Solar Thermal Propulsion for an Interstellar Probe

The conceptual design of an interstellar probe powered by solar thermal propulsion (STP) engines will be described. This spacecraft will use STP engines to perform solar gravity assist with a perihelion of 3 to 4 solar radii to achieve a solar system exit velocity (see Figure 1). Solar thermal propulsion uses the sun's energy to heat a working fluid with low molecular weight, such as hydrogen, to very high temperatures (around 3000 K). The stored thermal energy is then converted to kinetic energy as the working fluid exits a diverging nozzle, resulting in a propulsion system with a high specific impulse (ISp). Figure 1. Proposed trajectory for an Interstellar Probe This paper will evaluate the feasibility of using STP for an interstellar probe. The technology readiness and required system improvements will be identified. INTRODUCTION Although traveling to the stars is still beyond the reach of known technology, a probe capable of penetrating the interstellar medium may be within that reach." Using a high-thrust, high specific impulse solar thermal propulsion system this probe could potentially achieve an asymptotic escape speed of 20 astronomical units (AU) per year. This mission would utilize a Jupiter flyby and powered perihelion gravity assist.'" Solar thermal propulsion is an innovative concept that uses the Sun's energy to heat a low-molecular weight fluid such as hydrogen." The thermal energy stored in the heated fluid is then converted to kinetic energy by expansion through a diverging nozzle. This results in a high efficiency propulsion system. Spacecraft using STP have been proposed for orbital transfer, interplanetary, and other delta velocity missions. An interstellar probe using solar thermal propulsion to perform a perihelion maneuver has been proposed.'" The probe itself is small (~50kg). Prior to the solar gravity assist, it is surrounded by a much more massive carrier or cocoon. This cocoon protects the probe and consists primarily of the heat shield and STP system. The thermal model is shown in Figure 2. This diagram shows a vehicle with a *© 2001 ATK Thiokol Propulsion Corp. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission Figure 2. Proposed Interstellar Probe with the pre-maneuver heat shields deployed 1 American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. deployable exterior heat shield (shown in red) that protects the interstellar probe and propulsion system during its perihelion maneuver. The shield, made of carbon-carbon composite, is opened and oriented toward the sun as the spacecraft approaches perihelion. This hinged device will form a 30° shadow (at 4 solar radii) to thermally protect the probe/carrier assembly. The center third of the heat shield is the STP engine. This engine is a channeled heat exchanger as shown in Figure 3. During the perihelion maneuver, low-pressure hydrogen (< 1380 kPa) flows through this heat exchanger at 232 g/s for 15 minutes. The channels are sized to achieve a high heat exchanger effectiveness, resulting in a hydrogen exit temperature approaching the theoretical limit for the solar heat source, while minimizing pressure drop of the hydrogen as it passes through the device. by rocket pioneer Hermann Oberth. Measuring the AV in km-sec" the asymptotic escape speed from the solar system is approximately Figure 3. Section of the STP Engine (Channeled heat exchanger) Once the hydrogen exits the heat exchanger, it is transferred through a plenum to a centroidal nozzle, where the stored thermal energy is converted to thrust. It should be noted that in this feasibility study the designs of the entrance or exit plenums have not been addressed, although a mass allocation has been made. PROPULSION SYSTEM The goal for the system is to achieve an escape velocity from the solar system of 20 AU-yr" (94.8 km-s"). The mission requires a fast transfer to Jupiter and flyby, where an unpowered gravity assist reduces perihelion of the transfer orbit. In addition, the gravity assist accomplishes a plane change that places the trajectory where a perihelion change in velocity, a delta velocity (AV) maneuver, yields the desired solar system escape direction. The AV required at perihelion is the difference between the perihelion velocities of the hyperbolic solar-systemescape and elliptical Jupiter-to-perihelion transfer trajectories. The idea of using a high-ISP, high-thrust maneuver close to the Sun was first identified in 1929 74 yr where rp is the distance from the center of Sun in terms of Sun radii (R$). This equation gives an approximation to the required propulsive maneuver for a given asymptotic speed. With this relationship the required AV is approximately 14.6 km-s" for a 20 AU-yr" escape velocity from a 4 Rs perihelion. With a closer approach, 3 Rs perihelion, the required AV is reduced to approximately 12.6 km-s". However, there are additional factors that must be considered for closer approaches, such as thermal protection methods along with the associated mass increase. As a baseline, the value of AV =15 km-s" was chosen for the propulsion "target" to enable an asymptotic solar system escape of roughly 20 AU-yr" from a 4 Rs perihelion. The relationship between the required AV, the system ISp and the mass ratio (MR) or the mass fraction (Q is where g0 is 9. 8 1x1 0" km-s" for a AV measured in terms of km-s". The mass ratio, MR, is defined as the ratio of the final or dry mass (mf) to the initial or wet mass (mi), and the mass fraction, £, is the ratio of the propellant mass (mp) to mj. The relationship shows that the required AV is directly proportional to ISP, and a logarithmic relationship exists with the mass ratio or the mass fraction. To achieve the AV maneuver goal of 15 km-s", without an unrealizable propellant mass, the ISP must be maximized. An analysis was made to explore the possible ISP capabilities of potential propellants. (These are propellants in that an external thermal source heats chemically inert material for expulsion in STP instead of providing heat chemically by "burning" a fuel.) These include liquid hydrogen (LH2), ammonia (NH3), and methane (CH4). The baseline propellant selected for reference is LH2 since it has the potential for the highest ISP. Each candidate was analyzed at various pressure and temperature conditions at the nozzle inlet and allowed to expand through a range of nozzle expansion ratios of 20:1 to 100:1. Pressures of 517 kPa (76 psia) and 1380 kPa (200 psia) were evaluated in combination with a set of temperatures ranging from 1500 K to 3500 K. The feasibility of reaching these temperatures is discussed American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. later in this report, but it is assumed that at least 2400 K is possible and a 3500 K maximum allows a margin for carbon-carbon material that melts at approximately 3800 K. The analysis results for H2 (see Figures 4 and 5) show the highest ISp levels of the three propellants studied. Results of this analyses indicates that at the lower temperatures the pressure does not affect the ISp, while at the upper end of the temperature range, the lower pressure allows for more dissociation of the propellant and thus higher ISP values. The maximum ISP predicted for temperatures of 2400 K, 3000 K, 3300 K and 3500 K are 860 sec, 1037 sec, 1166 sec and 1267 sec, respectively. An observation that can be made from the analysis results is that the nozzle area expansion ratio does not have a large affect on the ISP level at ratios greater 50:1. There is less than a two percent increase in ISp from an expansion ration of 50:1 to 100:1. Hydrogen (H2) Specific Impulse (Vacuum ) Relative to Temperature and Nozzle Expansion Ratio Pchamber = 517 kPa Figure 4. Hydrogen ISp at 517 kPa Initial Pressure Hydrogen (H2) Specific Impulse (Vacuum ) Relative to Temperature and Nozzle Expansion Ratio Pchamber = 1380 kPa