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Supersonic Retropropulsion for High Mass Entry Systems

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Retropropulsion is basically using the thrust provided by the rocket engines against the velocity vector so as to decelerate the object or specifically the space craft. Consequently, supersonic Retropropulsion is deceleration of objects having supersonic speeds using Retropropulsion methods. This is a method lately being considered by many space organisations such as NASA or SpaceX for landing high mass payload safely on the surface of Mars. While it may seem that since man has already conveniently landed heavy payloads on the moon, landing it on a planet like Mars should be similar, it is not really the case. One of the main factors contributing to this dissimilarity is the fact that Mars exerts nearly twice the gravitational pull as compared to the moon making it much more difficult to decelerate. A few other method such as parachutes and airbags have their own disadvantages such as the fact that Mars’ atmosphere is much thinner than Earth’s, hence rendering convenient sized parachutes unusable. Similarly as the mass of the payload increases, the required size of airbags also increases hence making them impractical for use.

Introduction:

Investigation into the interaction of supersonic retropropulsion with blunt body aerodynamics began in the early 1950s. Experimental work with small-scale wind tunnel models by Love, Huff and Abdalla, and Moeckel focused on shock-boundary layer phenomena and the effects of nozzle flow on boundary layer transition. These investigations were among the earliest observations of the aerodynamic drag reductions and associated flowfield stability transitions for configurations with a nozzle located along the body centerline. Moeckel ] was among the first to observe an aerodynamic drag reduction in supersonic retropropulsion configurations with a centerline nozzle application. Moeckel also observed flow separation on the forebody with the same configuration. Both effects were later observed in other experiments with a supersonic retropropulsion nozzle along the centerline of a blunt entry body.

These early works are consistent in observing that increasing thrust coefficient moves the boundary layer transition closer to the nose of the body. Additional work on supersonic jet flow and jet-body interactions laid the groundwork for future wind tunnel testing of supersonic nozzle exhaust effects on body surface pressure distributions and flowfield stability. While many of these body shapes were not the blunted-cone entry vehicle shapes flown in the 1960s and 1970s, these works established the fundamental physics of shock-boundary layer interactions and motivated later investigations to apply such interaction effects to blunt body entry vehicles for planetary exploration.

In most current EDL systems, decelerating the vehicle from hypersonic to subsonic speeds is achieved using the aerodynamic drag of the entry vehicle and other aerodynamic decelerators such as parachutes. At Mars, high entry masses and insufficient atmospheric density often result in unacceptable parachute deployment and operating conditions. An alternative deceleration approach is to initiate retropropulsion while the vehicle is still traveling supersonically. Supersonic retropropulsion may be an enabling technology for systems with high ballistic coefficients operating in thin atmospheres such as at Mars.

The ballistic coefficient (β) of a body is a measure of its ability to overcome air resistance in flight. It is inversely proportional to the negative acceleration: a high number indicates a low negative acceleration—the drag on the vehicle or projectile is small in proportion to its mass. The ballistic coefficient of an atmospheric reentry vehicle has a significant effect on its behaviour. A very high ballistic coefficient vehicle would lose velocity very slowly and would impact the Earth’s surface at higher speeds. In contrast, a low ballistic coefficient would reach subsonic speeds before reaching the ground. Ballistic coefficient is the ratio of entry mass to hypersonic vehicle mass area i.e.

The six successful Mars landers have all had ballistic coefficients under 100 kg/m 2. Requirements for substantial (one or two orders of magnitude) increases in landed mass and aeroshell packaging constraints result in significantly higher ballistic coefficients for human-scale Mars entry systems (400 kg=m 2 and higher) . For these high ballistic coefficient vehicles, the thin Martian atmosphere and the challenge of extending supersonic DGB parachutes to the required dimensions and deployment conditions combine to severely reduce the timeline available for deceleration and transition from a hypersonic entry vehicle to a terminal landing configuration. System-level studies to assess the required performance of high-mass entry systems recommend the development of alternative supersonic decelerators, a challenge potentially addressed by supersonic retropropulsion (SRP).

SRP is the initiation of a propulsive deceleration phase while the vehicle is traveling supersonically. The SRP flowfield is a complex interaction between a typically highly underexpanded jet and an opposing supersonic freestream. The aerodynamic characteristics of the vehicle are affected by both the thrust of the retropropulsion and by the interaction between the retropropulsion and the supersonic freestream [3]. The retropropulsion plumes obstruct the oncoming freestream, and a bow shock forms. The supersonic freestream flow is decelerated to subsonic by the bow shock, and the supersonic jet flow is decelerated to subsonic by a Mach disk (normal shock). A contact surface separates the subsonic retropropulsion flow and the subsonic freestream flow behind the bow shock. The shape of the contact surface acts as the effective vehicle geometry, or flow obstruction, seen by the supersonic freestream flow.

Retropropulsion Systems:

Starting with the two Viking landers in 1976 and continuing through the successful Mars Phoenix mission in 2008, NASA and its partners have used similar entry, descent, and landing (EDL) system architectures with incremental improvements to deliver robotic payloads to the surface of Mars. This architecture is based on a rigid, blunt-body aeroshell (spherically-blunted, 70-degree half-angle cone), a supersonic disk-gap-band parachute, and a subsonic propulsive descent system used in sequence to decelerate the payload in the thin Martian atmosphere. Viking-based systems have delivered robotic payloads to the Martian surface from both orbit (Viking) and from direct entry (Mars Pathfinder, Mars Exploration Rovers, and Mars Phoenix). Table 1 summarizes previous Mars EDL systems developed by the United States that have successfully delivered robotic payloads, all of which were less than 0.6 metric tons (t).

The focused technology development program preceding the Viking missions in the 1960s and 1970s developed supersonic retropropulsion to nearly the level of maturity the concept has today. The eventual selection of a supersonic DGB parachute system and subsonic propulsive terminal descent phase for the Viking landers ended much of the research efforts to develop supersonic retropropulsion. Only recently has interest in supersonic retropropulsion resurfaced. The applicability of Viking EDL technologies to the high mass planetary entries needed for human Mars exploration has been shown to be constrained by deployment conditions and performance at higher Mach numbers of supersonic DGB parachutes1. This resurgence of interest in human Mars exploration has resulted in systems-level studies to assess the required performance of these high-mass entry systems, and the conclusions of these studies, in general, recommend the development of alternative supersonic decelerators, a challenge potentially addressed by supersonic retropropulsion.

Of the past Mars EDL architectural studies, only the investigations by Christian, and Wells, initiate retropropulsion at supersonic conditions. In these two references, the aerodynamic interaction of supersonic retropropulsion was not modeled (i.e., the deceleration was assumed to be independent of nozzle location). In addition, the aerodynamic drag was set to zero, and only the axial force due to thrust was modeled. Past experimental evidence suggests that, for configurations with the nozzles located at the forebody periphery, inclusion of aerodynamic interaction effects into supersonic retropropulsion performance models will reduce the propellant mass required.

The general relationship between the required descent propellant mass fraction and vehicle ballistic coefficient is shown in Figure 1. Descent propellant mass fraction is defined as the ratio of propellant mass required to decelerate the vehicle from supersonic retropropulsion initiation to subsonic conditions to entry mass. This figure illustrates the increasing impact of aerodynamic effects on required descent propellant mass fraction as ballistic coefficient increases for a 15 m diameter Apollo aeroshell and CT = 1.0. The data presented is for the minimum Mach initiation to reach subsonic conditions at 3 km altitude.

Figure 1. Descent Propellant Mass Fraction Impact of Neglecting Aerodynamic Effects During Supersonic Deceleration for a Configuration with Nozzles Located at the Vehicle Periphery for CT = 1.0.

Supersonic Retro-Propulsion (SRP) is included in the MSL and involves initiating propulsive deceleration at supersonic Mach numbers by directing engine thrust into the oncoming freestream flow. The complexity of the interaction between the supersonic freestream and the retro-propulsion exhaust flow is illustrated notionally in Figure 2 for a single jet. For the case shown, the supersonic jet plume terminates in a shock behind the main bow shock, with a free stagnation point in between. The location and existence of these features are largely a function of the ratio of jet total pressure to freestream total pressure. The total effective drag coefficient using a system such as that shown in Figure 1 is derived from the aeroshell’s aerodynamic drag (CD = D/q∞Aref) and the propulsive drag provided by the engines (thrust coefficient, CT = T/q∞Aref):

CD,Total = CD + CT (1)

Depending on the entry system mass and deceleration requirements, especially for human-scale payloads, the propulsive drag term in the above equation may be the dominant contributor to the total effective drag. The EDLSA team considered SRP systems that delivered thrust coefficients above 10, which far exceed the typical bluntbody drag coefficient of around 1.5. Previous wind tunnel studies suggest that the use of retro-rockets can at best maintain the native aerodynamic drag, and at worst reduce the aerodynamic drag to nearly zero, depending on the retro-rocket configuration and thrust coefficient. In the case where aerodynamic drag is eliminated or is small relative to the propulsive drag term, the total entry vehicle drag is derived from the engine thrust alone.

[image: ]

Figure 2 : Interaction Between Supersonic Freestream and Exhaust from a Single Nozzle

The primary flow structures are the bow shock, free stagnation point, jet terminal shock, and the recirculation regions. The location, degree of formation, and stability of these features are a strong function of the ratio of jet total pressure to freestream total pressure. This total pressure ratio is often represented by thrust coefficient, assuming a fixed freestream stagnation pressure for a given test condition.

The entire flowfield structure is dependent on the formation of the stagnation point. The freestream must decelerate to zero velocity, first from supersonic to subsonic through a shock wave, then from subsonic to zero velocity at the stagnation point18. The nozzle flow undergoes a similar deceleration through mixing, viscous dissipation, or a normal shock, depending on the strength of the nozzle exhaust flow. For the case of supersonic retropropulsion, the stagnation region consists of two supersonic regions, the freestream and nozzle flow, and a subsonic region divided by a contact discontinuity, the stagnation point.

The interaction of the jet with the opposing supersonic freestream in central configurations has been observed to cause the flowfield to transition from stable to highly unstable and back to stable as the total pressure ratio (or thrust coefficient) increases. A stable flowfield occurs when the bow shock is close to the body, and the jet flow does not penetrate the bow shock; in this case, the flowfield structure is not oscillating. An unstable flowfield occurs when the jet shock penetrates the bow shock and the total shock displacement is significantly greater than the displacement characteristic of the stable condition. This displacement increases to a maximum with increasing total pressure ratio and then collapses back to a displacement similar to the original stable case.

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