Combustion, Explosion, and Shock Waves, Vol. 46, No. 4, pp.

408417, 2010

Problems of Implementing Ramjet Operation

V. M. Levin1
UDC 629.735.33.016

Translated from Fizika Goreniya i Vzryva, Vol. 46, No. 4, pp. 4555, JulyAugust, 2010. Original article submitted August 31, 2009; revision submitted December 1, 2009.

This paper discusses an approach to implementing the operation and designing the ow duct of a ramjet combustor. Methods of air compression, ignition, ame stabilization, and fuel combustion in the ow are considered for the purpose of implementing an eective process on a short length with moderate total pressure losses. Advantages of narrow-mode ramjet engines are noted. Comparative results of experiments on the burning of hydrocarbon fuels in the ramjet combustors are given. Key words: ramjet engine, combustion chamber, aviation kerosene, operation.

INTRODUCTION The article published in 1987 [1] considered perspectives and gave recommendations for investigations of liquid fuel scramjets. Despite the fact that the article was published an impressively long time ago, many problems stated in it have not been investigated in sufcient detail due to their complexity. Nevertheless, in comparison with 1987, the situation has changed significantly. Serious advances in experimental investigations [2, 3], along with the sum of the results of mathematical modeling [4, 5], have led to considerable progress in this eld of study. Certainty about success has been supported by the results of the summer tests of the Kholod hypersonic ying laboratory of Central Institute of Aviation Motors (CIAM) which proved the feasibility of operation of the combustion chamber of a dual-mode ramjet engine in the atmosphere [6].

DEVELOPMENT OF A RAMJET ENGINE WITH FIXED FLOW DUCT AREA The thrust-economic characteristics of a ramjet engine with xed ow-duct geometry over a wide range of external parameters are noticeably inferior to the similar parameters of an adjustable engine. Regardless of that, the design of a xed ramjet engine with high thrust-economic characteristics for tactical aircraft cur1

Moscow Aviation Institute (State University of Aerospace Technologies), Moscow 125993; vadimlevin@yahoo.com.

rently seems to be the closest solution to the problem of creating an engine of a new type and, at the same time, it is the main obstacle to moving to the next level of study. The choice of the shape of the ow duct of a xed ramjet engine, including the air inlet, diuser, combustion chamber, and nozzle is dictated by the thrust characteristic of the engine in some computational domain of the designed range of its operation determined by its purpose. In this range of parameters, the mode of maximal economy and maximal specic impulse occurs. On the range boundaries, the thrust-economic characteristic worsens; therefore, the choice of the design point should be well grounded to provide the eciency of the engine over the whole range of its operating conditions. The problem of designing a xed engine which is able to provide a high specic impulse over a wide range of speeds and which is called a dual-mode engine involves designing a chamber which is eective over a wide range of parameters of its component parts. The problem of designing a hydrocarbon ramjet combustor is particularly dicult. The following important stages of solution of separate aspects of this problem can be singed out. 1. Moscow Aviation Institute (MAI), 1989. Development and testing of a narrow-mode high-performance short combustion chamber of a scramjet which runs on kerosene under conditions which model ight with a Mach number Mf = 6. Fuel barbotage, a pneumatic throttle valve for forced ignition, and micropylons were used for the rst time [79].

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0010-5082/10/4604-0408 c 2010 Springer Science + Business Media, Inc.

Problems of Implementing Ramjet Operation 2. CIAM, 2001. Development and testing of the combustion chamber of a dual-mode ramjet engine which runs on gaseous methane under conditions which model ight with Mf = 34. Detailed methodological studies of the impact of methods of fuel supply, stabilization, and fuel temperature on the eciency of the process are given in [10]. The essence of the problem of developing a combustion chamber of a dual-mode ramjet engine lies in the synthesis of the best method for controlling the operation process. There are two fundamentally dierent versions of implementing the process in the chamber that provide very high performance over a wide range of inlet ow parameter. In the two-chamber ramjet engine model [11], the combustion process is implemented in two parallel combustion chambers, one of which is subsonic. This camera operates with an excess of fuel and is a gas generator which supports stable fuel burning process in supersonic concurrent air ow in the other chamber. The two-chamber scheme provides reliable stabilization of the jet over the entire range of operating modes. The region of use of this engine is up to Mf = 6.5. The prospect of the scheme may be associated with the development of a combined rocket-ramjet engine. In the second model, combustion is implemented in subsonic and supersonic ows in a chamber composed of two series-connected sections (two-section chamber) (shown schematically in Fig. 1). The choice of the shape of the ow duct adjacent to the insulator (upper chamber) corresponds to the optimal chamber geometry for supersonic combustion if Mf > 6 (scramjet mode). The geometrical parameters of a subsonic (lower) combustion chamber operating at Mf 6 (scramjet engine mode) are determined in accordance with the minimum ight speed. In both models under certain operation conditions, the necessary degree of compression and deceleration of the air ow in the diuser is provided (or sustained) by heat supply. It appears that the two-chambered model is more dicult to implement technically and it has no signicant advantages over the second model in both gas-dynamic and engineering characteristics (due to a number of problems related to thermal protection, complexity of fuel supply, and increasing ow duct length and weight). However, for the two-section model, a very serious problem is ame stabilization during subsonic fuel combustion under a great change in the consumption of the components. Nevertheless, this model is preferred by most researchers. Let us consider the key elements of the ow duct of the two-section combustion chamber of a ramjet engine. Insulator

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In a dual-mode ramjet engine there is no diuser in its classical design as a diuser of a scramjet. Protection of the ow in the are inlet throat from the backpressure impact in the chamber, i.e., the protection of the engine from surge, is partly performed by the device which a short prediuser (insulator) with a small expansion angle (13 ). This section is required to attach the boundary layer separated in the inlet throat because of shock interaction typical of wide-range xed inlets and to level out the velocity prole. The length of the insulator depends on many parameters. In particular, if a bow shock of the deceleration zone is to be places therein, the length of the insulator can depend on the shape of the inlet section. According to [12], the smaller the relative width of the cross section of a rectangular duct or the larger the radius of corner rounding in the cross section, the greater the pressure gradient during ow deceleration in a pseudo-shock. Reducing the thickness of the boundary layer in the corner zones or producing internal or external kinks in the duct improves the conditions for xation of a bow shock, which makes it possible to reduce the insulator length. Method of Fuel Supply in a Dual-Mode Ramjet Engine In the presence of supersonic ow at the inlet of a combustion chamber, the time of operation is limited to microseconds. The maximal time the fuel can be present in the chamber is provided by fuel supply at the insulator outlet. That is why, all the fuel is supplied here over the whole range of inlet air parameters and the air-tofuel ratio. Fuel injection is performed with thin pylons along the normal to the vector of the concurrent ow. Despite some disadvantages of this method compared to concurrent supply (pressure losses increase due to the interaction of the jets with the ow, and the impulse of the fuel jets is not used), there are also notable advantages for the implementation of the operation process. Normal fuel injection increases the time period during which the fuel is in the ow, provides deeper penetration of drops, and the best parameters of dispersion and evaporation, i.e., better preparation of the mixture on a shorter length, as well as a reduction in the distance to the ignition zone in comparison with concurrent supply. Fuel barbotage i.e., fuel saturation with small gas bubbles, can be considered useful from the point of view of fuel ignition in a combustion chamber. The ow of such jets follows the gas laws, which simplies the problem of automatic control of consumption at supercritical

410 gradients. It should be noted that the interaction between a liquid with gas bubbles and supersonic concurrent ow has interesting properties. At the nozzle exit, a liquid with a large number of dissolved (or mixed) gas bubbles discharges into a medium at a greatly reduced pressure. The released potential energy of the compressed gas breaks the gas-liquid phase into small droplets with an eect of explosion, which leave the injection zone. In comparison with other methods of normal injection of liquid into a supersonic ow, this method, as shown by a experiments, provides the most extensive area of spraying and the maximum depth of penetration with dispersion comparable to the dispersion produced by using in a centrifugal nozzle [7, 8]. Furthermore, in this method of supply, the pressure losses are minimal. Apparently, this is due to the fact that the pressure losses occurring when each individual small drop intersects the supersonic concurrent ow with high kinetic energy are negligible. The individuality and disconnection of the drops contribute to rapid evaporation. The kerosene jet saturated with bubble hydrogen (mass ratio 0.1%, supplied under a pressure of 25 atm along the normal from a 0.3 mm diameter nozzle into a supersonic air stream owing out of a air heater with parameters Min = 2.5 and Tin = 1650 K) is quickly rotates and evaporates at a distance of 2025 mm. It should be noted that the interaction of droplets with shock waves also leads to their further fragmentation.

Levin eration apparently occur in the case where, in a wide range of chamber throttling, a pseudo-shock is xed in the interference zone generated by injection array. It can block this area, but its further movement upstream is undesired because of the possible ow separation at the air inlet. In the outlet zone of the upper chamber there is a pylon a ame stabilizer used to implement combustion in the scramjet mode. It should be noted that with an appropriate choice of the stabilizer shape, it is possible to may correct the law of area variation of the upper chamber by length. This is particularly necessary in two cases: 1) for eective operation in the scramjet mode; 2) for placement of a pseudo-shock in the scramjet mode (it is known that during ow deceleration in divergent ducts, the length of the pseudo-shock increases [13] and sound speed is attained only in the presence of an area of constant cross section in the duct outlet). Another function of the inlet section is a combustion chamber (upper chamber), which runs in the scramjet mode under conditions that correspond to Mf 6. Here at Min = 2.52.8, the deceleration temperature exceeds 1650 K. With an increase in the thermodynamic parameters, the length of the mixture preparation path decreases dramatically and heat release occurs is realized in the upper chamber. The characteristic of the process can be judged by the results of successful bench tests performed at MAI with a supersonic combustion chamber at Mf = 6 using kerosene [7]. During those experiments, kerosene was for the rst time delivered into a duct zone with a strong initial expansion, which ensured a decrease in the thermodynamic parameters of the ow. First, this reduces the risk of blocking the chamber at air-to fuelratio () close to unity. Second, this makes it possible to reduce the burning rate in the initial section, i.e., as a matter of fact, to change the heat release to t the mode p = const, where heat supply occurs at a constant pressure in the middle and lower zones of the chamber. Third, this increases the pressure at the nozzle inlet and, hence, the integral of the jet pressure on the nozzle walls. This method made it possible to design a short combustion chamber which runs on kerosene and provides high eciency at 1. For an inlet height of 50 mm (width 100 mm) and ow inlet parameters are Min = 2.5 and Tin = 1600 K, the experimental combustion chamber was 375 mm long. The bow shock-wave preceding the heat release area is xed in the injection area, but the level of total losses is negligible. From the result of a re experiment for = 13, the total pressure recovery factor at the chamber exit obtained from measurement of the deceleration pressure was 0.3 for

Upper Section The insulator is adjacent to the main section of the engine combustion chamber, which, depending on the ramjet speed range, can be made with an extension up to 7 (or more) in the primary zone. The functions of a diuser and combustion chamber are combined here. Let us consider one of the two major functions of this section. This is a diuser which provides placement of the deceleration zone of air ow to subsonic speed in front of the ame stabilizers and which precedes the combustion zone for engine operation in a scramjet engine mode, i.e., when Mf 6. As in a traditional scramjet, the closer the bow wave of a pseudoshock to the prediuser, the lower the pressure losses are in the deceleration zone and the higher the ramjet performance. The diuser expansion provides a negative pressure gradient in the duct precisely where the placement of the pseudo-shock compensates it by its strong positive gradient. This leads to an important eect a reduction in the displacement of the pseudo-shock, i.e., positioning (xation) of its bow wave in a certain range of backpressure. The most favorable conditions for op-

Problems of Implementing Ramjet Operation completeness of combustion = 0.940.96. In the specied range of conditions, the chamber worked steadily. The results of the chamber tests and three-dimensional calculations show that the specied shaping method as a way of achieving high characteristics of a scramjet chamber is one of the best. The dierence between the upper chamber of a dual-mode ramjet engine and the MAI chamber described above is the presence of ame stabilizers placed in it to implement the operation process of the lower chamber. Lower Combustion Chamber. Flame Stabilization in the Scramjet Mode Implementing stable fuel combustion in subsonic ow is known to be related to the use of niches or highdrag bodies, such as wall ledges, pylons, etc. In a twosection model, the tool of stabilization, in particular the case (pylon) of the stabilizer, is placed in the upper chamber so that its cross section is located in the plane of the lower chamber inlet. One method of decreasing the length of the lower chamber is the use of pylons with a developed (nonrectilinear) shape of the rear edge, which helps increase the perimeter of the ignition line and shorten the linear size of the ame front. In view of the aforesaid, the length of the lower chamber can be signicantly reduced without lose of its eciency. In [10], a study was made of the eect of the concentration of an air-fuel mixture in the wake zone behind pylons and wall stabilizers on the stability limits of methane burning in a combustion chamber. It can be suggested that transition of the stabilization zone from the wall into the ow core, i.e. into the zone of high fuel concentration in the free mixing layer, contributes to the extension of the stable operation range of the combustion chamber. The placement of the pilot ame zone into the ame core also has a positive eect on the thermal protection of the walls of the combustion chamber.

411 example, by normal high-pressure injection of any gas through a series of holes in the walls of the lower part of the camera). Short-time throttling (0.5 sec) leads to a rapid decrease in the ow speed and an increase in the density in both the zone of ow over stabilizers and downstream to the point of pneumatic impulse. If there is an active source of temperature increase in the zone of reverse ows, this time may be sucient for the occurrence of favorable conditions for mixture ignition and stable diusion burning after the shutdown of the injection system. This method is quite eective, reliable and in a well-shaped duct of a dual-mode ramjet engine with = 1.52.5, it provides stable ring of the combustion chamber without blocking at almost any ow temperature. This method was recognized after its rst application for kerosene ignition and combustion in a short scramjet chamber [6], where the impulse point was at a distance of 0.6 of the chamber length from the inlet cross section. Operation of a Ramjet Engine In a chamber of xed geometry, the system shock (deceleration zone) + combustion (heat release zone) takes its position in the duct in accordance with the specied shape of the ow duct and the activity of chemical processes. For ow parameters corresponding to Mf = 36 (scramjet mode), due to low ow temperature, even with good mixture formation, the chemical processes in the upper chamber go without heat release. The burning zone is in the lower chamber behind the mechanical ame stabilizers. If the lower chamber is long enough, the process of heat release is almost completed here, and the subsonic ow is accelerated to the speed of sound in the critical section. Under chamber inlet conditions corresponding to Mf 6 (Tin > 1650 K), the chemical processes are dramatically accelerated and the heat release process is moved upstream. Activation of throttling increases the intensity of the bow wave which is xed at the injector array with the formation of Mach intersections in the ow core. Mach disks become a mean of stabilization in the ow core of the subsonic combustion areas. If a bow shock structure consisting of one or two strong interactions is xed, a system of concurrent (subsonic and supersonic) ows arises. As in the case of stabilization in a scramjet, where burning is sustained by the pilot ame in the wake zone, the fuel burning in the free subsonic zone becomes a means of ignition and stabilization of combustion of the mixture in the outer supersonic eld, in the system of oblique shock waves. As shown by computational and experimental results

Means of Forced Ignition When Mf 6, the thermodynamic parameters of the air ow in the combustion chamber do not provide conditions for ignition of the mixture. Because of the speed of the air-fuel mixture in ow over the stabilizers, its ignition is impossible in the zone of reverse ows even when using means of forced increase of temperature. A very eective method for attaining reliable ring of a chamber in experimental studies is pneumatic throttling of the ow in the combustion chamber (for

412 [2, 4], the heat supply in these parallel ows on average leads to the attainment of the speed of sound at the nozzle throat section. With this ow scheme, one can obtain a chamber operating at 1 with completeness of combustion = 0.940.96 and a total pressure recovery factor = 0.3. This combustion chamber operation mode occurs with a carefully chosen shape of the ow duct, and it can be considered an optimal operation regime in terms of heat release rate and eciency. Until a certain point, supersonic ow is sustained around the zone of subsonic ow. As the throttling increases, so do the lateral dimensions of the subsonic zone of stabilization. When the degree of thermal throttling approaches the limiting value, the subsonic ow covers the entire cross section of the upper chamber. The pressure wave in the duct overlaps the injection zone and penetrates into the pre-diuser. The completeness of combustion approaches unity. In the limit, with increasing fuel consumption above the critical point (for the chamber of this geometry), the ow regime is characterized by the destruction of the supersonic ow structure in the air inlet throat, the occurrence of a knockedout shock wave, and a catastrophic decrease in air consumption by the engine. The value of () decreases sharply to 0.050.1. Thus, for an aircraft speed corresponding to Mf 6, the operation process can be completed in the upper chamber at 1 with very high eciency. In this case, at its narrowest outlet section (cross section of mechanical stabilizers), speed of sound is attained i.e., the critical section of the combustion chamber moves into the ow duct. Because the cross-sectional area of the outlet of the lower chamber far exceeds this section, the gas ow in it is accelerated is serves as part of the engine nozzle. However, since gas acceleration in the lower chamber occurs not in the shaped duct, this leads to a serious loss of pressure and thrust. In fact, the lower chamber becomes a ballast. The total pressure recovery factor in a dual-mode ramjet engine reduces from 0.3 (for a well-shaped combustion chamber) to 0.130.15. According to the aforesaid, in a dual-mode ramjet engine with a xed shape of the ow duct in the specied range of inlet parameters, the critical gas ow velocity is reached at known sections of the ow duct. In the rst case, it is reached in the subsonic combustion mode in the exit nozzle throat, and in the second case, at the upper chamber outlet in the scramjet operation mode inside the ow duct. In the scramjet mode, apparently, some intermediate ow scheme is realized with an acceleration up to the speed of sound, i.e., with a oating critical section. The aforesaid to several important conclusions. The engine design should provide satisfactory thrust-

Levin economic characteristics in both subsonic and supersonic combustion modes. At the same time, the upper chamber of a dual-mode xed ramjet engine is shaped to a supersonic combustion mode in the hypersonic speed range of aircraft ight, leading to a loss of the performance in the scramjet mode, in particular, due to high ow velocity. At high velocity of the airow over the central stabilizers, the eciency of ame stabilization reduces, resulting in a reduction in the combustion stability and eciency over a wide range of (). We can say that in a dual-mode ramjet chamber in the lower speed range (subsonic combustion mode), it is impossible to attain performance similar to that in a pure scramjet. As the ight altitude increases, if Mf > 6, the characteristics of the combustion chamber of a xed scramjet (as well as the upper chamber of a dual-mode ramjet engine) vary in the same way as the properties of the process in the scramjet mode. That is, to maintain the pressure and the eciency of heat release in the combustion chamber, it is required to decrease its outlet section and to slightly increase its length. Therefore, if a combustion chamber is shaped for a mode Mf = 6, then, an increase in the ight speed will lead to a reduction in the operation eciency, for example, from mixed complex ow (subsonic and supersonic) and a high-performance process at Mf = 6 to inecient pure supersonic combustion at Mf = 7, and will be even lower at Mf = 8. The above consideration of the eciency of a dualmode ramjet engine with xed geometry of the ow duct leads to the following conclusions: the performance of a dual-mode ramjet engine operating in both subsonic and supersonic combustion will always be lower than the thrust-economic performance of supersonic and hypersonic ramjets operating in the range of ight speeds of the specied dual-mode ramjet engine. the narrower the specied range of Mf a ramjet engine, the higher its thrust-economic performance. The currently developed foreign samples of liquid fuel ramjet engines are designed mainly for single-mode limited speed ranges Mf = 34.5 and 46 (supersonic ramjet), Mf = 5.57.5, 5.58, and 78 (hypersonic ramjet). RESULTS OF INVESTIGATION OF THE OPERATION OF A RAMJET COMBUSTOR The formulated idea of the operation of a ramjet combustor and the above-listed methods for aecting its performance were implemented in bench studies of an experimental hydrocarbon combustor with an at-

Problems of Implementing Ramjet Operation

413

Fig. 1. Schematic circuit of a combustion chamber which runs on kerosene and the distribution of the static pressure along the length of the ow duct.

tached air duct (Fig. 1). The combustion chamber is two-sectioned, uncooled, and made of a powder alloy based on FeCrAl. The air ow velocity at the chamber inlet was modeled by three replaceable shaped nozzles: Min = 1.4, 2.1, and 2.9. The main fuel was cool aviation kerosene with a relative mass composition of 13.4% hydrogen and 86.6% carbon, a stoichiometric air coefcient of 14.75, and a caloric value of 42,914.7 kJ/kg. Kerosene was saturated with air to a relative mass content of 0.1%. Under investigation conditions corresponding to ight at Mf = 3, the proportion of free oxygen in the gas coming from the heater in the chamber combustion was 18%, and in all other experiments, it was 23%. The chamber of rectangular cross section 0.78 m long and an 30 70 mm inlet had a constant width of the ow duct. Four injector three-duct pylons (each provided with 40 jet nozzles) 1.8 mm thick were mounted in a transverse line at the inlet section and provided 8.8% blockage of the section. Barbotaged kerosene (K) was supplied through them to the combustion chamber at ow rates of 30300 g/sec.

In the basic design version, the following scheme of ame stabilization was investigated. At the end of the upper chamber of length 0.423 m on the upper and lower walls, two shifted wedges of rectangular cross section 16 mm thick were mounted. In the ow duct, an additional critical section was arranged in the middle of the lower chamber by mounting four additional stabilizers on its sidewalls. Under the assumption of a linear change in the size of the conventional critical oating section, the size of the additional stabilizers was chosen so that the open ow area of the duct in the plane of their edge was equal to the half-sum of the abovementioned critical sections. The sidewalls in the lower chamber had holes for air injection in order to implement forced fuel ignition by means of pneumatic throttling (Th). For the same purpose, the re block (FB) supplies hot gas jets to the bottom area behind the central stabilizers. This zone was simultaneously supplied with hydrogen (H2 ). The above-listed means for implementing the start were used only to ignite the main fuel.

414 TABLE 1
Mf Min Tin , K 740 730 704 3 1.4 726 730 715 983 980 968 4 2.1 965 928 920 1129 1141 1156 5 2.1 1173 1183 1112 1735 6 2.9 1711 pin , 105 Pa 14.7 14.48 14.37 14.29 14.5 14.4 23.65 23.71 23.33 23.31 23.18 23.1 14.06 14.13 14.06 14.21 14.45 14.45 24.15 23.93 1.81 1.45 1.43 1.53 1.1 1.2 2.55 2.48 1.56 1.57 1.15 1.19 1.26 1.11 1.58 1.44 1.26 1.35 1.21 1.1 hcr , mm 70.4 70.4 70.4 70.4 70.4 70.4 65 65 67.7 67.7 68.5 68.5 70 70 59 58 57 63.1 70.5 70.5 Fire block + + + + + + + + + + + + + + + + + + Throttling + + + + + + + + + + + + + + + + + + Supply of H2 + + + + + + + + + + + + + + + 0.814 0.822 0.834 0.74 0.8 0.8 0.887 0.95 0.972 0.985 0.773 0.826 0.988 0.809 1 1 1 1 0.98 0.87 0.63 0.68 0.703 0.65 0.63 0.64 0.31 0.31 0.357 0.36 0.356 0.36 0.356 0.35 0.373 0.391 0.41 0.37 0.138 0.135

Levin

The design provided an opportunity of a discrete change of the critical cross section of the chamber nozzle (vertically) within hcr = 5775 mm. During the experiment, we measured all the basic parameters of the supplied components and the longitudinal static pressure in the combustion chamber. The ame of the combustion products was recorded on the outlet of the combustion chamber by two video cameras. The investigation conditions and the main results are given in Table 1. The above-mentioned measures provided reliable ignition, stabilization and fuel burning throughout the investigation. It should be noted that in some of the experiments with a nozzle of Min = 2.1, ignition occurred without throttling. In the tests with a nozzle of Min = 2.9, the means promoting ignition were not used, self-ignition of the fuel occurred in the upper chamber. In some experiments after ignition, we stopped the supply of barbotaged gas, and that did not lead to combustion quenching. The distribution of the static pressure along the combustion chamber p = pw /ph (pw is the current value of the static pressure and ph is the pressure at the outlet

of the combustion chamber) is given in Fig. 1. In experiments at Min = 1.4 and = 1.21.8 and Min = 2.9, ow modes were implemented without blocking of the chamber. In some experiments in scramjet modes with a nozzle of Min = 2.1 the burning process was accompanied by disturbance of the ow at the duct inlet, resulting in a shift of the beginning of static pressure increase to the heater nozzle. The length and shape of the upper chamber were chosen based on the research experience of the burning process at Min 6, and so that a pseudo-shock is located here in scramjet modes. However, blocking occurred in almost all modes of the chamber when > 1.2. Judging by the results, the selected length of the upper chamber proved insucient for placement of the deceleration zone (pseudo-shock), which led to ow disturbance in the inlet section. Installation of the lowers stabilizers naturally impacted the longitudinal distribution of the static pressure in the duct during its cold mode operation (see Fig. 1c). Figure 2 gives the results of processing of the tests performed using a one-dimensional procedure taking into account the proportion of oxygen in the gas

Problems of Implementing Ramjet Operation

415 insucient cross-sectional area of the secondary camera in the heat supply zone (decit is about 10%). We can assume that in this series, the operation process is exposed to a negative eect of the lower stabilizers. The narrowing of the ow duct led to ow acceleration in the combustion zone. As stated above, the purpose of the lower stabilizers is to improve the intermediate characteristics of the scramjet (if Mf 4.5) ow disturbance at the inlet of the combustion chamber and, consequently, worsening of the fuel spray parameters (explained below). It is of interest to compare the results on kerosene combustion with the results of studies of methane combustion in the combustion chamber of a dual-mode ramjet engine at CIAM [10] (Fig. 3). Kerosene and methane are hydrocarbon fuels with very similar caloric values. As shown by experimental studies, their use in ramjet engines involves the problem of mixture preparation and implementation of ignition. At CIAM, they studied the combustion of gaseous methane heated to 550880 K in a two-section combustion chamber under conditions modeling a Mach number Mf = 34, i.e., at an air temperature at the combustor inlet of 650910 K. The investigations were performed on an attached duct with a nozzle at the inlet Min = 2.0. The length of the rectangular combustion chamber was 1 m, and the inlet size was 40100 mm. At a constant width (100 mm), the height of the chamber was variable: the upper chamber had a smooth divergence to a height of 50 mm, then followed two symmetrical ledges, each 15 mm (combustion stabilizers), and the lower chamber was a duct of constant cross section 80 mm high, i.e., without a critical cross section. We should mention the high quality of the methodology of performing this experimental work. In one of the directions of the study, the main fuel was supplied normal to the ow at the beginning of the upper section through 8 holes on the upper and the lower walls in the wake past acute-angled pylons (M). In addition, methane in a small amount (up to 12 g/sec) was supplied to the zone behind the ledges. Three inclined pylons installed in front of the ledges on the top and bottom walls at the end of the upper chamber in were used as additional ame stabilizers. The static pressures distribution in the duct during methane combustion for various values of () is given in Fig. 3. The results of the studies of combustion of kerosene and methane in both chambers demonstrate a high level of increase in the static pressure. For > 1.2, of thermal choking of the combustion chamber was observed the experiments at MAI and CIAM. This may be responsible for a decrease in the completeness of combustion in this range of air-to-fuel ratio. As noted above,

Fig. 2. Calculated completeness of kerosene combustion at the outlet of the combustion chamber versus test conditions.

Fig. 3. Schematic of the combustion chamber which runs on heated methane and static pressure distribution along the duct length.

supplied to the chamber. It is seen that under the conditions Min = 2.1 and Tin 920 and 1120 K, the completeness of combustion reaches values = 0.80.99, and at Min = 2.9 and Tin 1711 and 1735 K, it is = 0.870.98. The relatively low completeness of kerosene combustion ( 0.8) at Min = 1.4 and Tin 720 K may be explained by a number of factors: low activity of the operation process associated with a deciency of oxygen in the heater gas in that series of experiments,

416

Levin In summary, the tests of a short ramjet chamber at MAI demonstrated the following: The shape of the air-gas part, including the means of stabilization, provides stable operation in the range of 12.5; The combustion eciency is very high ( = 0.81.0) in the range of 1.22.5. For all modes in the range of external conditions for > 1.2, the completeness of combustion reduces, which conrms the data of [10]; The pressure recovery factor varies in a satisfactory range: from 0.7 (at Mf = 3) to 0.134 (at Mf = 6); In this combustion chamber, the size of the critical section hcr (see Table 1) has no signicant impact on the ignition characteristics of the fuel; The range of means used to ignite kerosene is suciently eective; It is conrmed that the type of gas used for barbotage has no noticeable eect on the characteristics of the process, and stopping the barbotage of fuel does not aect the sustainability of its combustion in the tested range of operating parameters; The expediency of using secondary stabilizers should be explored. One of the possible ways to optimize the design of the air-ow duct of a short ramjet chamber is a three-dimensional calculation of the duct ow.

Fig. 4. Dependence of the total pressure recovery factor in a combustion chamber on testing conditions.

the best mixing conditions occur when the fuel sent to supersonic ow crosses the system of shock waves, where there is a further reduction of droplet sizes. In the case of blocking of the combustion chamber, where the bow shock wave of the deceleration zone is upstream of the injector location, the fuel is supplied to low-speed ow, which leads to an increase in the droplet size, deceleration of evaporation, and a reduction in completeness of combustion. On the other hand, regardless of the ow structure in the inlet section, the ow characteristics in the zone of airow over the ame stabilizers (gas ow rate and temperature) and the stabilization conditions remain almost unchanged. Figure 2 gives a curve of the limiting completeness of methane combustion obtained from experimental data processed using a dimensional procedure. One can see that in the range Tin > 900 K, the completeness of combustion in a short kerosene chamber is comparable to the values in the chamber which runs on methane. Figure 4 shows calculated values of the total pressure recovery factor in the MAI experiments. As a whole, the values of in this range of studies are satisfactory: the coecient varies from 0.630.7 (at Min = 1.4) to 0.134 (at Min = 2.9). We should note the almost equal values = 0.310.36 during combustion of kerosene under conditions modeling Mach numbers Mf = 4 and 5 (Tin 940 and 1150 K). This is due to the fact that these tests were performed at the same ow rate coecient at the chamber inlet (nozzle of Min = 2.1). In the experiments at CIAM at Tin = 860880 K, the value of was 0.5.

REFERENCES
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Problems of Implementing Ramjet Operation

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