Introduction

The Icarus probe’s main purpose will be to communicate its findings back to Earth. To do so, it needs to be able to generate a powerful signal, and point it accurately at the Earth using a directional antenna. The probe might also benefit, for the first few years, from two-way communication with Earth. This would allow software upgrades and patches to be sent to continuously upgrade the probe’s computer software, and even some ship hardware using the on-board fabrication facilities, during its trip.

The original Daedalus study proposed using the probe’s second stage nozzle as an antenna dish. However, this proposal both constrained the operation as a nozzle and limited the maximum aperture. This was considered acceptable because the Daedalus probe was moving, at target, at high velocity. However, the Icarus Interstellar probe will be stationed at the target system, with plenty of time to deploy or even build an external antenna. Therefore, the proposition is to deploy a large antenna composed of self-assembling swarms or built by Spiderfabs [1] allowing for high bandwidth communication and lower transmitter power.

Radio waves are attenuated in proportion to the square of the distance. The distance to the target system, Alpha Centauri, at 4.37 light years, is a factor 10⁹ greater than the distance between a geosynchronous satellite and the Earth’s surface. As a result, the attenuation will be 10¹⁸ greater, and transmitter powers and antenna diameters must be much larger in order to cope with that increased distance. The required transmitter powers and antenna diameters also depend on the data rate. The target data rate, used for this chapter, between the Icarus probe and Earth is 20 Gbps, the equivalent of 13 high definition TV channels (at 1.5 Gbps each). This will enable the real-time, data-rich exploration of the target worlds. It is also around 1000 times greater than was planned for Daedalus. Some of the probe designers may choose different rates as explained in the ship sections.

Attenuation increases with frequency, but antenna gains also increase with frequency (for the same diameter) and, since there is an antenna at each end of the communications link, there is an advantage to choosing a high frequency. A high frequency also allows for the use of a wide bandwidth, which assists in supporting a high data rate.

An alternative would be to use a laser to support this communications link. In principle, the higher frequency would allow a higher data rate to be transmitted over the link. The possibilities for employing this technology are explored after first considering a Radio Frequency (RF) link.

TRANSMITTER TECHNOLOGIES

Figure 1 illustrates the major classifications of transmitter technologies and their capabilities. For maximum transmitter power at the highest frequency, the choice must be Gyrotron.

Figure 1, transmitting technologies [2]

Current communications spacecraft utilize Helix TWTs because these offer good efficiency at power levels on the order of 100 W, at microwave frequencies. Solid state amplifiers are also flown on spacecraft, but they tend to be available at the lower frequencies.

The capabilities of the different transmitter technologies are determined by a number of factors, including heat dissipation and voltage breakdown. The dimensions of the RF structures scale inversely with frequency, and the maximum power that can be handled by a particular technology is limited by the maximum temperature that the internal surfaces of the cavity can safely reach. As the structures become smaller, their heat dissipation capabilities are reduced. Gyrotrons can handle higher power levels than Klystrons because they operate with higher order RF modes and are, therefore, relatively large for the same frequency.

Figure 1 indicates that Gyrotron devices can handle power levels of around 1 MW, but there are a number of different technologies within the descriptive term Gyrotron. The type of Gyrotron being developed as an amplifier is the Gyro-TWT, because of its high gain and wide bandwidth characteristics. However, the Gyro-TWT also suffers from instabilities, in particular at high beam currents. These can be controlled by introducing distributed losses into the device and by modifying the internal structure of the device. For example, one development has involved the use of a helically corrugated interaction structure, resulting in an output power of 1.1 MW although not at as high a frequency as selected for this application. [8] Further development work is required before suitable amplifiers will be available for Icarus.

However, in order to maximize the power and frequency, Gyrotron technology is most appropriate for the communications link between Icarus and Earth, and that will potentially provide a transmitter output power of the order of 1 MW at frequencies up to 100 GHz. An RF frequency of 32 GHz has been selected for this application, because it is already allocated for space research, thus minimizing any potential interference.

For wide-band RF equipment, such as amplifiers, the maximum usable bandwidth is normally considered to be 10 % of the frequency. For the purpose of this analysis, a bandwidth of 3 GHz has been chosen.

TWT, Klystron and solid state technologies are appropriate for other communications links within this mission, and those are detailed when considering the Communications Subsystems.

LINK BUDGET CALCULATIONS

Icarus to Earth Communications Link

With the data rate, transmitter power and RF frequency chosen, the link budget calculations can be simplified to (mainly) a choice of antenna diameters.

Transmission Antenna

A large antenna cannot be installed on the Icarus probe for use throughout the mission. During the acceleration and deceleration phases the stresses may be too great for the antenna to survive, and during the coast phase (at high velocity) impacts from interstellar dust could destroy the antenna structure. So separate link budgets for each phase of the mission have been constructed, employing antenna diameters suitable for that phase.

The target data rate will also be reduced for the early phases of the mission, since only the en-route science observations would need to be transmitted. Uplinking to the Icarus probe would also not require high data rates, since real-time commanding is not foreseen.

The maximum antenna diameter for the Icarus probe, once on station, is 1 km. This antenna would have to be fabricated after the probe arrives. Tethers Unlimited has already proposed a design for a large parabolic reflector in the 1 km range using their Spiderfab technology [1]. This development can be applied directly to the construction of a reflector dish at Alpha Centauri. The reflector is composed of a hoop structure constructed of carbon composite truss elements, a reflective mesh spread out inside the hoop and a network of tension lines that enforce the correct parabolic shape on the mesh. Figures 2, 3 and 4 illustrate this concept.

Figure 4 Dish Mass vs Dish Diameter [1]

There would be some development required beyond the parameters used by Tethers Unlimited, as we propose to change the wavelength from 30 cm (1 GHz) to 1 cm (32 GHz). This would require more precise fabrication to ensure that the signal is not degraded.

The mass of a 1 km reflector using the construction methods proposed by Tethers Unlimited is 40 tonnes. Two to three Spiderfab robots working in parallel could construct the antenna in six months. The material would come from Earth on the Icarus probe, and therefore take up a significant part of the payload. The payoff would be the very high transmission rates described in this paper.

Receive Antenna

At the Earth end of the communications link, an off-planet communications gateway is required, in order to maintain the antenna pointing towards the Icarus probe. If the antenna were to be on the Earth’s surface, it would rotate with the Earth and not support a continuous communications link. Even if multiple antennas were to be used, they would need to be steerable and that would become hugely problematic with the necessary antenna sizes. In order to ensure minimal blocking of the line of sight to the Icarus probe, this analysis has assumed that the gateway would be located at either the L4 or L5 Lagrange point of the Earth-Moon system. Then the maximum antenna diameter is constrained only by what would be reasonable to fabricate, at that distance from Earth and in the timescales of the mission.

In fact, the receive antenna for the gateway could be created from an array of smaller antennas, with additional antennas added as the mission continues. The effective area of the combined antenna would be the same as the additive areas of the individual antennas. The signals from the individual antennas could be combined after amplification. The required size of the receive antenna can also be reduced, for the on-station phase, by transmitting two carriers from the Icarus probe, each on the opposite polarization to the other, and splitting the data to be transmitted between them. For the gateway transmit function, however, a single antenna must be employed because combining signals at high power, whilst maintaining the phase relationship of those signals, is much more difficult. But, this transmit antenna need not be as large as the receive antenna since the target data rate is less.

The Gateway. A large antenna farm in high Earth Orbit. The central antenna, 3 km wide, provides the area required for signal transmission to the Icarus probe, but can also be used as part of the receiver. (ML)

It will be necessary to ensure that the two antennas are pointing accurately at each other. This can be accomplished if each antenna system is equipped with a monopulse tracking system, using the carrier transmitted by the other antenna as its source. The beam from each antenna should be wide, in comparison with the short-term movement of the other antenna, and so little variation in received carrier power should be expected.

Thermal Noise

In addition to the reduction in carrier power, over the distance between Icarus and Earth, the received signal will be degraded by thermal noise in the receiving system. This noise results from the thermal environment of the receiving equipment, including the antenna. For this application, however, the antennas will be pointing towards deep space and so their contribution to the overall receive noise temperature will be minimal. A contribution of 50 K from the antenna (in each case) has been assumed. For the receiving equipment, an environmental temperature of 250 K has been assumed. This is typically the minimum temperature for space-borne equipment while operating in Earth orbit, and the temperature at which components are qualified for use in space. The use of components at lower temperatures may have been qualified by the time of the Icarus mission, which will allow this environmental temperature to be reduced, but that may have consequences for the thermal control subsystem of the spacecraft. A noise figure of 3.1 dB has been assumed for the initial amplifier of the receiving equipment. This is typical of the current performance for space-borne Ka-Band low noise amplifiers.

Transmission Through the Ion Plume

One potential problem for this communications link occurs during the boost phase. The plasma ejected to provide the thrust may interfere with the communications link. The interference is caused by the free electrons in the plasma, which can absorb the RF energy depending on their density and range of movement. There is a cut-off frequency, below which RF signals are (effectively) blocked, which varies with the electron density. Re-phrased, that means that there is a critical electron density which must not be exceeded if a particular RF frequency is to pass through the plasma.

That critical electron density may be calculated from the formula; N= (4π2f2εome)/e2, resulting in a critical density of 1.11 x 10¹⁷ electrons/m³ at a frequency of 30 GHz.

The density of electrons emitted from the Icarus spacecraft will depend on the specific design being considered. For the Firefly design, the plume will form a cone with divergence 45^o, but the electrons will detach from the magnetic field lines (of the nozzle) and recombine with the positive ions after a certain distance from the spacecraft. The working assumption is that the electrons will detach at a distance twice the radius of the nozzle, forming a disc 50 m deep and 50 m radius, 50 m from the spacecraft. The electron flux is 1.5 x 10²⁵ electrons/second, with a speed of 12,000 km/s. The electron density within that disc will then be 1.6 x 10¹⁴ electrons/m³.

So, the electron density produced by the Firefly spacecraft is three orders of magnitude less than the critical density. The fact that Firefly uses a magnetic nozzle is also advantageous, because studies have shown that applying a magnetic field to a plasma creates windows through which RF signals can pass, even if they are below the cut-off frequency for that plasma [5]. More work is required in this area, to examine the specific impact of the spacecraft plasma plume on the communications links, but this initial assessment suggests that, at the frequencies assumed for the Icarus study, there should be no problem.

Link Budget Results

The achievable data rate is calculated from the Shannon-Hartley theorem [6], which determines the maximum rate at which information can be transmitted over a communications channel in the presence of noise, with an adjustment for the practical implementation (the Tracking and Modulation loss).

Table 1 illustrates the resulting on-station link budget. In order to achieve the target 20 Gbps data rate from the Icarus probe to Earth, using two carriers at 10 Gbps each, the equivalent of a 15 km diameter receive antenna is required. For a 1 Gbps uplink signal, a 3 km diameter transmit antenna is required. This transmit antenna must then be employed throughout the mission. The values used for some parameters have been based on realistic performance of current equipment designs and a number of assumptions. However, margin has been included (for example the maximum output power of a Gyrotron amplifier is well above the 1 MW that has been considered) and should be sufficient to allow adjustments when a detailed design is performed.

Table 1 On-Station Link Budget

Downlink to Earth

Uplink from Earth

Units

Frequency

32.0

34.5

GHz

Wavelength

0.009368514

0.008689636

m

Vehicle Antenna Diameter

1000

1000

m

Efficiency

0.5

0.5

Vehicle Antenna Gain

107.50

108.15

dB

Gateway Antenna Diameter

15.0

3.0

km

Gateway Antenna Area

176.7

7.1

km²

Efficiency

0.5

0.5

Gateway Antenna Gain

131.0

117.7

dB

Distance

4.37

4.37

ly

4.13141E+16

4.13141E+16

m

Path Loss

394.9

395.5

dB

Transmitter/HPA OP Power

1.0

0.8

MW

60.0

59.0

dBW

Received Carrier Power at Antenna

-227.4

-218.8

dBW

Typical Ka Transponder Noise

3.1

3.1

dB

2.0

2.0

Environmental Temperature

250

250

K

Transponder Noise Temperature

260

260

K

Antenna Noise Temperature

50

50

K

Total Noise Temperature

310

310

K

Receiving G/T

106.1

83.2

dB/K

Received C/N_o

107.3

93.0

dBHz

Channel Bandwidth

3000

500

MHz

C/N

12.6

6.0

dB

Tracking & Modulation Loss

1.1

1.1

dB

Received C/N

11.5

4.9

dB

13.98264184

3.120532396

Theoretical Max Channel Capacity

11,715,660,379

1,021,415,377

bps

11,716

1,021

Mbps

Target Data Rate

10

1

Gbps

Performance Ratio

1.17

1.02

Sensitivity to the link budget parameters can be judged from Table 2.

Table 2 Effect of Varying Link Budget Parameters

Uplink

Gateway Antenna Diameter

km

3

3

3

3

3

3

3

3

Gateway Transmitter Power

MW

0.8

0.8

0.8

0.8

0.8

0.8

0.8

0.8

Icarus Antenna Diameter

km

1

0.75

0.5

0.25

1

1

1

1

Data rate

Gbps

1.021

0.731

0.416

0.129

1.021

1.021

1.021

1.021

Downlink

Gateway Antenna Diameter

km

15

15

15

15

15

15

15

15

Icarus Antenna Diameter

km

1

0.75

0.5

0.25

1

1

1

1

Icarus Transmitter Power

MW

1

1

1

1

1

0.75

0.5

0.25

Data rate

Gbps

11.716

9.444

6.506

2.718

11.716

10.566

8.995

6.506

Similar link budgets can be constructed for the other mission phases, concluding with the data rate capacities and antenna diameters illustrated in Table 3. In each case, the link budget is calculated for the maximum distance for that phase:

• 0.2 light years, for the Boost phase,
• 1.61 light years for the initial coast phase,
• 3.02 light years for the second coast phase, and
• 4.37 light years for the deceleration phase.

Since the antenna for the deceleration phase is sized for a maximum distance of 4.37 light years, it can continue to be used (with the data rates applicable to the deceleration phase) once the spacecraft is on station, and while the main communications antenna system is being constructed.

The Gateway receive antenna diameter increases as the mission duration increases, compensating for the increased path loss for the downlink signal. There is some modest increase in the antenna diameter on board the Icarus probe, once the boost phase has been completed, but the increased path loss for the uplink signal is compensated, mainly, by increasing the transmitter output power at the Gateway. In all cases, the target data rates are exceeded.

The link budget is a multi-dimensional problem, in that (at the simplest level) three parameters can be varied independently to achieve the desired data rates. These are the transmit and receive antenna diameters and the transmitter output power. Table 2 provides an indication of the achievable data rates, as either the transmit antenna diameter or the transmitter output power values are reduced.

Intra-System Communications Links

Several different communications links will be employed to support the exploration of the target star system. At the target star system, communication requirements between the probes are orders of magnitude lower than for the Icarus-to-Earth link [3]. The recent successes of New Horizons at Pluto and Ceres, and the continuous operation of Voyager, shows that this is already possible at the largest possible scale in the solar system. However, the power supply from a large probe like the Icarus fusion starship allows for data rates unseen until now, above the Earth’s atmosphere.

The target star system, Alpha Centauri, is a binary system. As a result, one concept is for there to be two communications hubs; one in orbit around each of those two stars. This minimizes the link budget requirements for the sub-probes associated with each star. Those hubs are designated Icarus-A and Icarus-B, with Icarus-A providing the communications link to Earth and Icarus-B communicating with Earth via Icarus-A. An alternative arrangement, with a single communications hub, remains possible with a different set of link budgets.

In addition to the communications link between the two communications hubs, four distinct communications requirements are foreseen. These are:

• An orbiter, communicating with either Icarus A or Icarus B,
• A planetary probe (a rover or a glider), communicating with an orbiter,
• A stellar probe, communicating with either Icarus A or Icarus B, and
• A flyby probe to Proxima Centauri, communicating with Icarus A.

Link budgets for these four cases can be constructed in the same manner as for the main mission, Icarus to Earth, communications link. For the case of a planetary probe communicating with an orbiter, however, an additional element must be added to the link budget to take into account any atmosphere. No information is available regarding the composition of any atmosphere, so (for the purpose of this study) an Earth-like atmosphere has been assumed. As a result, a different frequency at around 8 GHz has been chosen to reduce atmospheric attenuation and rain fades, but also to minimize the diameter of the antennas. An allowance of 10 dB has been included for rain fades.

The achieved data rates from these link budgets are illustrated in Table 4. (In the case of the Icarus-B to Icarus-A link, the term “probe” corresponds to Icarus-B.) The target data rates are exceeded in all cases.

Gateway to Earth Communications Link

The Earth gateway for the communications link to the Icarus spacecraft cannot be located on the Earth’s surface since that is rotating and a constant link would not be possible. For the same reason, it cannot also be located on the surface of the Moon. If it were located in orbit around the Earth, but within the Moon’s orbit, then there would be the potential for blocking of the link by the Moon. As a result, in order to minimize blockages, the gateway location has been assumed to be at either the L4 or L5 Lagrange point of the Earth-Moon system. As a result, the distance from Earth to the gateway is the same as the Earth to Moon distance.

The link budget for this communications link is constructed in the same manner as for the other communications links. However, it requires the inclusion of atmospheric losses and a rain fade margin, since the link passes through the Earth’s atmosphere. The environmental temperature for the gateway RF system can be kept cold, as for the environment onboard the Icarus spacecraft, but “room temperature” is appropriate for the equipment, and the antenna, actually on Earth. The antenna noise temperature for the gateway antenna supporting this link can remain cold, however. Although that antenna is pointing at Earth, its field of view will also include deep space.

Table 5 illustrates the results of this link budget, confirming that the required data rates can be achieved with a gateway antenna diameter of 30 m and an Earth antenna diameter of 15 m. In reality, a number of ground stations on the surface of the Earth will be required, to maintain constant communications with the gateway at the Lagrange Point. An antenna diameter of 15 m is consistent with current antennas used for satellite communications hub stations.

Table 3 Antenna Diameters and Transmitter Powers for Mission Phases

Proxima Centauri Flyby

Included within the Icarus mission plan is exploration of Proxima Centauri. This star is a considerable distance from the Alpha Centauri system, but may be gravitationally linked to it. Whilst it is closer to Earth, Proxima Centauri is not on the flight path between Earth and Alpha Centauri. The intention is to deploy a probe from the main Icarus spacecraft during its cruise phase but, as a relatively simple probe, this will offer only a flyby opportunity for exploration.

Since the probe will be moving at high velocity, it will not be possible to deploy a large antenna, since that would not be protected by the probe’s dust shield. In consequence, the data link with this probe cannot be established until after the Icarus spacecraft arrives at Alpha Centauri and has deployed at least part of its antenna array. Although the probe will continue to move at high velocity, at the distance between Alpha Centauri and Proxima Centauri, the angular rate of change will remain small, allowing antenna tracking.

DATA LINK REQUIREMENTS

As a reference point, the following assumptions have been made about the numbers of data links and their capacities:

• 1 data link between Icarus A and Earth, at 20 Gbps
• 1 data link between Icarus B and Icarus A, at 10 Gbps
• 3 data links between planetary orbiters and Icarus A, each at up to 3 Gbps
  • 6 concurrent data links between planetary probes/rovers/gliders and Icarus A, via the three orbiters, each at 1.5 Gbps
• 1 data link between a stellar probe and Icarus A, at 1 Gbps
• 3 data links between planetary orbiters and Icarus B, each at up to 3 Gbps
  • 6 concurrent data links between planetary probes/rovers/gliders and Icarus B, via the three orbiters, each at 1.5 Gbps
• 1 data link between a stellar probe and Icarus B, at 1 Gbps

The actual number of planetary probes may be higher, but not all will have visibility of the appropriate orbiter at the same time and not all may be operational at the same time. Alternatively, a greater number of probes may transmit at reduced data rates, perhaps with buffering, when capacity limits are reached. Nominally, one orbiter will be assigned to each planet.

A data link is also required between the Proxima Centauri flyby probe and Icarus A. It is assumed that the flyby probe will transmit its data before the probes are deployed within the Alpha Centauri system. If data were to be transmitted concurrently with data from the Alpha Centauri system, however, it would result in only a small constraint on the data rates available for the planetary probes. The capacity of this data link is constrained by the antenna diameters on each vehicle, both of which will be travelling at relativistic velocities. The calculated link budget shows that a data rate of 100 Mbps is possible.

COMMUNICATIONS SUBSYSTEMS AND GATEWAY

Icarus Spacecraft

Icarus-A Link to Earth

From the link budget in section 3.1, a 1 MW transmitter is required at the Icarus spacecraft, once it has arrived at the target star system. This necessitates the use of Gyrotron technology for the high power amplifier (HPA). The same amplifier can be used during the earlier mission phases, with backed-off output power. Although a Gyrotron amplifier can support the required output power, its gain is less than 40 dB and so a pre-amplifier will also be required to produce an appropriate input to that HPA, at up to 1 kW. That implies (from Figure 1) that the input to the Gyrotron amplifier must be fed from a Klystron amplifier.

Also from the link budget in section 3.1, it will be necessary to transmit two carriers (each 1 MW) from the Icarus spacecraft to Earth, in order to limit the required diameter of the gateway antenna. The consequence of this is that there must be two active HPAs at the Icarus spacecraft, each with its own, independent input since the carrier frequencies will be the same. There will be on-board maintenance facilities, but a standby HPA will be necessary to ensure ongoing transmission while any repairs are carried out. Therefore, the transmitter section of the Icarus communications link to Earth must consist of three amplifier chains, two of which are active at any time once on station.

A single carrier will be transmitted from Earth to Icarus, and so the receive section requires only one active low noise amplifier (LNA) plus a standby unit.

For both transmit and receive sections, the RF loss between that equipment and the antenna must be minimized. That implies that the distance to the antenna must be minimized, and all RF equipment should be mounted close to the antenna feed. Since the link budget requires an antenna diameter of 1 km, it is not practical to mount this antenna on the body of the Icarus spacecraft and so a remote antenna with a power umbilical is the solution. Telemetry data from the Icarus spacecraft would also be carried by that umbilical, plus the scientific data if the data pre-processing is performed on-board the Icarus spacecraft.

The simplest antenna design would be a dual reflector, such as the Cassegrain, in which the feed is located at the centre of the main parabolic reflector with a convex sub-reflector in front of the feed. This would allow the RF equipment to be mounted on the rear of the main reflector, close to the antenna feed. That sub-reflector creates a small blockage of the signal, however. Some increase in gain can be achieved through the use of an “offset” antenna design, in which the main reflector is not a symmetrical section of the parabolic surface, and its focus and the secondary reflector are located to one side of the main reflector, removing the blockage. The feed can remain mounted at the edge of the main reflector, allowing the RF equipment to be attached to that structure.

It has been proposed that this antenna could be constructed by self-assembling swarms or built by Spiderfabs [1]. In any event, care must be taken to ensure that the surface accuracy of the main and sub reflectors is sufficient. The typical rule of thumb is that the surface profile should be within 10 % of the wavelength of the theoretical profile. At a frequency of 32 GHz, that 10 % of the wavelength is .0009 m, ie less than 1 mm over a surface 1 km in diameter! As the surface accuracy decreases, the antenna gain will also decrease. If the surface accuracy cannot be achieved then a larger diameter antenna will be required at one or both ends of the communications link. Ruze’s equation [4], can be used to determine the degradation of the antenna gain as the surface roughness increases. The gain decreases in proportion to the square of the ratio of the roughness to the wavelength, so a doubling of the roughness (to 2 mm) would decrease the gain by 6 dB and decrease the achievable data rate by a factor of approximately 4.

The mounting of the RF equipment, and the HPAs in particular, on the rear of the main reflector will also have some impact on the reflector’s surface accuracy as a result of the heat dissipated from that equipment. The thermal gradient imposed by that dissipation will cause some distortion of the surface, unless sufficient heat can radiated away. That thermal control will also be necessary in order to maintain the LNAs at a cold temperature.

The efficiency of Gyrotron amplifiers has been improving, and some reports indicate an efficiency of up to 60 % has been achieved. However, the cases considered have been for pulsed operation. For the purpose of this study and for continuous operation, an efficiency of 40 % is assumed. That results in a total power consumption of 2.5 MW and a heat dissipation of 1.5 MW for each active amplifier. Each Klystron pre-amplifier could have a similar efficiency, perhaps up to 50 %, resulting in a power consumption of 2.5 kW and a heat dissipation of 1.5 kW. The other equipment will have power consumption and heat dissipation values several orders of magnitude less, and so (to a first approximation) the total power consumption, for the pair of amplifier chains, will be around 5 MW with a heat dissipation of around 3 MW.

The mass of a typical Gyrotron amplifier is around 1,000 kg. A Klystron amplifier would have a mass of around 15 kg. The mass for a 1 km aperture antenna, built by Spiderfabs, could have a mass of around 40,000 kg. The aggregate mass for the complete sub-system is likely to be in the region of 44,000 kg, including cables and waveguides, etc.

Figure 5 provides a basic illustration of the RF sub-system at the Icarus-A spacecraft, for the communications link with Earth.

The two Icarus spacecraft would be separated by 30 AU, so the link budget does not require the transmitter powers or antenna diameters of the communications link to Earth. The data rate from Icarus-B to Icarus-A would be 10 Gbps. For each spacecraft, the antenna diameter has been selected to be 75 m and the Icarus-B transmitter power is 60 kW. The Icarus-A transmitter power is 3 kW.

For both spacecraft it should be possible to use Klystron HPAs. In both cases a single carrier is to be transmitted, so a single active amplifier, only, is required plus a standby amplifier. As a result, the total mass for each spacecraft would be around 50 kg. For Icarus-A, the power consumption would be 7.5 kW, with a heat dissipation of 4.5 kW. For Icarus-B, the power consumption would be 150 kW, with a heat dissipation of 90 kW.

Each 75 m diameter antenna would have a mass of around 70 kg, plus 50 kg for the RF equipment.

Icarus Spacecraft Links to Probes

Each Icarus spacecraft will support communications links to three orbiters, orbiting an object of interest, assumed to be a planet, plus a stellar probe. The distance between an Icarus spacecraft and an orbiter would be 6 AU, while the distance to the stellar probe would be 3 AU. The orbiters would transmit at 3 Gbps to the nearer Icarus spacecraft, and the stellar probe would transmit at. However, the requirement for each Icarus spacecraft would be less stringent, at 100 Mbps to each of these vehicles.

On board each Icarus spacecraft, a 50 m diameter antenna would be required to support the link to each of the probes. For the 100 Mbps links to the orbiters and stellar probes, a transmitter power of 0.5 kW and 0.1 kW, respectively, would be sufficient. A Klystron HPA would be appropriate for the link to the orbiters, but a TWTA could be used for the link to the stellar probe. Each antenna should be supported by its own pair of HPAs.

For the four links, from an Icarus spacecraft, the total power consumption would be 2.4 kW, with a heat dissipation of 0.8 kW.

Each antenna mass would probably be similar to that for the 75 m antenna, at 70 kg, plus the same 50 kg mass for RF equipment as noted for the Icarus-B to Icarus-A communications link.

Proxima Centauri Flyby

The maximum antenna diameter for the flyby probe has been considered to be 25 m, and the required data rate is 100 Mbps. For a 200 m diameter antenna at the main Icarus spacecraft, the probe’s transmitter power would need to be 750 kW. This would require a Gyrotron high power amplifier, with Klystron pre-amplifier.

The total power consumption on the probe would, then, be around 2 MW (including all other RF equipment) with a heat dissipation of 1.125 MW. For a 25 m diameter antenna the mass is estimated to be 80 kg, and the mass of a single Gyrotron would around 1,000 kg, so the total mass of the communications subsystem would be approximately 2,200 kg (assuming a redundant spare amplifier chain is also installed).

At the main Icarus spacecraft, a 200 m diameter antenna is used, with a mass of approximately 900 kg. A 75 kW Klystron transmitter is sufficient to support a 10 Mbps link to the probe. The mass for this communications subsystem would, then, be no more than around 1,000 kg (including a redundant spare Klystron amplifier).

Aggregate Constraints for Icarus Spacecraft

The antenna array at Icarus A will comprise the 1 km antenna for the data link to Earth, four 50 m antennas for data links with the orbiters and stellar probe, and one 75 m antenna for the data link with Icarus B. It will also include a 200 m antenna for the data link to the Proxima Centauri flyby probe. The antenna array at Icarus B will comprise four 50 m antennas for data links with the orbiters and stellar probe, and one 75 m antenna for the data link with Icarus A.

The aggregate mass, power and dissipation values for each, are as follows:

Link

Icarus-A

Icarus-B

Mass (kg)

Power (kW)

Dissipation (kW)

Mass (kg)

Power (kW)

Dissipation (kW)

Earth

44,000

5,000

3,000

A-B

120

7.5

4.5

120

150

90

Alpha Centauri Probes

480

2.4

0.8

480

2.4

0.8

Proxima Centauri Probe

1000

187.5

112.5

Total

45,600

5197.4

3117.8

600

152.4

90.8

Orbiters and Stellar Probe

A subprobe orbiter deploying a small lander at an hypothetical Alpha Centauri planet. The large 13 m antenna communicates back to the Icarus probe, the smaller one will track the lander. (ML)

Each orbiter must support a communications link to the nearer Icarus spacecraft. In addition, it must support two concurrent links to planetary probes. The stellar probe must support only a link to the nearer Icarus spacecraft.

It has been assumed that the structural designs for the orbiters and stellar probes will be at least similar. As a result, the antenna diameter for the link to the Icarus spacecraft has been assumed to be identical, at 13 m.

For the orbiters, the link to the Icarus spacecraft will be at a rate of 3 Gbps. That requires a transmitter power of 20 kW. A Gyrotron HPA will be required at the chosen frequency. A lower carrier frequency could be employed, which could enable an alternative HPA technology to be employed, but that would limit the possible modulation bandwidth and hence the data rate that could be supported. In addition, two links to planetary probes are to be supported. For transmission from the orbiter the target data rate is 100 Mbps, requiring a transmitter power of 50 W and a 1.5 m diameter antenna. So, the orbiters have three antennas installed (one 13 m and two 1.5 m) and three pairs of high power amplifiers. The power consumption and heat dissipation of the Gyrotron will dominate, at 40 kW and 20 kW, respectively.

The link from the stellar probes is at a rate of 1 Gbps, requiring a transmitter power of 2 kW. This is within the capability of a Klystron amplifier. A pair of amplifiers is necessary, as before, with a total power consumption of 3.3 kW and a heat dissipation of 1.3 kW.

Planetary Probes

Each planetary probe supports a data link to an orbiter, at 1.5 Gbps. A planetary atmosphere has been assumed but, since no information is currently available, this is assumed to be earth-like. The carrier frequency has been chosen to be 8 GHz, to avoid rain attenuation whilst also maximizing available bandwidth.

In conjunction with the 1.5 m antenna on the orbiter, this data link requires a 0.3 m diameter antenna on the probe, or an equivalent phased array, coupled with a 20 W transmitter output power. A solid state amplifier should be sufficient for this purpose. Typically, the mass for the RF equipment would be less than 10 kg, the power requirement would be of the order of 130 W, and the heat dissipation would be around 110 W.

Gateway

The purpose of the Gateway is to receive the downlink signals from Icarus-A, and relay them to Earth. In addition, the uplink from Earth to Icarus-A is to be relayed.

For the link with Icarus-A, the transmit antenna must be complete at the start of the mission but the receive antenna can be formed by an array of smaller antennas that grows as the mission progresses. This is because the received signals can be combined at low power, but the splitting of the transmit signals, at high power, to multiple antennas would be very difficult. For the receive antenna array, the aggregate area of the implemented antennas will be the equivalent to the area of the single antenna used in the link budget analysis.

On-board maintenance and repair will be essential for the Gateway, as for the Icarus spacecraft, because of the distance from Earth. However, as a relay station, little other “intelligence” would be necessary.

For the transmissions to Icarus, a 3 km diameter antenna is required. Once Icarus is on station the required transmitter output power is 0.8 MW, so a Gyrotron HPA is necessary. A Klystron pre-amplifier will also be necessary, as for the Icarus spacecraft. Only two amplifier chains are necessary, however, since only one carrier is to be transmitted.

For the transmissions to Earth, a 30 m diameter antenna is required, with a transmitter output power of 100 kW. This, again, will require a Gyrotron HPA and Klystron pre-amplifier. Only two amplifier chains will be necessary since for this relatively short distance a 20 Gbps signal can be carried on a single carrier.

The 30 m antenna facing Earth can be shared between transmit and receive, but two antenna systems are required for the link to Icarus. For the purpose of this analysis, a single receive antenna of 15 km diameter is assumed.

Scaling up the mass of a 3 km antenna from the 1 km antenna for the Icarus spacecraft, results in a value of 450,000 kg. Scaling up the mass of a 15 km antenna results in a value of 11,250,000 kg, to be achieved in stages as the antenna array is implemented. Scaling down the mass of a 30 m antenna results in a value of 125 kg.

The two transmitter systems will result in a mass of around 4,000 kg. The two receive systems will have insignificant mass in comparison.

So the total mass of the Gateway, in its final configuration, will be close to 12,000,000 kg.

The combined power consumption will be around 2.5 MW with a heat dissipation of 1.5 MW.

LASER OPTICAL TECHNOLOGY

The possibilities for using laser optical communications systems over interstellar distances have been considered for some time. Such systems are now being introduced operationally in Earth orbit, in particular for inter-satellite communications links.

Optical communications links have similar properties to those of RF links, with a path loss proportional to the square of the frequency, but with gains in the transmitting and receiving systems also proportional to the square of the frequency and to the collecting area of the mirror or lens. However, the surface profile of that mirror, or lens, must be accurate to 10 % of the wavelength as was the case for an RF antenna. For an optical communications link the wavelength could (typically) be of the order of 0.5 µm, and so that surface accuracy would need to be 0.05 µm.

An optical communications link is considered in [7]. In order to achieve a symbol error rate of 10^-3, it determines that at least 11.4 photons per pulse must be received. As proposed by Lesh [7], A Pulse Position Modulation scheme is employed, in which a pulse is transmitted in one of 1024 positions, with each position representing a different value. 1024 positions is equivalent to 10 bits of information, and so this modulation scheme can transmit 10 bits per frame. In the case considered in [7] a frame was transmitted once per second, with each pulse 10 ns wide and 0.99999 s of dead time once all possible 1024 pulse positions had passed. This results in a data transmission rate of 10 bps.

Sufficient power is transmitted by concentrating the available laser power into each pulse, such that the average power over the complete second was within the capability of the device for continuous use. Thus a 20 W laser device could actually transmit 2 GW for the duration of a 10 ns pulse, once per second.

The maximum pulse repetition rate depends on the ability of the laser device to dissipate the heat created when transmitting at that high power. Effectively, all of the waste heat must be dissipated before the next frame can be transmitted. Technology has improved since [7] was published and it is now possible to transmit a 10 ns pulse every 10 µs. This results in a frame of (of the order of) 1024 positions being transmitted every 10 µs or 10 bits every 10 µs which is equivalent to a data transmission rate of 1 kbps. For convenience, a data rate of 1 kbps is assumed. The 20 W laser device, above, would be able to transmit each pulse at a power level of 20 kW.

However, it is still necessary to achieve the minimum 11.4 received photons per pulse, and so a higher power device, and/or a larger collecting area, is necessary in order to achieve a working communications link.

The data rate cannot be increased by increasing the transmitted power, as is the case for the RF communications link. An increase in data rate can be achieved only by increasing the number of frames transmitted per second (or increasing the number of bits transmitted within each frame), which is limited by the ability of the laser devices to dissipate their waste heat. But, the symbol error rate will improve with increasing power. Conversely, if the received power drops below the threshold, the symbol error rate will increase and the link will become unusable.

A practical example link budget, for the Icarus mission, and after [7], is shown in Table 6. For this example, the current technology, allowing a bit rate of 1 kbps, is assumed and with an average transmit power of 1 kW. A transmitting mirror of 10 m diameter is assumed, and the receiving mirror is assumed to have a diameter of 200 m. With these values, the number of received photons per pulse is calculated to be 14, well above the required 11.4, and so a 1 kbps link from Alpha Centauri could be maintained.

Also shown in Table 6 is the link budget for continuous transmission at 1 kW. In this case the resulting C/N is -12.47 dB and the signal will be significantly below the level of the background noise. If a spread spectrum modulation scheme were in use, it could be possible to recover the signal in this situation, but some means of synchronizing the receiver and transmitter would be required. Alternatively, approximately 20 dB improvement in the received C/N would be required, by increasing the transmitter and/or receiving mirror diameters, or by increasing the transmitter power level.

Table 6 On Station Optical Link Budget

Link Distance

4.36

ly

4.12E+16

m

TX Laser Power

1,000

W

30.0

dBW

Pulses per Frame

1,000

Pulsed Tx Laser Power

1,000,000

W

60.0

dBW

TX Laser Wavelength

0.532

µm

TX Aperture Diameter

10.00

m

TX Obscuration Diameter

0.20

m

TX Optics Efficiency

0.90

-0.5

dB

TX Bias Error

0.03

µrad

TX RMS Jitter

0.03

µrad

RX Aperture Diameter

200.00

m

RX Obscuration Diameter

2.00

m

RX Optics Efficiency

0.80

-1.0

dB

Narrowband Filter Transmission

0.80

-1.0

dB

Detector Quantum Efficiency

0.80

-1.0

dB

Atmospheric Transmission Loss

1.00

0.0

dB

Background Power

9.23E-12

W

-110.3

dBW

Transmitter Antenna Gain

3.49E+15

155.4

dB

Receiver Antenna Gain

1.39E+18

181.4

dB

Waist Size

4.90

m

3.456E-08

Directional Error

0.03

rad

Transmitter Loss Ratio

2.216E-01

-6.5

dB

Free-Path Space Loss

1.05E-48

-479.8

dB

Net Loss

1.021E-01

-10.0

dB

Received Power

5.23E-10

W

-93.0

dBW

Number of Photons per Joule

2.68E+18

Detected Photon Events per Second at Pulsed Power

1,400,153,976

/s

Detected Photon Events per Pulse

14.00

Received Average Power

5.23E-13

W

-122.82

dBW

Detected Photon Events per Second at Average Power

1,400,154

/s

Bandwidth

3E+09

Hz

C/N

-12.47

dB

If the transmitting mirror were to be enlarged, care would be required to ensure that it remained within the protection of the dust shield during the inter-stellar transit. Because of the required surface profile accuracy, it is unlikely that this mirror could be manufactured after arrival. Higher power laser terminals have been demonstrated, at powers up to 30 kW, but these are actually devices employing multiple 1 kW laser transmitters. If such multiple devices and their signals could be synchronized, then the combined power could overcome part of the shortfall in the link budget. A larger collecting area at the receiver would be essential. A 5 km receiving mirror, combined with the 1 kW transmitter and 10 m transmitter mirror could achieve a viable link budget, with a 12 Gbps bit rate. However, even if the transmitting mirror is increased to 20 m diameter and the receiving mirror to 30 km diameter, the maximum bit can only increase to approximately 12.5 Gbps. These mirrors would be challenging to manufacture, given the required accuracies. Increasing the transmitter power, similarly, offers limited opportunity to improve the achievable bit rate. With all of the increases in performance mentioned above, the maximum achievable bit rate is still only around 15 Gbps.

A laser system could offer a useful communications link between Alpha Centauri and Earth, but it could not, with foreseeable development, achieve the data rate required for the Icarus mission. However, current development activities are aiming for laser communications within (at least) the inner Solar System. In the timeframe of the Icarus mission it may well be possible to use laser communications within the Alpha Centauri star system.

CONCLUSION

An outline design for communications subsystems has been presented, using RF technologies and conservative assumptions. With the exception of the remote construction of large structures and further development of high power Gyrotron amplifiers, the design is feasible using current technologies. Estimates for the mass, power consumption and heat dissipation for the RF equipment and the antennas have been generated. Laser technologies could provide a communications link over the distance to Alpha Centauri, but not at the required data rate.

Gyrotron and Klystron high power amplifiers have not yet been flown in space, but that should not be an insurmountable obstacle. As sufficiently high power amplifiers become available at higher frequencies, it will be possible to increase the data rates and/or to reduce the diameter of the antennas for the inter-stellar communications links.

Over the 4.37 light year distance, between Earth and the target star system, a 20 Gbps downlink data link can be maintained. At the target star system, communications requirements between the probes are orders of magnitude lower [3]. The recent successes of New Horizons at Pluto and Ceres, and the continuous operation of Voyager, show that this is already possible at the largest possible scale in the solar system. However, the power available from a large probe like the Icarus fusion starship allows for data rates so far unseen above the Earth’s atmosphere.

For missions to more distant star systems, a similar design of communications system would remain feasible, but with lower data rates supported.

The missing ingredient will be the human operators and mission specialists who continuously adjust and monitor the space probes. The ways by which the Icarus Interstellar probe can execute these types of tasks are explored in the Computers and Software chapters.

RF propagation

The power of an electromagnetic transmission decays proportionally to the square of the distance travelled. This is sometimes referred to as the spreading loss, where the power is assumed to be transmitted in all directions. Otherwise it is known as the path loss. When the power is concentrated into beams, the power transmitted in the direction of the beam is increased and the antenna is said to have “gain”. Antennas are reciprocal devices, and their transmit and receive gains (at the same frequency) are the same. A transmission between two antennas benefits from the gain of both antennas.

The received power of a transmission can be calculated from the following formula for the Free Space Path Loss (which does not include other effects such as atmospheric absorption, or scattering, etc)

PR=PTGTGR(λ4πd)2

where,

P_R is the power at the output of the Receive Antenna,

P_T is the power at the input of the Transmit Antenna,

G_T is the gain of the Transmit Antenna,

G_R is the gain of the Receive Antenna,

λ is the wavelength of the transmitted signal, and

d is the distance between the two antennas.

λ and d are both expressed in the same units, metres. P_R and P_T are both expressed in the same units, watts.

It is common in RF engineering to use logarithmic units, so that gains and losses can be added and subtracted, rather than multiply and divide large numbers. These units are decibels, which are calculated as 10 times the logarithm of the value (to base 10). So, for example, a power of 20 watts would become 13 dBW. If the power was better expressed in milliwatts, then that would become 43 dBm. Gains and losses, as ratios in “real life”, are expressed simply as dB.

Thermal Noise

Thermal noise is created in all electrical circuits, caused by random movements of the electrons. This tends to be at a constant level, and so affects low power signals more than those at high power. Devices can be described as having a Noise Temperature from which the level of noise can be calculated. The Noise Temperature is related to the physical temperature by a parameter called Noise Factor. Alternatively, Noise Figure may be quoted, and this is the Noise Factor in dB.

The Noise Factor, F = 1+ TETO , where T_E is the environment temperature and T_O is the standardized temperature of 290 K. Lossy devices (such as attenuators, cables and waveguides) have a Noise Factor related to their loss, =1+(L-1)TETO ,

where L is the loss of the device.

The Noise Power resulting from that thermal noise is calculated from the formula

PN=kBTE

where,

k is Boltzmann’s Constant (1.3806×10⁻²³ J/K) and

B is the bandwidth in which the signal and the noise are being measured

A noise density can be calculated by removing the bandwidth from that calculation.

Data Link Capacity

The Shannon-Hartley Theorem can be used to determine the maximum data rate that can supported over a communications link in the presence of noise. That theorem can be stated as

C=Blog2(1+SN)

where,

C is the channel capacity in bits per second,

B is the channel bandwidth in Hz,

S is the received signal power, in Watts, and

N is the noise power in the bandwidth, in Watts

S/N is referred to as the Signal to Noise ratio, although its dB equivalent is often the actual value quoted.

References

[1] R. P. Hoyt, J. I. Cushing, J. T. Slostad, “Spiderfab: Process for On-orbit Construction of Kilometer Scale Apertures”, Tethers Unlimited Inc, NASA Innovative Advanced Concepts, Grant # NNX12AR13G, 2013

[2] R. Temkin, “High Frequency Gyrotrons and Their Applications”, Plasma Physics Colloquium, Columbia University 28 February 2014.

[3] Internal project papers by Shankar and Tziolas

[4] J. Ruze, “The effect of aperture errors on the antenna radiation pattern”, Nuovo Cimento Suppl., Vol. 9, No. 3, 364-380, 1952.

[5] R. M. Manning, “Analysis of Electromagnetic Wave propagation in a Magnetized Re-Entry Plasma Sheath Via the Kinetic Equation”, NASA/TM—2009-216096, 2009

[6] C. E. Shannon, “Communication in the presence of noise”, Proc. Institute of Radio Engineers 37 (1): 10–21, 1949

[7] J. R. Lesh, et al, “Space Communications Technologies for Interstellar Missions”, JBIS, Vol. 49, pp. 7-14, 1996

Computers and data management