Project Longshot/Mission Profile

2.0 MISSION PROFILE

2.1 Objective Star System

The star system chosen as our objective system for this mission is the Alpha Centauri trinary star system. There are three main reasons for choosing this system as the destination of the probe: (1) the system's proximity to our solar system; (2) the scientific interest of a trinary system; (3) sending a probe to this star system provides an opportunity to make great advances in the field of astrometry.

One of the main concerns regarding the success of the mission is the ability of the probe to be in good working order when it reaches the objective system. While any interstellar mission will require transit time in excess of a century, the Alpha Centauri system is the closest to our own, and thus provides the best opportunity to deliver a functional probe to another star system.

While it would be a great achievement to do a close-up study of any other star system, the Alpha Centauri system promises to be a a particularly interesting objective. Alpha and Beta Centauri, which are type G2V and dK1 stars, respectively, orbit each other with a separation of 11 to 35 AU. The third star of the system, Proxima Centauri, orbits the other pair at a distance of about 1/6 of a light year. Proxima Centauri is a type dM5e red dwarf, which exhibits sudden changes in magnitude. It is also one of the smallest known stellar masses. Proxima's position should provide and opportunity for the probe to pass relatively close by on its way to its final destination, a close orbit about Beta Centauri. There is also the possibility that this star system contains planets. Advances in astronomy before the probe is launched should provide much more information about the system to help plan the probe's exploration.

Perhaps the greatest contribution that the mission will make to the scientific community will be in the field of astrometry. Sending a spacecraft to the Alpha Centauri system provides an opportunity to make parallax measurements with a baseline of 4.34 light years (see Fig 2.1a). This is over 63,000 times longer than the present method, which uses the semimajor axis of Earth's orbit as a baseline (see Fig 2.1b). At this time, parallax measurements are only accurate to about 20 parsecs from the Sun. The longer the baseline would allow accurate measurements of stellar distances of more than 1.2 million parsecs. If the probe lasts long enough, it has the potential to accurately determine the distance to hundreds of trillions of stars. Knowing the distance to a star is vital in determining its properties. Such an accomplishment would keep astronomers busy for quite some time.

PARALLAX DETERMINATION

THE BASELINE FROM CENTAURI SYSTEM IS LONG

Fig. 2.1a

PARALLAX DETERMINATION

$\mathbf {d(parsecs)={\frac {b(AU)}{\mathbb {P} (arcsecs)}}}$

From Earth the baseline, b, is short.

Fig. 2.1b

2.2 Operations

Assembly of the probe will occur in two phases: major components and subsystem will be assembled and tested on the ground, and then sent to LEO for final assembly and integration into the spacecraft. Transfer to orbit will require multiple launches with initial estimates indicating that launches will be carried out in the following manner, (based on current projections):

• Fusion engine - one launch on an ALS-class vehicle.
• Main structure - one launch on a Shuttle-class vehicle.
• Fuel tanks - one launch on an ALS-class vehicle.
• Fuel - one launch on a Growth ALS-class vehicle.
• Payload - one launch on an ALS-class vehicle.
• Upper stages for the initial phase of the mission will be launched as required based upon advances in upper stage technology.

The fusion drive will be launched as a finished unit due to its complexity. It will be equipped with a plug-in type interface for integration with the remainder of the ship. The main structure will be a collapsible space frame, which will be erected from its stored configuration by personnel at the space station. Additional structure will also be orbited in collapsed form for erection at Space Station. The fuel tanks will be built of sheet aluminum with formed end caps. The end caps will be launched in a completed and stacked configuration and the sheeting for the bodies will be rolled into a solid cylinder for launch. Assembly will occur in orbit using advanced adhesives technology. The fuel for the fusion engine will consist of He3 and deuteruim. The He3 will probably be manufactured on earth using particle accelerators, and then both components of the fuel will be launched in liquid form to orbit for processing into pelletized reactor fuel. The payload, like the fusion drive, will be launched fully-assembled, and have a simple plug-in interface to the rest of the system.

By the time that this mission will be flown, the experience gained through the assembly of Space Station, and the ongoing orbital operations which it is expected to support, will provide an adequate base of technology for personnel to assemble the probe. During assembly, systems checks will be run as each component is integrated into the spacecraft, to prevent failures which would necessitate disassembly of components assembled in orbit. Following assembly, a comprehensive system check-out will be made prior to ignition of the upper stages.

The fission reactor will be used to provide the initial power required to start the fusion reactor. Once the drive is ignited it will become self-sustaining and provide enough excess power to operate the systems in use during the transit. During this phase, fuel will be drawn from symmetrically located pairs of fuel tanks to avoid instability caused by non-homogeneous mass distribution. As each pair of tanks is emptied, they will be discarded. Tank volume has been calculated so that four of the ship's six tanks will have been jettisoned by the time it reaches the turn-around (see Figs 2.2a and 2.2b). This point has been chosen so that reversal of the spacecraft and reignition of the main drive will result in arrival at the target at proper insertion speed. When the turn around point is reached, the drive will be shut down, and the probe rotated 180 degrees with respect to its velocity vector using the attitude control system. Following the turn around maneuver, the fission reactor will again be used to provide power to restart the drive. These activities will occur at the 71.27 year point in the mission (see Figs. 2.2c and 2.2d for graphs and appendix for calculations).

Throughout this phase of the flight, data from experiments on interstellar space will be returned to earth at low data rates. This will serve two purposes, providing scientific data, and ensuring that contact will be maintained with the probe. Since the probe will be fully autonomous, any problems with the communications system, due to degradation of the transmitting equipment or faulty link analysis will have to be corrected by making improvements to the receiving equipment. Maintaining constant contact will allow the required lead time to implement any necessary corrective measures in the receiving equipment.

After deceleration is completed, the two remaining fuel tanks and the fusion drive will be discarded and the instruments deployed. The attitude control system which was used to rotate the entire spacecraft will be used as necessary to place the

INTERSTELLAR FLIGHT PHASE

1.

INITIAL CONFIGURATION

2.

TANKS ONE AND TWO JETTISONED AT TIME T=+33.35 YEARS

3.

TANKS THREE AND FOUR JETTISONED AT TIME T=+66.7 YEARS

2.

AT T=+71.266 YEARS, THE SPACECRAFT ROTATES. AT T=+100 YEARS (ARRIVAL AT TARGET STAR), TANKS FIVE AND SIX JETTISONED, ENGINE AND SHIELDS SEPARATE.

Fig. 2.2c

Fig. 2.2d

probe into an elliptical orbit around Beta Centauri and to maintain this orbit. Once in orbit, high data rate collection and transfer will begin based on both a preprogrammed series of studies and a set of prioritized experiments (determined by what is found in the system). This will continue until the spacecraft's nuclear plant ceases to produce enough power to operate the communications lasers, which will constitute the end of the mission.

2.3 Orbits

Once the probe is assembled in orbit at the space station, it will be nudged into an independent but similar orbit to prevent damage to the station due to the exhaust from the first upper-stage burn (see Fig. 2.3a). This burn will be made to increase the inclination of the probe orbit from 28.5 degrees to 37.5 degrees so that it will sum with the obliquity and result in an orbit inclined 61 degrees relative to the ecliptic. This stage will be jettisoned, and the spacecraft will then escape Earth with a second burn which occurs at the ascending node of the probe's orbit about the earth (see Fig. 2.3b). This point then becomes the ascending node of the probe's heliocentric orbit, which is a circular orbit at 1 AU and at an inclination of 61 degrees. The second stage will then separate, and three months (i.e. one-fourth of an orbit) later, the third and final upper-stage will burn. This will send the spacecraft on an escape trajectory toward the Centauri system (see Fig. 2.3c), which is located at a declination of −61 degrees to the ecliptic. Upon completion of the interstellar phase of the mission the probe will be inserted into an eccentric orbit about Beta (see Figs. 2.3d and 2.3e).

Assuming a space station orbit of 300 kilometers at an inclination of 28.5 degrees, the following values were determined for required velocity changes for the probe. The Alpha Centauri star system is at a

SYSTEM INJECTION

FINAL ORBIT ABOUT BETA CENTAURI

ALPHA-BETA SEPARATION: 11-35 A.U.
R_p ABOUT BETA: 1 A.U.
R_A ABOUT BETA: 2.5 A.U.
SEMI-MAJOR AXIS: 1.75 A.U.
ECCENTRICITY: 0.429
PERIOD: 2.5 EARTH YEARS
VELOCITY AT APOGEE: 13.1 km/sec
VELOCITY AT PERIGEE: 32.9 km/sec

declination of −61 degrees relative to the ecliptic of our solar system. At the optimum launch position, the obliquity (23.5 degrees from the ecliptic) will sum with the inclination angle of the spacecraft's orbit (28.5 degrees from the equator). This will make the probe's orbit 52 degrees from the ecliptic. Thus, a 9 degree plane change is required to put the probe into a 61 degree orbit around the earth (relative to the ecliptic). From this orbit, another velocity change is necessary to escape the earth's gravity and inject the probe into a circular orbit about the sun at 1 AU and a 61 degree inclination. At this point, there are two options to be considered, depending on the fuel source, Earth or Jupiter. For the first option, the probe will leave from this orbit directly for Alpha Centauri, in which case it will change velocity at the perihelion of the transfer orbit between our solar system and Alpha Centauri. For the second option the spacecraft will transfer to a Jupiter-distanced heliocentric orbit, using a velocity change at the perihelion of the Earth to Jupiter transfer orbit, and inject into an orbit around Jupiter at a 61 degree inclination (to take on fuel). The probe will then escape back into a Jupiter-distanced heliocentric orbit at 61 degrees relative to the ecliptic. Finally, it will escape the solar system and head for Alpha Centauri, with a burn at perihelion of the Sol-Centauri transfer orbit.

(see appendix for actual number calculations)

1.	Delta-V for plane change of 9 degrees:	1.2123 km/s
2.	Delta-V to escape earth:	3.2002 km/s
3.	Delta-V to escape solar system:	12.4273 km/s
4.	Total Delta-V for Option #1:	16.8398 km/s

For Option #2, the first two numbers are the same.

5.	Delta-V to enter Earth-Jupiter transfer:	8.8223 km/s
6.	Delta-V to orbit around Jupiter:	7.4814 km/s
7.	Delta-V to escape Jupiter:	6.1705 km/s
8.	Delta-V to escape solar system:	5.4087 km/s
9.	Total Delta-V for Option #2:	32.2954 km/s

Obviously, the option of picking up fuel mined from the atmosphere of Jupiter would be impractical because of high cost and complication.

For the objective orbit within the Centauri system, Beta was chosen as the target star because it is a dK-type star, about which we have very little data, while Alpha is a G2 type star like our own, which we have studied extensively. the orbit chosen is based on an assumption that Alpha and Beta (which vary in distance between 11 and 35 AU) will be at the lesser of the two distances. The orbit chosen is an elliptical orbit around Beta for which the aphelion lies on the line between Alpha and Beta. The perihelion radius was set at 1 AU. The aphelion was determined based on a requirement that the gravitational effect of Alpha would not exceed 5% of the gravitational effect of Beta. Accordingly, an aphelion of 2.5 AU was chosen. At this aphelion the gravitational force on the probe due to Alpha will be equal to 3.3% of the gravitational force due to Beta. Based on the chosen perihelion and aphelion radii, the orbital parameters were determined:

a = 1.75 AU
e = 0.428572
E = −216.95 kmE2/secE2
T = 2.4977 Earth years
V_a = 13.1 km/s
V_p = 32.935 km/s

The perihelion velocity will be the velocity to which the probe must be slowed for proper insertion into this orbit.