ESA Science & Technology - Q&A for Solar Orbiter Payload AO

Concerning the development of CZT, we have concentrated on building up the capability of growing high quality material in Europe.
There is a contract running with Durham, Scientific Crystals (C 20493) under the ITI initiative with the following targets:

Phase 1 ITI Demonstration of the vapour phase growth of single crystal cadmium zinc telluride material – 6 months duration

• Demonstrate the feasibility of growth of CZT using the MTPVT method
• Determine the growth parameters using a bench top simulator
• Expand existing CdTe models to simulate CZT growth conditions
• Validate model with bench top simulator results

Status: This first phase has successfully been concluded

Phase 2 ITI Production of single crystal cadmium zinc telluride (CZT) for detector applications – 12 months duration

• Fabricate a fully operational vapour growth system for the production of single crystal CZT from two binary sources
• Adjust growth conditions to optimise the crystals for use as high energy radiation detectors

Status: Industrial reaction vessel built. Producing large area single xtal high resistivity. We hope to conclude this activity in Q1/2008.

We are still far from having a reliable source for CZT detectors in Europe, but at least, the first steps have been taken.

The maximum length for remote sensing instruments is 1 metre, as stated in the EID-A, section 4.6.2.

Instrument teams are allowed to contact Galileo Avionica for a quote for the entrance window cost. The contact person is

	Mr Giovanni Taglioni phone: +39.08639062202 e-mail: giovanni.taglionigalileoavionica.it
(who may then refer you to a commercial officer).

A Structural Mathematical Model is required per EID-A section 6.3.1.1.
Thermal Mathematical Models are required per EID-A sections 6.3.2.4, 6.3.2.5, 6.3.2.6.
A Structural & Thermal Model (STM, hardware) and an Engineering Model (EM, hardware) are required per EID-A section 6.8.1.2.
The schedule of deliveries (TBD for the mathematical models and TBC for the hardware models) is per EID-A section 8.8.3.

We would like to bring to your attention that the baseline measurements for solar wind ions (i.e., the SW-PAS instrument) stated in Table 3-3 of the HELEX JSTDT report are significantly different when compared with those stated in the Solar Orbiter PDD. Please clarify if the information given in the HELEX report now represents the actual requirements.

The PDD requirements for SW-PAS are given on page 73:

Normal Mode
A typical time resolution of 3s for 3-D velocity distribution functions;
FOV: ± 45° to Sun, ± 15° north/south at 5° deg resolution

Burst Mode
Measurements of 2-D distributions of solar wind protons and alpha particles to resolve typical growth times of instabilities, i.e., at a cadence of 0.1s

In contrast, the Prioritized Solar Orbiter Measurements for solar wind ions as stated in Table 3-3 (page 48) of the JSTDT report for HELEX are:

Baseline Measurements for solar wind ions:
Detailed 3-dim velocity distribution functions at 1 s time-resolution;
FOV: ± 45° to Sun, ± 15° north/south, angular resolution of 2°

Augmented Measurements for solar wind ions:
Burst mode at 0.1s time resolution or better for protons for 1 dim

• In their response to the ESA Solar Orbiter Payload AO, U.S. proposers provide their proposal to ESA, according to the procedure described in the ESA AO. The selection will be performed by an ESA Payload Review Committee (= selection committee) which will include U.S. members appointed by NASA in consultation with ESA. Selection by ESA does not entail any funding by ESA. The procedure for U.S. proposers is identical to the procedure for any other (non U.S.) proposer.
• Simultaneously, U.S. proposers provide their proposal to NASA in order to secure NASA funding, according to the NASA SMEX/FOSO procedure described in the NASA SMEX/FOSO. This procedure is similar to the procedure for any other (non U.S.) proposer, who also has to secure funding from a national agency.
• Because of the respective AO schedules, U.S. proposers are required to provide within their proposal to ESA a NASA document certifying that their proposal has been received and accepted by NASA and declared compliant with the SMEX/FOSO rules. (Other proposers are required to include a full-fledged Letter of Commitment from their funding agency.)

Section 10.2 Reference Payload states that the "Nominal accommodation of all instruments has been defined with the exception of a wide-field heliospheric imager. With the caveats mentioned in the previous paragraph, it is anticipated that the available resources for a wide angle coronagraph/heliospheric imager will be in the range of 5-10 kilograms in mass, 5-10 Watts in power, and a telemetry bandwidth of ~20 kilobits per second." In section 3.6.3 of the ESA Payload Definition Document (PDD), there is a chart that defines the instrument accommodation maximum resources as slightly different. Which document (the SMEX AO Addendum or the ESA PDD) should we consider as the master for developing our proposal concept? 

NASA Response to Question 11:
The FOSO Amendment to the SMEX AO takes precedence over the ESA PDD.

ESA complementary response: The evaluation of proposals will be based on the ESA documentation including the EID-A, which contains a payload resource table. The EID-A is applicable. Since spacecraft resources are at a premium, a proposal with a higher mass will be graded lower than a proposal with an EID-A compliant mass. The mass range indicated by NASA's response is the range beyond which a proposal will not be examined. The same applies to the power and telemetry. Concerning telemetry, one should take into account that remote-sensing instruments will not operate along the whole orbit, therefore the bandwidth figure depends on the proportion of the orbit on which this particular instrument generates data. This resolves partially or wholly the discrepancy between the ESA and NASA bandwidth figures.

Section 3.6.3.2 of ESA's Payload Definition Document, shows the Heliospheric Imager instrument mounted to the same side as a solar panel in figure 3.6.2. Straylight from the solar array makes it extremely difficult to meet filed of view requirements associated with the Heliospheric Imager. During the accommodation study, would ESA consider moving the solar panels to another plane while keeping the Heliospheric Imager instrument in the ram or anti-ram side and on the ecliptic plane? 

NASA Response to Question 13:
Moving the solar panels may be considered during an accommodation study.

ESA complementary response: The accommodation study will not start from scratch. It will be aimed at optimizing the accommodation of the instruments (including GHI/WFC) within the spacecraft, but not at fundamentally re-designing the spacecraft. In addition, in this particular case, alternatives to the baseline solar panel location have been considered already to a significant extent and result, as far as we know today, in solutions worse for the instrument than the current baseline.

The (satellite) PFM is not equivalent to the traditional QM but it does mean that part of the (satellite) qualification will be performed on the (satellite) Flight Model, hence its designation as Proto-Flight Model. However, we anticipate that relatively little will be qualified on the satellite PFM, since the primary structure will be qualified on the satellite STM and the electrical aspects will be for the most part qualified on the satellite EM. Designating the satellite flight model as PFM gives us some limited flexibility in this respect at satellite AIV level, for the case programme development identifies beneficial options in schedule and/or cost e.g. by re-using some part of the satellite STM into the satellite PFM. This also explains why the (satellite) FM is absent from the list on page 129.

At instrument level, instrument teams have some leeway to propose the model and AIV philosophy they deem most suitable. However they must remain compatible with the satellite-level philosophy as described by the EID-A, hence:

• the instrument STM has to be put through thermal vacuum testing to ensure its capability to participate in satellite-level thermal vacuum testing; and the test should include thermal balance phases to ensure its representativity for the instrument.
• the instrument EM is not required to go through thermal vacuum testing, since it will be used at satellite-level in the satellite EM, which is not scheduled to go through thermal vacuum. However, instrument teams may elect to perform thermal vacuum qualification testing on this model, making it an EQM. This is a classic approach. It greatly reduces the risk of discovering thermal problems too late, on the FM. It also avoids stressing the FM by acquiring qualification on it (thus making it a PFM). Depending on the make-up of the instrument, if it significantly saves resources without unduly increasing risk, it may also a good compromise to adopt a STM/EQM/FM approach for some more critical units and a STM/EQM/PFM approach for some less critical or better known units e.g. those only slightly modified from existing qualified units.
• the FM and FS have to be put through acceptance thermal vacuum / thermal balance testing. If qualification has not been acquired on a QM or EQM, then the FM is actually a PFM and has to be put through qualification thermal vacuum / thermal balance testing.

• instrument models are an STM, an EM (not vacuum-tested), an EQM (flight-standard), an FM.
• The qualification testing is performed on the EQM.
• The FM is acceptance-tested.
• After completion of qualification, the EQM becomes the FS (with minor refurbishment if/as applicable). Note that in this case a separate EM is required, since the EM will remain with the spacecraft until after completion of payload in-orbit commissioning as per EID-A section 6.8.1.2.

Does ESA have a position on the use of fold-outs? Would it be acceptable to include 2 fold-outs in addition to the 40-page A4 limit?

When a unit is mounted outside the spacecraft, it is assumed to be thermally decoupled e.g. with washers, CFRP legs from the spacecraft structure. In this case, the instrument will control by itself the temperature of the URP. Per definition in 4.7.1.1, the URP is "on the unit side close to the mechanical interface to the spacecraft". On the spacecraft side, there will be a Spacecraft Reference Point (SRP) whose temperature will be controlled by the spacecraft Thermal Control Subsystem.

• Illumination from direct sunlight at 0.22 AU for short duration (15 degree off-pointing for up to 20 seconds, TBC).

1. It is acceptable to put the details of the funding agencies and Co-Is in the management plan.
   
   
2. It is acceptable to exceed the number of pages in chapter 3 if the total number of pages of chapter 1 and chapter 3 remains 30.

• As long as the proposal is made available in electronic form on the proposer's FTP server (or sent by email if the proposal is small enough) by the stated deadline of 15 January 2008 24:00 CET, ESA will accept receiving the paper copies no later than 17 January 2008 close of business.
• ESA will accept the required funding agency letters of endorsement until 08 February 2008.