perturbations by a
pulse gun device will be favorable. In this study, a pulse gun device having two ports of explosive charges shown in Fig. 3 has been adopted. To prevent abnormal explosion at an unwanted instant due to overheating caused by hot gases in the combustion chamber, the nitrogen purging system was introduced [9]. Shock waves generated from this pulse gun device are vertically introduced into the combusting flow. In general, pressure fluctuations by external perturbations, p(t) show following traces with an exponential decay shown in Fig. 4 when a stable combustion response is attained, p(t) = pmax e−αt sin(2πf0 t)
(1)
where pmax is the initial maximum amplitude by an external perturbation, α the decay rate, and f0 the frequency of a con-
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cerning unstable mode. Various parameters for evaluating the stabilization capacity of rocket engines in relation to external perturbations are introduced with schematics shown in Fig. 4 [3, 10,11]. One of these various parameters is An,max , the ratio of the maximum amplitude of an artificial disturbance to the mean pressure fluctuation in a steady state. Another of these parameters is the maximum amplitude of an artificial disturbance with respect to mean chamber pressure, AR, which can be a good indicator of stabilization capacity. In this study, these two amplitude-ratio parameters have been investigated to elucidate dynamic stability characteristics. Generally, An,max and AR should be larger than the threshold value (Acr in Fig. 1) to properly evaluate a stability margin. But if the practical value of An,max and AR do not exceed this threshold value, it would be favorable to make these amplitude-ratio parameters as large as possible. Roughly An,max is recommended as O(10–102 ) and if combustion stabilization is established, the maximum of An,max can be regarded as the maximum dynamic stability capacity of the combustor. The other parameter is the well-known decay time estimated by transient pressure fluctuations. In this study, the decay time is quantified as the elapsed time when the amplitude of pressure fluctuations becomes 1/e times of pmax in Eq. (1). It has been
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conservatively recommended that the decay time be estimated reversely in the time domain [10,11]. In the case of Russian RP-1 engines operating at a high pressure range, a reference decay time of ττ = 15–25 ms is recommended [11]. Alterna1/2 tively, criterion such as ττ 1250/f0 has been established by JANNAF standards [12], and by adapting this for the first tangential mode, ττ is approximately 30 ms in our case. In the present study, ττ has been conservatively selected as 15–25 ms. 3. Results Combustion instability phenomena have occurred in the initial development stage of the present combustor. These instabilities were identified as typical spontaneous instability phenomena due to the first tangential mode having a 1680 Hz resonant frequency and its harmonics (without any higher acoustic modes), which can be easily observed in Fig. 5 [7,8]. To eliminate these phenomena, a baffle with a 1-hub/5-blade configuration has been adopted in this study and the stability capacity of the axial baffle length, LB , has been investigated. First,
(a)
Fig. 3. The schematic of two pulse guns, a cavity and an adaptor for the chamber installation.
(b)
Fig. 4. A conceptual pressure trace and some important parameters which can be evaluated from the experimental data.
Fig. 5. Time history of dynamic pressure and FFT signals in the initial development stage of the present combustor showing the characteristics of combustion instabilities [7,8]. (a) Time history. (b) FFT of (a).
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artificial external disturbances were introduced in the baffled combustion chamber of LB = 85 mm (LB /Dch ≈ 20%) under nominal conditions (pch = 13.8 bar, O/F = 2.34). The unperturbed pressure oscillation is about ±3% with respect to the steady chamber pressure, and this condition can be assumed to be statically stable. Two pulses of 1.0 and 1.8 g explosive charges were generated 1.0 s and 0.5 s respectively before normal shutdown. Consequent raw pressure fluctuations, pressure fluctuation filtered through 1 k–2 kHz bandpass, and the acceleration data are shown in Fig. 6. The sensitivity to acceleration of the present transducers is about 0.002 psi/g. The maximum acceleration was less than 700 g during combustion tests which
can be found in Fig. 6(c). Therefore, the maximum pressure amplitude induced by acceleration is about 1.4 psi or 0.097 bar. Consequently, the pressure induced by acceleration via present dynamic pressure sensors can be negligible. An,max caused by these two perturbations was 150 and 85, and the AR was 168 and 73%, respectively. These values were much higher than the inherent noise level. No particular unstable combustion modes have been observed. The corresponding decay time for each pulse is smaller than 2 ms, and it can be concluded that a baffle of LB = 85 mm can sufficiently suppress relatively large external perturbations, showing dynamically stable characteristics under the An,max and AR of O(102 ). Next, a stability rating test was done for the case of LB = 50 mm (LB /Dch ≈ 12%) under the same conditions as the previous case. A small LB implies that the baffle cannot compart active combustion zone sufficiently so the combustion field could be prone to instability. Therefore, to minimize the possible damage to the combustion chamber and the test facility, decreased explosive charges of 0.6 and 1.4 g have been used. In
(a)
(a)
(b)
(c)
(b)
Fig. 6. Various dynamic pressure characteristics and acceleration data of an 85 mm baffled combustion chamber under nominal conditions. (a) Raw signal. (b) Filtered (1 k–2 kHz) signal. (c) Acceleration data.
Fig. 7. Filtered dynamic pressure characteristics and FFT of a 50 mm baffled combustion chamber under nominal conditions. (a) Filtered (1 k–2 kHz) time history. (b) FFT waterfall of (a).
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addition, the pulse guns were triggered 0.6 s and 0.3 s before normal shutdown. Due to electrical cut-off in the second pulse gun, only the first pulse gun generated shock waves. The time history of the pressure fluctuations is shown in Fig. 7(a). Frequency analysis shows that there exists an unstable mode in the range of 1.2 k–1.6 kHz, which can be seen in Fig. 7(b). The phase difference of two dynamic pressures measured by FCC1 and FCC2 sensors is about 155◦ , and this mode was identified as the first tangential (1T) mode coupled with combustion. On the contrary, there was a time gap, 1.8 ms shown in Fig. 7(a) between the local pressure maximums in the two dynamic pressure measurements. This characteristic seems to result from a certain tangential flow or a flame oscillation in the combustion chamber. Similar behavior was observed in the single injector’s and sub-scale combustion test previously [13,14]. The flames exhibited oscillating lift-off shapes with asymmetric configurations, despite being under atmospheric conditions. By considering these characteristics, flames can be thought not to have stable lift-off heights and these phenomena may create pressure fluctuation characteristics as shown in Fig. 7(a). Low frequency pressure fluctuations of about 170 Hz resonance were also observed, shown in the waterfall of Fig. 7(b). These low frequency oscillations were observed throughout the test although unstable combustion responses attenuated and a stable operation was achieved with less than ±5% oscillation. Also, similar oscillations were monitored during the tests of a different combustor. This is assumed to be due to the coupling with the feeding system of our test facility. These unstable responses can be explained by combustion characteristics. One of the possible reasons is the hydrodynamic character of the adopted impinging injectors [13–15]. As in Fig. 8(a), the F-O-O-F impinging injector has two collision planes, at the axial distances about 30 mm and 50 mm from the faceplate which can be evaluated geometrically by considering the mass flow rates of kerosene and LOx. And the axial profile of the heat release rate can be shown schematically like Fig. 8(a) [13]. After injection, mixing, atomization and vaporization processes, the zone having the maximum reaction rate is assumed to exist near the second collision plane which is axially 40–50 mm away from the faceplate and plays an active role in interacting with acoustic modes. In this respect, a baffle of LB = 50 mm cannot completely separate the active flame zone from these calculated planes. Moreover, it is known that there is the longitudinalization of transversal acoustic modes in the baffle tip region [16,17], so a stiff pressure gradient can occur near the baffle tip, observed easily in Fig. 8(b), which shows the acoustic pressure contours of the first tangential mode. This can cause strong coupling between a stiff acoustic pressure gradient and an active combustion zone, resulting in a high susceptibility to a certain combustion instability phenomenon. In this regard, the 50 mm baffle cannot damp out pressure fluctuations sufficiently and pressure oscillations remained for a relatively longer duration than in the previous case when a baffle of LB = 85 mm was used.
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(a)
(b) Fig. 8. Hydrodynamic characteristics of a F-O-O-F impinging injector [13] and acoustic field illustrating a possible acoustic-flame coupling [16,17]. (a) Hydrodynamic characteristics of a present injector. (b) Acoustic pressure contours of the first tangential mode showing the longitudinalization near baffle tip.
Although pressure oscillations of high amplitude were observed during a certain time period, the decay time was about 6 ms shorter than generally recommended values of 15–25 ms [1,3,6,12]. We can now judge that the simple comparison of decay time is not a necessary and sufficient condition for combustion stability assessment. Instead, it is necessary to consider overall combustion characteristics. To further investigate these characteristics, an off-design condition prone to combustion instability was chosen. The possibility of the occurrence of combustion instability increases generally with pressure. According to the laminar flamelet model [15,18,19], flames near an extinction regime have a large amplification effect on acoustic pressure. In the case of a kerosene/LOx system, a low O/F ratio condition with respect to the nominal operating condition corresponds to a fuel-rich near-extinction regime. Therefore, a high pressure and low O/F ratio condition can be assumed to be the most prone to combustion instability. Although not mentioned herein, other off-
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design conditions were tested and confirmed to be relatively more stable than the above-mentioned off-design conditions. The combustion chamber of LB = 50 mm (LB /Dch ≈ 12%) under this condition (pch = 15.9 bar, O/F =2.00) has been tested with the same explosive charge masses and triggering times, and the results are shown in Fig. 9 only for the first external perturbation. Highly unstable pressure fluctuations were observed and sustained with a large amplitude for 60 ms, a relatively long time period compared with the previous results under nominal conditions. However, after 60 ms, the unstable oscillating pressure suddenly attenuated and disappeared, and returned to a stable condition. Responses to the second perturbation have shown qualitatively similar characteristics with those of the first one, having a relatively short decay time. In Fig. 9(b), the unstable pressure response characteristics occurred in about 170 Hz and 1.2 k–1.4 kHz frequency ranges. The former is thought to be due to the coupling with the feeding system, as explained earlier. The latter frequency range is
(a)
not exactly the same as in the previous test, the reason of which may be due to the dynamic characteristics of the impinging injectors adopted in this study. On the whole, impinging injectors are known to have relatively broad-band frequency characteristics [20]. Also, a temperature decrease due to a low O/F ratio could cause this frequency difference. According to the equilibrium analysis [21], the averaged sonic velocity under nominal condition is about 1256 m/s and that under low O/F ratio condition is about 1228 m/s. So the frequency shift by the decrease of sonic velocity can be estimated to be about 50 Hz and this characteristic would play considerable role in decrease of resonant frequency. Consequently, the combustion chamber with a baffle of LB = 50 mm is thought to be unstable. Although the combustion response using a 50 mm baffle is not fully unstable, it cannot be considered as stable. To support this tentative conclusion, a chamber of LB = 70 mm (LB /Dch ≈ 16.7%) was also tested under high pressure and a low O/F ratio. As in the case of LB = 85 mm, there were not any particular unstable modes observed, which can be seen in Fig. 10. Decay times were within 1.5 ms for both of the external perturbations. Therefore, both combustors, LB = 70 and 85 mm, are thought to have a large stability margin within the specified operating window. All test results are summarized in the operational working window, which are shown in Fig. 11. In the pre-specified working window, the present combustor has a large stability boundary with LB 70 mm. However, in the case of LB = 50 mm, this combustor exists in the vicinity of marginal stability under the nominal and off-design conditions. Characteristics of decay time, decay rate, An,max , and AR with respect to LB are shown in Fig. 12. For the cases of LB = 70 and 85 mm (LB /Dch ≈ 16.7, 20%), pressure fluctuations decayed within 1.6 ms and the amounts of decay rate were larger than 0.5, on the average. The decay time of LB = 85 mm is slightly larger than that of LB = 70 mm due to the different sizes of explosive charges used in the pulse gun device. Therefore, a clear quantitative comparison cannot be made. However,
(b) Fig. 9. Filtered dynamic pressure characteristics of a 50 mm baffled combustion chamber under off-design conditions. (a) Filtered (1 k–2 kHz) time history. (b) FFT waterfall of (a).
Fig. 10. Filtered (1 k–2 kHz) dynamic pressure characteristics of a 70 mm baffled combustion chamber under off-design conditions.
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(a)
Fig. 11. The operating conditions in stability rating tests.
combustion stabilization has been achieved and the one stable system having decay time of 1.25 ms (LB = 70 mm) is not always more stable than the other stable system having decay time of 2.5 ms (LB = 85 mm). They are thought to be all stable, irrespective of the amount of decay time and decay rate. Consequently, the present combustor under both cases is regarded as being quantitatively in the stable regime in the parameter space of decay time, decay rate and amplitude ratios. The test results showing stable combustion responses do not significantly deviate from each other regardless of the operating conditions such as the nominal and the off-design conditions. On the other hand, the results of LB = 50 mm (LB / Dch ≈ 12%) have long decay time, low decay rate and are scattered. Considering these characteristics and the data of the amplitude of pressure fluctuations An,max and AR, the combustor of LB = 50 mm is thought to be in the unstable regime in terms of practical application. In addition, An,max and AR can reflect the responses of a combustion zone to external perturbations and are much more dependent on the characteristics in combustion chambers than on the size of explosive charge used in the pulse gun [8,9,11, 14]. In other words, An,max and AR would be large if combustion characteristics were very sensitive to external perturbations. Therefore, large values of An,max and AR show that a combustion chamber can be moved to an unstable regime by any undesired changes to the operating conditions. The case of LB = 50 mm, which shows unstable characteristics, reveals considerably large and scattered values compared with the LB = 70 and 85 mm test cases. As mentioned before, the effect of an axial baffle length on the thrust or specific impulse of the present combustor has been analyzed and confirmed under unperturbed conditions to validate the optimizing process of an axial baffle length. Specific impulse, Isp with respect to length is shown in Fig. 13. Intuitively, the specific impulse of the present combustor decreases with axial baffle length. Under nominal conditions, the decrease of specific impulse is relatively steep, compared with off-design conditions. In other words, a long-length baffle for stability assurance may cause considerable performance loss
(b)
(c) Fig. 12. Decay time, decay rate and the amplitude ratios with respect to three different axial baffle lengths, LB . (a) Decay time. (b) Decay rate. (c) An,max and AR.
for the present combustor. Therefore, this length-optimization process would be very important for the determination step of design confirmation. Consequently, the stability boundary of the combustion chamber in this study is assumed to be between 50 and 70 mm in axial baffle length, LB . Quantitative optimization with only a few experimental results would be insufficient and difficult; however, by synthesizing the response characteristics of a decay time and two amplitude ratios practically, the actual stability boundary exists close to LB = 50 mm.
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Fig. 13. The specific impulse, Isp with respect to axial baffle lengths.
4. Conclusion Stability rating tests on a liquid propellant rocket combustor were carried out for the evaluation of stability capacity with respect to axial baffle length. A pulse gun device capable of generating two external perturbations was adopted to determine the dynamic stability boundary of an axial baffle length. Generally, the axial baffle length with respect to the chamber diameter is recommended to be about 20%. However, it is necessary to decrease the axial length to minimize the thrust loss and cooling problems caused by installing a baffle in a combustion chamber. A decay time and two amplitude ratios of a steady-state to the induced external perturbation have been used to elucidate the stability boundary of the combustion chamber. In the case of LB = 50 mm (LB /Dch ≈ 12%), unstable combustion responses caused by external perturbations were observed. These characteristics can be explained by comparing the axial baffle lengths and the axial distances of the second geometrical collision planes of the F-O-O-F impinging injectors in the viewpoint of heat release rate. On the other hand, pressure fluctuations of large amplitude were damped out within the recommended reference decay time in the cases of larger baffle lengths, showing satisfactory stabilization capacity under a specified operating window. Therefore, the stability boundary with respect to axial baffle length is assumed to exist around 50 mm in the present combustor. Finally, the optimization process of an axial baffle length and the indigenous establishment of stability rating test procedures have been set up. Acknowledgements This work has been financially supported by Korean Ministry of Science and Technology. The authors would like to express sincere thanks to the efforts of all members in Combustion Chamber Department at Korea Aerospace Research Institute for conducting these hot firing tests. References [1] D.T. Harrje, F.H. Reardon, (Eds.), Liquid Propellant Rocket Combustion Instability, NASA SP-194, 1972.
[2] E. Laudien, R. Pongratz, R. Pierro, D. Preclik, Experimental procedures aiding the design of acoustic cavities, in: V. Yang, W.E. Anderson (Eds.), Liquid Rocket Engine Combustion Instability, in: Progress in Astronautics and Aeronautics, vol. 169, AIAA, Washington, DC, 1995, pp. 377–399. [3] A.F. Agarkov, K.P. Denisov, M.L. Dranovsky, I.A. Zavorohin, V.N. Ivanov, V.P. Pikalov, A.A. Shibanov, Injector flame stabilization effects on combustion instability, in: V. Yang, W.E. Anderson (Eds.), Liquid Rocket Engine Combustion Instability, in: Progress in Astronautics and Aeronautics, vol. 169, AIAA, Washington, DC, 1995, pp. 281–305. [4] NASA, Liquid Rocket Engine Combustion Stabilization Devices, NASA SP-8113, 1974. [5] J.J. Fang, Longitudinal mode combustion instabilities of a high-pressure fuel-rich LOX/RP-1 preburner, in: Proceedings of the 22nd JANNAF Combustion Meeting, vol. 1, 1985, pp. 429–441. [6] J.C. Oefelein, V. Yang, Comprehensive review of liquid-propellant combustion instabilities in F-1 engines, Journal of Propulsion and Power 9 (5) (1993) 657–677. [7] C.H. Sohn, W.S. Seol, S.Y. Lee, Y.M. Kim, D.S. Lee, Application of combustion stabilization devices to liquid rocket engine, Journal of the Korean Society for Aeronautical and Space Sciences 31 (6) (2003) 79–87. [8] Y.M. Han, Y.S. Ko, S.H. Kim, K.J. Lee, Y.G. Chung, W.S. Seol, S.Y. Lee, Combustion stability rating test using a pulse gun device in a liquid rocket engine, in: The 4th Symposium on Space Launch Technology, May, 2003, pp. 33–37. [9] K.J. Lee, H.J. Kim, S. Seo, I.Y. Moon, H.M. Kim, S.Y. Lee, Y.S. Ko, Experiments on development of a pulse-gun pressure-wave inducing device for stability rating test technique, Journal of the Korean Society for Aeronautical and Space Sciences 33 (3) (2005) 99–106. [10] C.H. Sohn, S.K. Kim, Y.W. Moon, Y.M. Kim, Technologies for Suppression/Protection of Combustion Instabilities in Rocket Engines for Launch Vehicle, Ministry of Science and Technology, Korea, 2003. [11] Denisov, et al., KARI combustion chamber firing tests, Report under Contract No. KARI-01-032, Phase 4, NIICHIMMASH, 2002. [12] M.D. Klem, R.S. Fry, Guidelines for combustion stability specifications and verification procedures for liquid propellant rocket engines, CPIA Publication 655, 1997. [13] C.H. Sohn, W.S. Seol, A.A. Shibanov, V.P. Pikalov, Combustion stability boundaries of the subscale rocket chamber with split-triplet impinging jet injectors, Journal of Propulsion and Power 23 (1) (2007) 131–139. [14] S. Seo, K.J. Lee, Experimental study of combustion stability assessment of injector, Journal of the Korean Society for Aeronautical and Space Sciences 32 (4) (2004) 61–66. [15] J.S. Kim, F.A. Williams, Acoustic-instability boundaries in liquidpropellant rockets – theoretical explanation of empirical correlation, Journal of Propulsion and Power 12 (3) (1996) 621–624. [16] J.M. Wicker, M.W. Yoon, V. Yang, Linear and non-linear pressure oscillations in baffled combustion chambers, Journal of Sound and Vibration 184 (1995) 141–171. [17] S.-K. Kim, H.J. Kim, W.S. Seol, C.H. Sohn, Acoustic stability analysis of liquid propellant rocket combustion chambers, AIAA Paper 2004-4142, 2004. [18] C.H. Sohn, S.H. Chung, J.S. Kim, F.A. Williams, Acoustic response of droplet flames to pressure oscillations, AIAA Journal 34 (9) (1996) 1847– 1854. [19] H.J. Kim, C.H. Sohn, S.H. Chung, J.S. Kim, Nonlinear acoustic-pressure responses of oxygen droplet flames burning in gaseous hydrogen, Journal of Mechanical Science and Technology 15 (4) (2001) 510–521 (previously, KSME Int’l Journal). [20] Y.M. Han, S.H. Kim, I.Y. Moon, H.J. Kim, J.G. Kim, W.S. Seol, S.Y. Lee, S.T. Kwon, C.J. Lee, Combustion performance tests of fuel-rich gas generator for liquid rocket engine using an impinging injector, Journal of the Korean Society of Propulsion Engineers 8 (2) (2004) 10–17. [21] S.-K. Kim, Y.W. Moon, T.S. Park, Development of chemical equilibrium cfd code for performance prediction and optimum design of LRE thrust chamber, Journal of the Korean Society of Propulsion Engineers 9 (1) (2005) 1–8.