The Space Environment
Space missions must work under many constraints. The payload mass, the power available to run the instruments and the bandwidth available to transmit the observations are likely to determine what can and cannot be done. Often the first consideration in planning a mission is determining the orbit or trajectory that is required and is feasible to achieve. In Earth orbit one has to be concerned with the radiation environment (Fig. 2.1). At low altitudes, say below 500 km, and at low inclinations, ~28°, one can have a mild annual radiation dose ~3 krad with minimal shielding but at higher altitudes around 700 km in an orbit that crosses the pole one would have to shield the electronics with 40 mils of aluminum to achieve a dose as low as 10 krads per year. At higher altitudes such as geosynchronous orbit a reasonable compromise would be electronics that could survive 100 krads per year with 100 mils of aluminum shielding all around the box. The cost and availability of such radiation tolerant parts, however, limits what can be done in space.
One also has to be concerned about space debris and the presence of other spacecraft in low Earth orbit (LEO). Figure 2.2 shows the number of objects in a 10 km band with a diameter greater than 10 cm as a function of altitude. A 10 cm object traveling at 7 km/sec can disable a spacecraft quite readily. Geosynchronous orbit is also becoming very crowded.
If one is interested in imaging from LEO, one might wish to view the Earth under fixed lighting conditions. This can be done with a near-polar, retrograde orbit whose inclination depends on the altitude of the spacecraft. This is called a sun-synchronous orbit. Due to gravity torques this orbit plane precesses just the right amount to keep the spacecraft viewing a fixed local time on each of its two (north and south passes).
Another concern in Earth orbit is the maximum length of time a spacecraft spends in the Earth's shadow (Fig. 2.3). At low altitude this is about 40 minutes. Above about 2000 km this time climbs to reach about 70 minutes at synchronous orbit and continues to climb above that.
Most launches are assisted by the rotation of the Earth so these launches are performed at low latitude such as the Eastern Test Range near the Kennedy Space Center but, if one wishes to launch into a high inclination (Polar) latitude orbit, one may launch from a higher latitude site with a north-south path over water such as Vandenburg Air Force Base, CA. If one is interested in a distant planetary encounter, one must arrive at a time when the energy required for the mission is a minimum. This is usually measured in terms of the constant, C3, that is the energy of a kg of the spacecraft once it gets away from the Earth's gravitational field. The units of C3 are (km/s)2. Figure 2.4 shows contours of C3 for different launch and arrival dates for a 2007 Mars mission. Total flight time in days is shown by the straight lines.
All the energy required for a planetary mission need not be provided at launch. Some can be picked up at gravity assists and some can be provided by onboard propulsion systems. If the amount of onboard maneuvering is large and the maneuvers do not need to be quick, then advanced propulsion systems such as ion propulsion or solar sails can be considered. Gravity assists can be obtained during returns to Earth, during flybys of Venus, Mars and Jupiter and when in orbit about Jupiter and Saturn, from their moons. Figure 2.5 shows a trajectory for a Mercury mission with two gravity assists at Venus and two at Mercury. By using these gravity assists the onboard fuel usage is reduced by two thirds.
Part of the reason for the push to smaller spacecraft is that the cost of launching spacecraft is quite expensive. Figure 2.6 shows the possible launch vehicles and their lift capability to low Earth and geosynchronous orbits. A planetary launch would require greater C3 and hence a smaller mass could be sent on its way to a planet. The injected mass as a function of C3 is given for two members of a recently introduced series of launch vehicles in Figure 2.7 One way to reduce the size of the launch vehicle is to include propulsion on the spacecraft. A spacecraft that can carry a small but very efficient engine that over a long period of operation can make a major change in a trajectory. This is suited to trajectories that do not require a rapid change in the momentum of a spacecraft. Figure 2.8 shows the specific impulse of engines (a measure of the momentum that a kg of fuel obtains) plotted versus the thrust available from such an engine. An ion electric propulsion thruster can use fuel about an order of magnitude more effectively than a chemical thruster but provides orders of magnitude less thrust. Thus orbital and trajectory changes are much less rapid. The ion engine on the Deep Space 1 mission had a thrust of 90 mN and a specific impulse of 3100 s. The specific impulse when multiplied by the acceleration of gravity gives the velocity of the fuel. Thus, the xenon ions expelled from the Deep Space 1 thruster were accelerated to 30 km/s. The successor to the DS1 (and Dawn) mission thrusters are expected to have an ISP of over 4000 s, with a thrust upto 230 mN.
Another important aspect of mission design is the manner in which the attitude of the spacecraft is maintained. Figures 2.9 and 2.10 show the means usually employed. A spinning spacecraft is relatively easy to maintain because the spin axis remains quite steady in direction under most circumstances. Some instruments like to use the spin of the spacecraft to scan around the spacecraft but others like to point in a particular direction. This can be done by adding a despun platform so that most of the spacecraft spins but some of it is not spinning relative to the target. Another option is to stop the spin altogether. This can be done in two ways with reaction or momentum wheels and with thrusters. If there will be a lot of motion of the spacecraft back and forth, then reaction wheels are preferred because thrusters use fuel while reaction wheels use a renewable resource, power. In order to determine what the attitude of the spacecraft is, star trackers, sun sensors and horizon sensors are used.
Onboard propulsion is usually a serious complication to a mission but it is difficult to design a mission without at least some thrusters to spin up or down the spacecraft or to reorient the spin axis. A typical bi-propellant engine is sketched in Figure 2.11. It has a lot of plumbing. A typical mono-propellant (hydrazine) system is shown in Figure 2.12 and a typical solid-state thruster is shown in Figure 2.13.
The figure of merit of an advanced propulsion engine is the specific impulse, measured in seconds, that is the velocity of the exiting gas normalized by the acceleration of gravity (on the surface of the Earth). The ion engine used on the Deep Space 1 mission is schematically shown in Figure 2.14. The engine uses xenon for fuel, ionizes it and accelerates it through a grid. Electrons are then injected to neutralize the beam. It is estimated that the Deep Space 1 engines at full thrust can fire for about 12000 hours before eroding its grid and in that process accelerate about 150 kg of xenon. Two disadvantages of the ion engine are its need for a powerful source of electricity. Typically an ion engine requires 2.5 kW to run. On DS1 this is supplied with a large solar array. The second problem, especially on space physics missions, is electromagnetic contamination. The plasma plume will cause noise in the electric antennas needed to measure plasma waves and the strong magnets in the thrusters will require that any magnetometers be placed far away from the thrusters.
An alternate ion engine is the Hall thruster that also accelerates xenon but to a slightly lower velocity. The recent SMART-1 mission to the Moon by ESA used a Hall thruster (see Figure 2.15) to leave Earth's gravitational field and orbit the moon. Two developmental thrusters with powers of 8 and 3 kw are shown in Figure 2.16. These are at a lower level of readiness for use in an operational mission than the electrostatic strusters.
Solar sailing in which light pressure is reflected by the sail and thereby transfers a force to the spacecraft is a propulsion technique that is not yet qualified for space. Proposed solar sail concepts are shown in Figure 2.17. A practical solar sail on a realistic mission might be 50m across. These sails are deployed either with masts or an inflatable collar. None of these concepts have yet been tested. A simple solar sail mission involves a parachute-like sail. It requires little maneuverability when on station because of the reflected sunlight and just orbits the sun. Because the force from the reflected sun light is opposite solar gravity the spacecraft can orbit the sun within a one year period yet be much closer to the Sun than is the Earth. A three spacecraft mission, illustrated in Figure 2.18. could hover above and below the ecliptic plane and orbit the sun in small (constant latitude) circles while a third orbited in the ecliptic plane and monitor the structure of magnetic disturbances approaching Earth, allowing geomagnetic predictions to be made more accurately.
The successful return of data from space is necessary to complete any scientific mission. Operating a ground station to receive these signals is another expensive endeavor. Thus it is the practice now to store data for later transmittal. In any event, the data return requires an antenna of a certain size and a transmitter of sufficient power to return the data within a specified time with a bit error rate below some limit. Table 2.1 lists the radio bands in use in space. Past missions often used S band near 2 GHz. Deep space missions now most frequently use X-band near 7.5 GHz. K band near 14 GHz provides even greater bandwidth for data return but is affected by terrestrial weather. NASA's principal ground stations are at Goldstone, CA, near Madrid, Spain, and near Canberra, Australia. At these locations they have 70 m, 34m and 26m and occasionally 11m stations.
Table 2.1 Research Satellite Frequency Bands
|Band||Frequency Range [MHz]|
Each spacecraft has a command and data handling system that receives commands from Earth and collects the data from the instruments and sends it back to Earth. Terrestrial antennas are generally quite directional and can be accurately aimed at the telemetering spacecraft. In space, it may be difficult to point a high-gain (dish) antenna, especially if the spacecraft is spinning. Also if the spacecraft has some emergency and cannot point to Earth, a high-gain antenna may not be able to receive the signals broadcast from Earth that might help rectify the problem. Thus spacecraft usually have a series of antennas with different angular response functions and gains. These might be a low-gain, nearly omnidirectional antenna such as a dipole antenna, a less omnidirectional, belly-band, medium-gain antenna (on a spinner) and high-gain (dish) antenna. To minimize ground station costs, data are usually stored onboard and transmitted over short (4-8 hour) periods at the highest possible data rate.
The commands to the spacecraft and the data received on the ground are usually handled by a spacecraft operations facility. Such a facility takes the operations sequences planned by the scientists and turns them into digital commands for the spacecraft. It also directs the science data to the scientists' facilities where they are analyzed. While data are often transferred over the internet, NASA's commanding links are separate and are more secure. Ultimately the calibrated records and processed data are stored in archives such as the Planetary Data system and the National Space Science Data System.
When spacecraft enter eclipse behind a planet as seen from the sun, solar panels are ineffective and an alternative temporary source of power, such as a battery, must be used. For missions that go too far from the sun (about 3AU), they also must use some alternative power system but a battery does not hold sufficient power. Thus outer planetary missions such as Pioneer 10 and 11, Voyager, Ulysses, Galileo and Cassini have used radionuclide thermal generators or RTGs. These generate electricity that is then distributed to the spacecraft.
Thermal control is important on a spacecraft because some instruments need cool detectors. Other instruments must be kept warm to operate properly. Missions close to the sun tend to require active means to cool them while heaters are required by missions to objects far from the sun or that spend a long time in eclipse.
Many measurements, especially those of space physics, require electrostatic and electromagnetic cleanliness. The magnetic field of permanent magnets and current loops must be kept small. Alternating currents must not affect electromagnetic wave measuring devices. The electric potential of the spacecraft must be kept low and uniform so that the plasma may be accurately measured. Finally, while the radiation belt fluxes mentioned above can cause permanent damage to instruments, they also might cause temporary latchups. These too must be guarded against. Thus the environment of space is a difficult one in which to work and much caution must be exercised.
A Sample Discovery Mission: NEAR
The NEAR mission was one of the first two Discovery missions selected for flight, the other mission being Mars Pathfinder. NEAR was launched in February, 1996 and after an inflight motor misfiring that delayed arrival, it entered orbit about the asteroid 433 Eros in February 2000 with last data (taken from the surface of Eros) in February, 2001. NEAR weighed 788 kg of which 319 kg was fuel. It carried 56 kg of scientific instruments. The spacecraft structure weighed 102 kg; the harness 39 kg; the propulsion system 118 kg; the power system 64 kg; the telecommunication system 25 kg; the command and data system, 19 kg; the guidance and control system 34 kg, and the thermal system 11 kg. The spacecraft took a minimum power of 314 W to operate. A list of each of these systems and their subsystems, their masses and powers are given in Table 2.2. The components that comprised the guidance and control system is given in Table 2.3. Figure 2.19 shows a sketch of the NEAR spacecraft showing the location of the instruments.
Multi-spectral imager (MSI)
Near imaging spectrograph(NIS)
Propellant and pressurant
Power system electronics
High gain antenna
Medium/Iow gain antennas
Solid state amplifiers (2)
Command detector units(2)
Telemetry conditioner units(2)
RF switches coaxial cables
|Guidance and control
Reaction Wheels (4)
Inertial measurement unit
Digital sun sensors(5)
Attitude interface unit
|Command and data handling
Command and telemetry processors(2)
Solid state recorders(2)
Power switching unit
Spacecraft primary structure
Spacecraft secondary structure
Despin mass and balance mass
Thermal blankets, heaters, thermostats
Propulsion survival heaters
Spacecraft and instrument survival heaters
Instrument operations heaters
Harness and terminal boards
|Inertial measurement unit||Delco||Gyros (4):30 mm hemispherical resonator gyros
Rate bias < 0.01 deg hr-1over 16 hr
ARW <0.001 deg hr-1/2
Accelorometers (4): Sunstrant QA-2000
<100 g RMS noise
|Star tracker||Ball||FOV: 20 X 20°
Sensitvity: +0.1 to +4.5M
No. of stars tracked: 5
Output rate: 5 Hz
|Reaction wheels||Ithaco||Brushless DC motor
Momentum: 4 Nms (@ 5100 RPM)
Torque: 0.025 Nm
|Sun sensors||Adcole||Quantization: 0.5°
|Attitude interface unit||JHU/APL||Clock: 6 MHz|
Memory (16 bit words):
Processor: RTX 2010
|Flight computer||Honeywell||Clock: 9 MHz;|
Memory (16 bit words):
The multspectral imager shown in Figure 2.20, was one of the workhorses of the NEAR mission. A stepper motor rotates a filter wheel past the light path allowing pictures to be made in different colors. Figure 2.21 shows the transmission bands of these filters as well as the transmission efficiency of the optics and the response of the CCD that detects the signal.
The NEAR spacecraft carried a near-IR spectrometer to identify the minerals on the surface of Eros. Figure 2.22 shows the light path of this instrument and how it analyzed the signal as a function of wavelength with its grating. As illustrated in Figure 2.23, which gives a block diagram of the instrument, part of the spectrum was measured with a Ge detector and part with an InGaAs detector.
The X-ray and gamma-ray spectrometers were basically two separate sets of sensors with common electronics. These sensors are shown in Figures 2.24 and, 2.25a. As illustrated in Figure 2.25b the gamma rays were detected by a NaI crystal surrounded by a BGO shield with a small photomultiplier tube in front and a large one in the rear. A block diagram is shown in Figure 2.26 showing the pulse height analyzers applied to the signals from both the X-ray and gamma-ray sensors.
In order to measure the precise shape of Eros, NEAR carried a laser ranger shown in Figure 2.27. The light path in the laser is shown in Figure 2.28. A block diagram of the instrument is given in Figure 2.29 and the data processing unit is shown in Figure 2.30.
Finally, NEAR carried a magnetometer with sensors on the high-gain antenna. A block diagram of the instrument is given in Figure 2.31. The measurement was noisy because of the closeness of the sensor to the subsystems on the spacecraft. A boom on which the sensors could have been mounted would have reduced this source of noise.