European Solar Polar Orbiter Mission

Top-Level Baseline Science Objectives

The many potential science objectives and goals of a SPO mission have previously been discussed in detail in Ref. Error: Reference source not found. The purpose of this paper is to address the technology goals and requirements of such a mission and as such discussion of science goals is limited to top-level baseline objectives.
The solar wind, in addition to higher energy solar flare particles, can induce power line surges and radio interference on the Earth, as well as causing the well known aurora borealis. Observations from Ulysses shows that the 11-year solar cycle minimum causes the solar wind speed at polar latitudes to almost double the equatorial value, from a speed of order 450 km s^-1 to 750 km s^-1. The solar wind also appears to have a different composition at the solar poles. A close solar polar orbiter would thus enable further investigation into the polar solar wind data obtained by Ulysses and likely to be obtained from the Solar Orbiter mission over a range of inclinations up to 35 deg. Furthermore, it is important that we can obtain an understanding of the relationship between solar wind velocity and the solar magnetic field geometry, with the best location to accurately assess the longitudinal structure of the magnetic field in the corona being from polar latitudes. Solar polar observations would also address the scale over which the co-rotation of coronal plasma with the Sun is lost. Combined coronagraph data would thus allow the determination of three-dimensional structures and show the locations of streamers, rays, and plumes in the corona. Considerable fine structures, termed microstreams, were observed in high-speed flow from coronal holes at the poles by Ulysses. Relating the microstreams to polar plumes, supergranulation patterns and bright flares would be enabled by a spacecraft in a solar polar orbit.Error: Reference source not found

Mission Architecture

The mission is split into seven core phases, ranging from Launch through to sail jettison and the beginning of the science mission and then on into a potential extension to the science mission. The longest mission phase is the transfer trajectory, which is provisionally scheduled as 5 years, although this will vary depending on the final selected sail characteristic acceleration. We define characteristic acceleration as the acceleration the sail actually provides at a solar distance of 1 AU, with the sail normal to the Sun-line. Following the arrival of the spacecraft at the solar polar orbit the sail is jettisoned to allow the science operations phase to begin. The spacecraft attitude and orbit maintenance is from this point on performed using a hydrazine system as will be discussed later. Science operations are provisionally scheduled for 2 years. The target solar polar orbit is defined by the direction of the solar poles. Thus, the desired polar orbit is inclined at 82.75 deg with a right ascension of ascending node of 255.8 deg at J2000 (corresponding to the Julian day 2451545) with a drift rate of plus 0.014 deg yr^-1, within a standard ecliptic plane reference frame. Analysis of sunspots has revealed that the direction of the solar poles is less well defined than indicated above,⁵ however we adopt these values as the target orbit. Spacecraft orbit phasing with respect to the Earth must be carefully considered. Science returns are maximized when the spacecraft is positioned near to the solar limb as seen from Earth, allowing observation of the corona along the Sun-Earth line. Maintaining this alignment eliminates solar conjunctions and hence loss of telemetry. It is thus considered necessary that the spacecraft orbit is in resonance with Earth’s orbit about the Sun. Potential target solar orbits are defined as a circular polar orbit with radius N^-2/3 AU for integer values of N, where N is the orbit resonance number. Figure 1 illustrates the Earth – Sun – sail separation angle for 1 ≤ N ≤ 5. It is seen that as the orbit resonance number is increased the separation angle tends further from 90 deg, hence degrading mission science returns. At N = 1 the Earth – Sun – sail separation angle stays within ± 27 deg of 90 deg, while at N = 5 the Earth – Sun – sail separation angle peaks at over ± 73 deg from 90 deg. The choice of optimal resonant orbit depends on a number of factors. Far from the Sun, larger aperture instruments are required to maintain image resolution, with only infrequent passes over the solar pole. However, the Earth – Sun – sail separation angle stays close to 90 deg. Closer to the Sun we obtain frequent passes over the solar poles and very high resolution imaging, but the spacecraft thermal environment becomes increasingly severe, while also passing further from the solar limb. Thus, a balance must be sought based on spacecraft engineering constraints, cost and science goals. The N = 3 resonant orbit is defined as the target scientific orbit as this places the spacecraft close to the Sun, while also being in a relatively benign thermal environment compared to closer resonant orbits. This orbit also maintains the spacecraft within ± 30 deg of the solar limbs for the majority of the mission duration. Figure 1 also illustrates the trajectory of the solar polar orbit for N = 3. The trajectory is seen from an Earth-centered coordinate system, with the Sun fixed along the negative X axis. The asymmetry for N = 3 (and N = 1) can be reversed by a simple alteration of the initial conditions.

Figure 1 Earth – Sun – sail separation angle for 1 ≤ N ≤ 5, top, and trajectory plot for N = 3 solar polar orbit in an Earth-centered rotating reference frame, bottom.

Spacecraft Model

Within this paper the term “spacecraft” means the vehicle which will perform the science operations at the defined target orbit and does not necessarily include the solar sail. Conventionally this craft has been called the “solar sail payload” however the spacecraft will command and control the solar sail, which will not be capable of independent operation. The term spacecraft is thus more appropriate. The solar sail is a fully-integrated sub-system of the spacecraft, however for the purpose of technology definition requirements it is presented here as a separate entity.
Throughout the solar polar spacecraft design full redundancy is maintained, except for the high-gain antenna (HGA). We note further that many components of the solar sail, such as sail film, cannot be supported through redundancy due to mass and/or volume considerations however full redundancy is applied to the sail where possible. The spacecraft systems analysis is based on a minimum solar approach thermal limit of 0.48 AU, the baseline mission profile. However, an analysis will also be presented as to the effect of varying the thermal limit. An overview of the spacecraft mass budget is shown in Table 1 as part of a complete launch mass breakdown. We note that the total spacecraft wet mass is 247.9 kg, of which 41.2 kg consists of the science instrument allocation. Table 1 gives the current best estimate (CBE) mass which then has a design maturity margin (DMM) added to give the total sub-system mass allocation. The design maturity margin is added at equipment level, where > 5 % is added for off-the-shelf items (European Cooperation for Space Standardisation, ECSS, Category: A/B), > 10 % for off-the-shelf items requiring minor modifications (ECSS Category: C) and > 20 % is added for new design/development items, or items requiring major modifications or re-design (ECSS Category: D). We note in Table 1 that the added DMM can appear somewhat arbitrary. For example, the power sub-system DMM is 8.3 %. However, this is simply a result of averaging the DMMs allocated at the equipment level. We anticipate only limited technology issues with the spacecraft sub-systems as the space environment at 0.48 AU is relatively benign in comparison with other currently envisaged missions, including BepiColombo⁶ and the solar orbiter mission. Note however that direct adaptation of technology is rarely possible and thus some limited modifications and developments are inevitable.
Table 1 Mass budget, with scaling laws rounded to one decimal place.

System	CBE Mass	DMM	CBE Mass + DMM
Science Instruments	37.0 kg	10.8 %	41.2 kg
Attitude & Orbit Control System, AOCS (dry)	28.7 kg	5.0 %	30.1 kg
Telemetry, Tracking and Command, TT&C	48.3 kg	5.0 %	50.7 kg
On-Board Data Handling, OBDH	4.2 kg	10.0 %	4.7 kg
Thermal & Radiation	9.6 kg	10.0 %	10.6 kg
Power	42.9 kg	8.3 %	46.8 kg
Mechanisms & Structure	49.2 kg	7.5 %	53.0 kg
Spacecraft Nominal Dry Mass At Launch			237.1 kg
AOCS propellant, inc. sail separation allowance and a margin			10.8 kg
Spacecraft Nominal Wet Mass at Launch			247.9 kg
Solar Sail Nominal Mass at Launch (see also Table 5)			195.9 kg
Nominal Launch Mass			443.9 kg
ESA System Level Margin		20.0 %	88.8 kg
Total Mass At Launch			532.7 kg
Soyuz Fregat 2-1b launch capacity (C3 = 38.84 km^2 s^-2)			620.0 kg
Launch Margin			87.4 kg (14.1 %)