TechScape: The Science, Technology and Education Journal of IIT Jodhpur
Volume 01, Issue 03, August 2021

High Altitude Testing Facilities for Rocket Engines based on Ejector-Diffuser Systems

Ghanshyam Bharate & Arun Kumar R

The static testing of a rocket engine is one of the critical and mandatory test procedures for any space mission involving rocket propulsion. Among the static testing of various rocket stages, the most challenging is the testing of the upper stage rocket motors. This is because the upper stage rocket motor should operate at a vacuum pressure condition and maintaining such low vacuum pressure in a test chamber is a challenging task. The conventional method of using vacuum pumps may not provide the required vacuum level due to the enormous mass flux ejected by the rocket motor into the vacuum chamber. An alternative method for creating low back pressure is by pumping out the fluid from the vacuum chamber using the nozzle jet itself. It is well known that a high momentum nozzle jet (primary flow) inducts and entrains the surrounding fluid (secondary flow) due to the momentum exchange between the two streams. This principle can be used to create a vacuum condition at the nozzle exit by exhausting the nozzle jet into a confined chamber with upstream closed and such devices are known as the vacuum ejectors or zero secondary flow ejectors [1-3]. In general, the high-altitude test facilities (HAT) used for testing upper stage rocket motors employ the vacuum ejector principle for maintaining the required vacuum level at the nozzle exit. A schematic of a typical high altitude testing facility is shown in Fig.1 (a). As shown in Fig.1 (a), the nozzle will be kept in a vacuum chamber (secondary chamber) and is connected to a diffuser system for pressure recovery. Two configurations of ejector-diffuser systems are commonly used for HAT systems, 1) straight ejector diffuser (SED) and 2) the second throat ejector diffuser (STED) system [3]. The schematic of the two configurations can be found in Fig. 1(a) and 1(b), respectively. In a straight ejector diffuser (SED), a straight constant area duct is used as the diffuser section and in the second throat ejector diffuser (STED) system a convergent area section followed by a constant area duct is used as the diffuser section. Both these configurations utilize the shock cells developed in the duct for the pressure recovery.

It is seen from the literature that the ejector system in high altitude testing (HAT) facility operates in two modes during the initial transient starting phase, where the stagnation pressure in the combustion chamber builds up to a steady state. During this transient stage, the stagnation pressure (P0) at the inlet of the rocket nozzle increases and the jet expands continuously. In the first operation mode of the ejector, the jet boundary of the nozzle plume is not attached to the outer wall and this mode is called the un-started mode [3]. However, at a specific inlet stagnation pressure P0, the underexpanded jet coming from the nozzle impinges on the outer wall and produces shock cells. These shock cells help in pressure recovery and the outer duct downstream to the impingement point acts as a diffuser. The condition at which the supersonic jet attaches with the outer duct is therefore referred to as the starting of the diffuser and the corresponding nozzle inlet stagnation pressure (P0st ) is called the starting pressure of the diffuser [3]. As the diffuser attains the started mode, the jet boundary seals the vacuum chamber from the diffuser downstream and blocks further induction of fluid from the vacuum chamber. As a result of this, the minimum vacuum chamber pressure can be expected at the onset of started mode operation. The minimum vacuum level and the started mode pressure in vacuum ejectors depends on many geometric parameters, like, diffuser length to diameter ratio (L/D), nozzle position, nozzle area ratio, annular gap between the nozzle and the diffuser etc. There have been plenty of studies in the past to optimize these geometric parameters and yield better performance characteristics in vacuum ejectors used for HAT facilities. However, these optimization studies are limited by the inadequate understanding of the fluid dynamics of vacuum ejector diffusers, particularly during the starting process. Hence, this study investigates the process of vacuum generation in a second throat vacuum ejector system employed for the high-altitude testing of rocket motors using the computational fluid dynamics method.

Fig.2 shows the nature of vacuum generation for a typical axisymmetric second throat ejector diffuser system used in high altitude testing. It is seen that the vacuum generation progresses with four distinct stages. The initial stage consists of a gradual and oscillatory vacuum generation which is followed by a transition region (stage-2) where the vacuum pressure drops faster. The transition stage progresses to a rapid evacuation stage (stage-3), where a rapid reduction in vacuum chamber pressure is observed and this has been reported by many past studies. However, the present numerical study shows that the rapid evacuation is again followed by a gradual vacuum generation process, before reaching the minimum vacuum level or the started mode (labeled as stage-4 in Fig.2). A recent experimental study by Arun et al. [4-5] also reported the presence of a gradual evacuation stage after the rapid evacuation process. However, the reason behind the reappearance of gradual evacuation after the rapid evacuation was not properly explained in their study.

In order to further investigate the physics behind the staged evacuation, the static pressure histories at the nozzle exit and in the vacuum chamber have been compared (Fig.3 (a)), since the fluid induction from the vacuum chamber strongly depends on the pressure gradient between these two sections. The transient start-up process can be essentially divided into several quasi-steady processes and the vacuum chamber attains a quasi-steady pressure condition during each of these processes through a dynamic pressure equilibrium existing between the vacuum chamber and the primary jet at the nozzle exit plane. This can be clearly noticed from Fig. 3(a), which shows that the vacuum generation in the vacuum chamber closely resembles the pressure history at the nozzle exit. During the stage-1 evacuation, the nozzle exit pressure is relatively higher and it reduces gradually, as shown in Fig. 3(a). The static pressure contours during the stage-1 evacuation (Fig. 3(b)) show that the nozzle is at un-started mode with multiple shock cells appearing inside the nozzle. Due to these multiple shock reflections, the pressure at the nozzle exit will be larger during the stage-1 evacuation, which in turn results in a relatively larger vacuum chamber pressure (Fig. 3(b)). From Fig. 3(a), it is observed that at a time instant close to 0.414 seconds the slope of the static pressure curve suddenly changes and this is identified as the starting of a ‘transition region’ where the gradual evacuation starts changing to a rapid evacuation (point ‘c’ in Fig. 3(a)). It is observed that this ‘transition region’ starts when the second shock cell leaves the nozzle exit (Fig. 3 (c)). It is also seen from Fig. 3(a) that at a time instant close to 0.464 seconds, the ‘transition region’ changes to a rapid evacuation stage, where a rapid rate of vacuum chamber pressure reduction can be seen. (Point ‘e’ in Fig. 3(a)). The methodology by which the onset of rapid evacuation has been computed is mentioned in the appendix. This rapid reduction in static pressure at nozzle exit can be related to the dynamic movement of shock wave in the nozzle during the start-up process, as shown in the pressure contours in Fig. 3(e) to 3(h). As the primary jet total pressure increases the shock structure in the nozzle moves downstream with the shock cells exiting the nozzle. After a certain pressure ratio, only the first shock cell remains inside the nozzle, as shown in Fig. 3(d). A slight increase in stagnation pressure from this stage (point-e in Fig. 3(a)), results in a rapid movement of the first shock cell to the nozzle exit plane, as clearly observable from Fig. 3(e) to 3(h). The first shock cell exhibits a Mach reflection structure with an incident shock, a reflected shock, and a Mach stem meeting at a single point. As the first shock cell is pushed towards the nozzle exit plane, the reflected shock wave in the Mach reflection structure exits the nozzle, as shown in Fig.3 (e) to 3(h). It should be noted that in a Mach reflection shock system, the pressure rise will be maximum behind the reflected shock wave and as it rapidly moves out of the nozzle exit plane, the static pressure at the nozzle exit reduces suddenly which in turn reduces the vacuum chamber pressure. The comparison of the pressure history at the vacuum chamber (points ‘e’ to ‘h’ in Fig. 3 (a)) and the corresponding pressure contours in the nozzle (Fig.3 (e) to 3(h)) clearly shows that the rapid reduction in vacuum chamber pressure happens during the rapid movement of the reflected shock wave out of the nozzle exit plane.