Project Longshot/Spacecraft Systems

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Project Longshot
Spacecraft Systems

3.0 SPACECRAFT SYSTEMS[edit]

3.1 Power[edit]

Power for the instruments, computer, communication lasers, and star trackers will be supplied by a 300 kilowatt nuclear reactor. This reactor will be compact-sized, have a low specific mass, long life, high reliability, and a variable power output. Different systems were compared, and a design published by Jones, MacPherson, and Nichols, of the Oak Ridge National Laboratory was chosen and scaled down to suit the needs of the Centauri mission.

This power system concept defines a nuclear reactor with ceramic fuel, clad with refractory metal, and cooled by liquid potassium (1365 K). Also described is a direct, closed Rankine power conversion cycle, and a large tetraxial flywheel energy storage system featuring graphite composite materials and magnetic bearings. The fuel is is enriched uranium nitride pellets. The reactor and flywheel systems will be constructed as separate modules designed to fit in the shuttle cargo bay.

The reactor will boil potassium which will then be piped through a turbine that will convert the thermal energy to mechanical energy. The potassium is further cooled by flowing through the heat radiators and is then recycled back into the reactor.

The electrical current produced by the spinning turbine is directed to the two (for redundancy) energy storage systems located at either end of the main structural truss. The modules contain flywheels which store the energy in their spin rate, and also provide attitude control by absorbing external torques.

Mass
reactor 500 kg
shielding 830 kg
turbine 230 kg
piping and miscellaneous 680 kg
8 flywheels 3400 kg
flywheel motors, structure 760 kg
Total mass 6400 kg

Power output is variable but will need to be a minimum of 250 kW during the in-system phase of the mission to power the communication lasers and computer.

3.2 Propulsion[edit]

3.2.1 In Solar System[edit]

Since it will not be practical (or safe!) to ignite the fusion drive while in earth orbit, another propulsion system will need to be used for the plane change and to escape the Earth and the solar system (see Orbits). Advanced solid rockets will be employed to accomplish these velocity changes. it is assumed that they will be similar in size and fuel to the Space Shuttle's SRB's with twice the specific impulse (near 600 seconds). Based on this assumption, the following parameters for the various upper stages were calculated:

(see appendix for calculations)

Booster Mass (kg)
Fuel/Total
Length (m) Diameter (m)
Plane change 352/417 16 3.7
Escape Earth 605/715 25 3.7
Escape Sun 252/298 11 3.7
Total mass: 1209/1430

3.2.2 Interstellar Transit[edit]

Developing a Propulsion system capable of meeting the 100-year interstellar travel time is the most difficult part of the mission design. 4.3 light years is an easily misinterpreted distance. It is equivalent to 41,000 terra meters (41E15 km) which would take the space shuttle just over 190,000 years (assuming it had escaped the solar system at the speed of Low Earth Orbit). Although 100 years is a long time, this requirement expects a three-order-of-magnitude leap over current propulsion technology.

3.2.2.1 Choosing the System[edit]

After the initial inspection of potential Interstellar Drive candidates, it was decided that chemical fuels would not be able to produce a three order of magnitude leap over current systems in the near future. Five alternate technologies were compared for their potential as Interstellar Drive candidates: Pulsed Fusion Microexplosions, Laser-pumped Light Sails, Ion Drive, High Temperature Thermal Expansion of Gas, and Matter Anti-matter Annihilation (see Fig. 3.2a for a summary table). After a thorough inspection of each of the five candidates it was decided that only the Pulsed Fusion Microexplosion was adequately capable of carrying out the mission requirements.

INTERSTELLAR DIVE TRADE-OFF STUDY


SYSTEM Isp
(1000s)
FEASIBILITY
FUSION MICRO-EXPLOSION 1020 MEDIUM
LASER-PUMPED LIGHT SAIL N/A* LOW
ION DRIVE 3.5–10 HIGH
THERMAL EXPANSION OF GAS 39 HIGH–V. LOW
MATTER/ANTI-MATTER DRIVE 100 EXTREMELY LOW

* EXTERNAL DRIVE SOURCE--3.75 TERRAWATT LASER


The Matter Anti-matter Annihilation could potentially be capable of producing the necessary specific impulse of a million seconds, but it was not considered to be feasible to create a system adequate for storing the anti-matter for 100 years under the limited power constraints of a spacecraft.

The ideal rocket equation was used to determine the potential specific impulse of an extremely high temperature expansion of gas through a nozzle. (See calculations and assumptions in appendix.) Using the critical temperature for sustained Deuterium fusion (3.9E8 degrees Kelvin) a specific impulse of 39000 seconds was calculated (1/300th the specific impulse desired). Since this specific impulse is insufficient for the mission requirements (using an extremely optimistic temperature under ideal conditions), this candidate was dropped.

Advances in technology for an accelerated ion drive (using a magnetic/electric field to fire charged particles out a nozzle) have brought the specific impulse to 3500 seconds. Although this is a current technology that could be implemented now at a relatively low cost, it is felt that the two remaining orders of magnitude will remain out of reach in the near future. Therefore, this candidate was also discarded.

In determining the feasibility of a Laser-pumped Light Sail, another method besides comparing specific impulse becomes necessary, since the drive is external. The single impulse required to reach the designated system in 100 years was determined to be 13,500 km/sec. The size of a laser with continuous output, to accelerate the payload to 13,500 km/sec in a year, is 3.75 Terra Watts. Since a micropulsed 1 Terra Watt laser has been developed, it is conceivable (although extremely unlikely) that the necessary laser could be invented within the next 20 years. The low feasibility, coupled with the lack of a system for decceleration[sic] into the Centauri System, led to the cancellation of this system's candidacy (see appendix for calculations).

3.2.2.2 Pulsed Fusion Microexplosion Drive[edit]

The Pulsed Fusion Microexplosion Drive is not a current, but rather an enabling technology. The system concept, modeled after the British Interplanetary Society's project DEADALUS, is to fire high energy particle beams at small fusionable pellets that will implode and be magnetically channeled out the nozzle (see Figs. 3.2b and 3.2c). The expected specific impulse is 1.02E6 seconds. The specific mass breakdown for separate sections (including fusion chamber, particle beam igniter system, and magnetic nozzle/inductor system) is included in the structures section (2.3.2). Finally, the entire system is expected to gimble a full degree in two axes to enable navigational corrections in three dimensions.

The type of fuel used in the pellets is of critical importance. Due to the extremes of temperatures and duress inherent in fusion reactions, a magnetic field is required to supplant the casing around the fusion

PULSED FUSION CONCEPT

Project Longshot - Design Report - Fig 3.2b-1.png

1. THE PELLET IS LOADED AND FIRED UPON BY SEVERAL HIGH POWER PARTICLE/LASER BEAMS.

Project Longshot - Design Report - Fig 3.2b-2.png
2. THE FUSION EXPLOSION IS DIRECTED OUT OF THE EXIT VIA THE MAGNETIC "NOZZLE".
Project Longshot - Design Report - Fig 3.2c.png

3. THE HIGH VELOCITY PULSE (~10000 KM/S) INDUCES A CURRENT IN THE COILS THAT SURROUND THE EXIT PORT. THIS ENERGY IS USED TO RE-CHARGE THE PARTICLE/LASER BEAMS AND THE MAGNETIC FIELD.

chamber. The problem involved with using such a field (besides the obvious requirement for immense quantities of energy) is that only charged particles will be channeled out the nozzle. Although the extreme temperatures will instantly ionize all of the atoms and molecules, any neutrons produced in the fusion reaction will not be affected by the magnetic field. Instead they will irradiate the drive and the entire spacecraft over the 100 year transit, and reduce the drive efficiency. Since this is a highly undesirable result, a reaction which produces few to no neutrons is required (see appendix). He3 + H2 yields no neutrons (although realistically some of the deuterium will react with itself producing a limited number of neutrons in each implosion). The problem is not solved, however, since there is not enough He3 on our planet to fuel the spacecraft! Three methods of gaining the necessary He3 have been compared: mining the planet Jupiter; creating He3 through the bombardment of Lithium in nuclear accelerators; and capturing He3 from the Solar Wind. Another possibility is for a further technological breakthrough to enable using higher threshold-energy fusion reactions (higher than H,He) which use more abundant elements in a no-neutron reaction. None of the options seem very reasonable, and each should be explored and further developed to determine the best method for collecting the necessary fusionable material.

The pellet size, in order to obtain the proper mass flow through the nozzle, depends upon the pulse frequency. Smaller pellet size could potentially lower the coil mass as well as the igniter mass, although the higher frequency would complicate fuel injection in a system that must run for 100 years continuously, without repair. The appendix shows the spectrum available between the DEADALUS pellet size and frequency (since DEADALUS required a higher mass flow).

After the final upper stage separation, the nuclear reactor will be increased to full power in order to charge the interstellar drive capacitors for initial ignition. The Interstellar drive will then be used for both acceleration and deceleration. The system is to be turned off at the appropriate time (determined through an internal navigational calculation), rotated 180 degrees, and restarted, all while staying on course. The payload contains a 300 kw nuclear power reactor which must be also capable of starting and restarting. The nuclear reactor will have to be ignited, and rechanneled to repower the slowly draining capacitors of the interstellar Drive igniter system, after the spacecraft has fully rotated and stabilized in the proper alignment.

3.2.2.3 Feasibility[edit]

The entire Interstellar Drive is highly dependent upon enabling technology. Building an actual scale model that is capable of running continuously for 100 years will be a challenge by itself! Barring further significant technological breakthroughs, the collection of fuel will be the most difficult and time consuming portion of the building. Never the less, within 20 years, these projects should be possible with the proper funding. Current technology is already capable of creating singular microexplosions in the laboratory.

3.3 Instrumentation[edit]

3.3.1 Instruments[edit]

The instruments to be carried on board the probe head are as follows:

  • IR Imagers
  • Visual Imagers
  • UV Telescopes
  • High-energy Particle Detectors
  • Astrometrical Telescopes
  • Wide-Band Spectrophotometers
  • Magnetometers
  • Solar Wind Plasma Analyzers
  • Communications Lasers

Three of each item will be carried on the spacecraft for triple redundancy.

The total weight of the instrumentation package (including everything listed, except the communications lasers) is estimated at not more than 3 metric tons. The estimated weight of the communications lasers is 2 metric tons. The peak power requirement of the instruments is estimated at 300 kW.

The justification for the visual imagers is obvious. Everyone will want to "see" what another star system looks like. In addition to providing important scientific information, a picture will be worth a thousand words or a thousand pages of numbers when it comes to obtaining funding for follow-up missions to other star systems.

IR Imagers and UV Telescopes will provide the first exact data on the characteristics of stars other than our own. Also, one must not forget the possibility that the Alpha Centauri system contains planets. These instruments could also provide data on the radiation and thermal environment of any planets in the system.

The High-Energy Particle Detectors are one of the few types of instruments which will be active during the transit to the objective system. Hard data on the energy level and density of such particles could provide insight into the origin and eventual fate of the universe.

The Astrometrical Telescopes will be the backbone of the mission. By providing the data to accurately determine the distance to the further stars, these instruments will advance the study of stellar characteristics immeasurably. The only limit to this aspect of the mission will be the endurance of the spacecraft. (The section on the objective system contains an explanation of how this will be accomplished.)

The Wide-Band Spectro-photometers will determine the composition of the stars and any planet sized bodies which the system may contain.

The Magnetometers will also be in use during the entire life of the probe. These instruments will provide extensive data on what should prove to be the very interesting magnetic field of a trinary star system. Also, they will provide the first hard data on the galactic magnetic field and how it interacts with the magnetic field and how it interacts with the magnetic fields of our own solar system and the and the Alpha Centauri system.

The Solar Wind Plasma Analyzers will also provide some scientific "firsts". While the composition of the Sun's solar wind is already known, these instruments will accurately determine how far the wind extends into inter-stellar space. This will also be done for the Centauri wind, as well as determining its composition. Also, the close binary pair of Alpha and Beta Centauri should exhibit a very interesting pattern where their solar winds interact.

The communications lasers, in addition to performing their obvious function, will at the same time provide data on extremely long-range laser communication and the extent of spreading losses caused by the interstellar medium.

3.3.2 Instrumentation Configuration[edit]

Two possible configurations for the mounting of the instruments were considered. Both consisted of three booms attached to the probe body spaced at 120 degree intervals with each boom supporting a complete instrument package, but one was dynamic and the other static.

The dynamic boom configuration was designed with the intent of retaining the forward particle shield as a structural and operational part of the probe body after orbit is achieved in the target system. The particle shield was to be infused with pipes to provide additional radiating area for waste heat. The shield could also be used to house elements of the probe's central processing units. The major advantages of the dynamic boom configuration are its additional cooling capability and the additional shielding which would be provided in-system for the body of the probe. The major disadvantage is that the mobility of the booms would have to be maintained for the length of the mission. This configuration would require movable mounts at the base of each boom capable of handling the large torques caused by moving the boom. These torques would also pose an additional problem for the probe's attitude control system. Furthermore, the advantage of a larger cooling surface would be offset by the added thermal control complications.

The static boom configuration, (see Fig. 3.3a) was designed to discard the front particle shield upon approach to the target system. The instrument booms will be firmly attached to the probe body and the only movable parts will be the individual instrument mountings. The major advantages of the static boom configuration are the lower number of parts expected to move after the long interstellar transit, and the lower final mass of the probe in-system. The major disadvantage of this design is the need for pyrotechnics to eject the shield after transit. All things considered, the static boom design was adopted, mostly its the greater reliability.

STATIC BOOM PROBE CONFIGURATION

Project Longshot - Design Report - Fig 3.3a.png

Fig. 3.3a

3.4 Communications System Design[edit]

The major challenges for the communications system of the interstellar probe both occur when the probe enters the target system at a range of 4.3 light-years, or 4.109 x 10^16 meters. This is the maximum transmission range; a fairly high data rate must be maintained, since all probe instrumentation is returning data. The only type of communications system capable of the necessary directivity and data rate is a high-power laser using pulse code modulation (PCM).

Low background noise from the target system is necessary for a low power level, so a laser wavelength of 0.532 microns was chosen. Radiation of this wavelength is almost totally absorbed by the outer atmospheres of K and G type stars, leaving a hole in the absorption spectrum (no transmitted radiation). Laser radiation of this wavelength can then be produced by a frequency-doubled diode-pumped YAg laser with an optical attachment to provide a large initial aperture.

The transmitter aperture is 2 meters in diameter with receiving mirrors of 24 meters diameter. The spreading angle is 1.22*lambda divided by the aperture diameter, or 3.25 x 10^-7 radians (0.067 arcseconds). At 4.34 light-years, the spreading results in a footprint radius of 13.4 million kilometers, 8.9% of an Astronomical Unit (AU). Both the pointing accuracy of the laser mount and the attitude determination capability of the probe must be within 0.067 arcseconds, so very low error laser mounts and star trackers will be used.

A total input power of 250 kilowatts is needed for each laser that is transmitting. With an assumption of a 20% lasing efficiency, the transmitted power is 50 kilowatts. If the power is distributed isotropically over an area of 5.64 x 10^20 square meters (the area subtended by the laser beam when it reaches Earth), the power density is 8.87 x 10^-17 watts per square meter, or 222 photons per square meter per second. For a 12 meter radius receiving mirror (area of 452.4 square meters), the received power level is 4.01 x 10^-14 watts, or 100,000 photons per second. Using the assumption that a digital pulse 'on' level is 100 photons, the receiver sees 1000 pulses per second. A data rate of 1000 bits per second is low. Note that this rate is the minimum because the transmitter would be at maximum range. If extremely reliable lasers are used, each transmitter can operate at a slightly different wavelength, so the data rate would be up to six times greater depending upon the number of lasers used.

The communications system would use six 250 kilowatt lasers. Three would be placed on the outside of the fuel tanks with the star trackers for communications during the acceleration phase. Three more lasers would be attached to the probe head for communications during the deceleration and in-system phases of the mission. The receiving mirrors would be in geosynchronous orbit about the earth in a constellation of several mirrors with a central node serving as a relay station to TDRSS and the ground.

3.5 Data Processing[edit]

One of the significant "enabling technologies" required to perform the interstellar probe mission is that of advanced processing. When the spacecraft reaches its target, the probe will be 4.3 light-years away from command and control facilities on Earth, and will thus have to be completely autonomous and self-repairing. The processing system will ideally have low power consumption to reduce heat dissipation requirements. It will be multiply redundant with advanced shielding and survivability features, and also be able to control all facets of probe operations, including high level decision-making. As shown in the sectional appendix, the probe must evaluate given mission objectives to control the scientific instruments in order to explore the Alpha Centauri system most effectively. If the system can integrate high-accuracy attitude determination and scientific data instantaneously, the attitude control requirement can be relaxed to a level easily maintained by such a large structure. Finally, the data must be taken out of processor's memory and sent to earth via laser. The communications lasers must be pointed with an accuracy of 0.67 arcseconds, and a hard file of the position of the earth relative to Sol must be retained in memory to govern the pointing of the laser. Once the target system is reached, the processing unit must be able to achieve and maintain an acceptable orbit, and maneuver the investigate high priority phenomena, such as evidence of intelligent life. For a block diagram of the data processing system, see appendix.

3.6 Guidance[edit]

A system of star trackers will be used for both navigation and attitude determination. This system has been chosen for its high accuracy (4 arc sec) and adaptability, and for its low weight (7 kg) and power requirements (18 watts). Star scanners were not chosen since the spacecraft is not rotating, and they are less accurate. The trackers will be coupled to a computer system which will have a star catalogue of 200-300 stars' locations. "Adaptability" refers to this catalogue, because a "best guess" of star locations that the probe will "see" on the trip and in orbit in the Centauri system can be programmed into the computer before launch, thus increasing the accuracy of position/attitude determination. This best guess could even be updated enroute or in Centauri-orbit by the astrometry calculations that the computer will make.

Initially, during the transit between Earth and the point where the probe-head separates from the propulsion system, trackers located on the last fuel tanks will be used. These trackers will be oriented in different directions in order to gain a nearly complete field of view. 18 trackers will be used, 3 in each of the 6 axis directions to get the greatest field of view and triple redundancy. The power for these trackers will come from generators drawing energy from the propulsion system waste heat. In the final mission phase, 9 trackers located on the instrument booms will be used, 3 on each boom instrument head. (This will require a very accurate position determination of the boom rotation angles.)

Three 3-axis rate-gyro assemblies will determine the rate of change of any two pointing angles and the spacecraft roll rate. This data will supplement the trackers' information and increase the attitude determination accuracy.

Star tracker parameters:

  • Solid state (vice photomultiplier tube)
  • 4 arcsecond accuracy (future improvement is expected)
  • Magnitude range −1 to +8
  • Field of view 6 × 6 degrees
  • 6 seconds to search field of view
  • 1.2 seconds to search field in 1 × 1 degree search mode
  • Has a track mode during which it follows a specific star
  • Total weight — 189 kg
  • Total power in transit — 324 watts
  • Total power after probe separation — 162 watts
A summary of the attitude control systems available to choose from are listed in Fig. 3.6a. The probe's attitude control will be accomplished using two sets of flywheels arranged on 4 axes (described in section 3.3 Power System), and an auxiliary system of hydrazine thrusters. These flywheels will serve as momentum wheels, controlled by the computer using the attitude/rate information, providing torque to maintain spacecraft stability. The 4-axis configuration will enable the reaction wheels to absorb external torques from any direction. The magnitude of this reaction torque is easily modulated by electronic control of the reaction wheel motor current. One disadvantage of this system is the need to control wheel speeds in order to limit vibrational effects. "Unloading" the energy of

ATTITUDE CONTROL TECHNIQUE COMPARISON

TYPE COST ACCURACY MISSIONS COMMENTS
3-AXIS ACTIVE CONTROL
(1) REACTION WHEELS
VERY HIGH .007 ARC-SECONDS–1.0 DEG. ASTRONOMICAL, WEATHER VIBRATIONS AT HIGH RPM'S MAY REDUCE ACCURACY
(2) HYDRAZINE THRUSTERS
VERY HIGH .1–1.0 DEG DEEP SPACE FUEL LIMITED
SPIN STABILISATION LOW TO MODERATE .1–2.0 DEG. EARTH ORB, INTERPLANETARY SPIN RATE AND DIRECTION REQUIRE CONTROL
DUAL-SPIN STABILIASATION
(1) HALF SPIN, HALF STABILIZED
HIGH .01–.1 DEG. GEOSYNCH. COMMS. SATELLITES REQUIRES COMPLEX ELEC. AND MECH. CONNECTIONS
(2) INTERNAL MOMENTUM WHEELS
HIGH .01–.1 DEG. GEOSYNCH. COMMS. SATELLITES HIGH TECHNOLOGY REQUIREMENTS
MAGNETIC STABILIZATION VERY LOW 10–20 DEG. LOW ALT., SCIENTIFIC REQUIRES A MAGNETIC FIELD
GRAVITY-GRADIENT STABILIZATION LOW 1.0–10 DEG. CIRCULAR, LOW ALT. REQUIRES A GRAVITY FIELD AND A LARGE MOMENT OF INTERTIA FOR SATELLITE
the wheels is done by transferring momentum to the second set of wheels, discharging energy through the power system, and using the hydrazine thrusters. The hydrazine tanks are located within the main truss between the flywheels, encircled by the main fuel tanks (so they may be cooled by the same refrigeration units). The nozzles are located on the circumference of the spacecraft pointed in each axis direction. The flywheels will be used for attitude control in the star system phase of the mission, and in the Centauri system. The hydrazine system is primarily used as a backup.

3.7 Thermal Control[edit]

There are two significant problems to consider in the thermal design:

  • There is a huge amount of waste heat from the propulsion system and the nuclear reactor (both fission and fusion reactions);
  • The fuel will be stored in large tanks at near-absolute temperatures, and must be shielded from the waste heat of the nuclear reactions.

The spacecraft will require highly efficient radiators to dissipate the thermal energy released by the interstellar drive, namely the inductors, particle beams, and fusion reaction. The radiation from this dissipation process will be reflected away from the rest of the spacecraft by a mirror specifically engineered to reflect infra-red energy. Additionally, conduction will be buffered by special ceramic materials between the power and propulsion units.

The nuclear reactor will dump its waste heat to the same radiators used by the propulsion unit. Ceramic buffers will also be located between the power unit and the fuel tanks.

The fuel tanks will contain pelletized helium and deuterium which must be shielded from conductive and radiative heat energy. During the initial phase of the mission the tanks will be shielded from the sun by a shroud which will be blown off at a sufficient distance from the sun where the solar radiation becomes negligible. The mirrors and ceramic buffers will keep the spacecraft's waste heat away from the tanks, while refrigeration units will keep the fuel vapor pressure low enough to remain in pelletized form for the duration of the mission. The tanks will be painted black to emit as much radiation as possible in deep space.

The probe head will be protected from the radiation from Beta in the Centauri system by a thermal blanket. Heat from the computer, instruments, and lasers will be convected through heat pipes to the radiators at the rear of the spacecraft.