5. Technical Reports

5.1. HST10X Structural Dynamics Assessment

The HST10X Structural Dynamics Assessment was conducted to investigate the mass properties, interaction with the Pointing Control System, thermal deformation, and potential line-of-sight jitter of the Hubble Space Telescope with the HST10X enhancement. The dynamic mathematical model of HST developed after the first servicing mission (SM-1), was modified with added large mirrors and light-shield to approximate the properties of an example HST10X configuration.

Conclusions: The increased Moment-of-Inertia (MOI) will require increased slew times and more reaction-wheel torque but within acceptable margins, light-shield resonance may require minor changes to the control law, pointing jitter can be mitigated with the line-of-sight control of the secondary mirror if necessary, and the enhancement will add about approximately 3,500 pounds to the current HST weight budget of approximately 24,000 pounds which is about 15%.

The (MOI) will approximately double which will slow slews and require more reaction wheel momentum. The only significant resonance mode seems to be the light-shield up around 4 to 6 Hz which may require some control law changes. We can stand considerable jitter in the primary control system if we create a l0 Hz Line-of-Sight (LOS) control loop on the Secondary Mirror. Because the New Secondary Mirror (NSM) is so light (Small MOI), this loop should have almost no effect on the Primary loop.

A technical description of the model and results produced by it follow.

5.1.1. The HST10X Dynamic Model

The dynamic model of the HST was produced following the first Service Mission 1 (SM-1). It is in the form of mode shapes, natural frequencies, and damping. Several modifications were done to make it suitable for the HST10X enhancement. First, the modes from the Solar Array 2 (SA-2) were removed below 0.6 Hz. Above this frequency, coupling of the modes made it too difficult to identify individual modes of the SA-2 alone. The mass and moment of inertia of SA-2 could not be extracted without causing instabilities in the model, so it was left. Hence, the HST model is overweight by about 710 lb which is acceptable at this coarse level of simulation. Next, the Aperture Door (AD) and the forward 3.5-m section of the existing Light-Shield Assembly (LSA) were removed. This was done by eliminating the AD mode at 0.9 Hz and rigid body mass subtraction of 520 lb for the LSA and 105 lb for the AD.

Then, new parts were added to bring the model up to its current condition. This included addition of the Solar Array-3 (SA-3) wings (in their 00 orientation), the cooling system equipment Electronics Support Module (ESM) and NICMOS CryoCooler (NCC), and radiators for the NICMOS Cooling System (NCS) and Aft Shroud Cooling System (ASCS). This added over 2,547 lb to the aft bulkhead and aft shroud of the System Support Module (SSM). The resulting mass properties are shown in Table 5.1.1.

5.1.2. HST10X Primary and Secondary Mirror Assemblies

The example HST10X primary mirror is modeled as a thin mirror surface supported on actuators that are fixed to a contoured shell support structure. The mirror (2 mm thick) and light-weight actuators (50 per m²) are modeled as non-structural mass at 5.0 kg/m². The shell is 5" thick aluminum-core honeycomb with 0.060" Glass/Epoxy (G/E) face sheets. This thickness provides enough lamina to develop a minimum Coefficient-of-Thermal Expansion (CTE) structure, and is light weight and stiff. By careful design, the actual CTE can be kept to about 2x10^-7/^oC or lower. Radial stiffeners connect the shell to a central stiff ring. This ring is attached to the existing Optical Telescope Assembly (OTA) metering truss at four points modeled as ball joints. The New Primary Mirror (NPM) is modeled as hinged at two places to fit in the Orbiter cargo bay. It is assumed that the backup structure is tightly bolted along these hinge lines when deployed to eliminate slop and dead-band. The honeycomb plate, however, is modeled as discontinuous at the hinge lines.

A mast locates the New Secondary Mirror (NSM) relative to the primary. The mast of G/E is a 25" OD tube with 0.50" wall. Radial gussets connect it to the primary mirror structure. The four gussets are 0.65" thick in order to match the shading caused by the existing OTA secondary-mirror spider. They extend only part way up the mast because of packaging constraints when stowed in the Orbiter cargo bay. At some section, a bolted joint is assumed that is open when stowed. Otherwise, the mast is a single-piece structure without joints or subassemblies.

The New Secondary Mirror assembly, like the primary, is a thin mirror surface supported on about 230 small actuators that connect to a contoured honeycomb-plate. The mirror and actuators are modeled as 5.0 kg/m2 non-structural mass. A 50 lb lump is included to account for a yet to be designed sun blocking mechanism to serve as an aperture door. The honeycomb is 3" thick with 0.060" G/E face sheets and aluminum core. Six radial stiffeners connect it to the central mast.

Properties of the combined new primary-secondary assembly are shown in Table 5.1.1 for the mass, and Table 5.1.4 for the structural dynamics. Here, as elsewhere, modal damping of 0.5% of critical is assumed. Weight contingency of 20% is used on the structural mass portion. This is expected to be reasonably conservative.

5.1.3 HST10X Light-Shield Assembly

The HST10X New Light-Shield (NLS) is modeled as made entirely of deployable ABLE booms and Multi-Layer Insulation (MLI) blankets. In the model, three booms span out from a stiff central ring to the 9-m outer diameter at eight equally spaced locations. The central ring connects to the System Support Module (SSM) Forward Shell with four ball joints. Then, eight booms extend in the axial (V1 or X) direction for nine meters to support the light shield. The seven layer MLI (0.07 lb/ft2) covering is loosely connected so that its thermal deformation is largely de-coupled from the booms. The ABLE booms are standard items with beam-like properties. These properties are shown in Table 5.1.2

The modeled structural dynamic properties of the Light-Shield Assembly are listed in Table 5.1.6. It shows the first natural frequency is at only 0.166 Hz. A set of major interactive modes begins at about 1 Hz. Because of symmetry of bending, the first five modes have no coupling to the base point, as evidenced by their modal effective weight values of zero.

5.1.4 The HST10X-HST Modeled System

The subassemblies were combined by standard techniques. The HST10X primary/secondary mirror assembly in Craig-Bampton form was coupled to the OTA metering truss forward ring (at station 450.65), and the HST10X light-shield, also in Craig-Bampton form, was coupled to the SSM forward shell (at station 455.30). Since the HST (SSM and OTA) model was in the form of free system modes shapes, 564 modes were used up to 100 Hz to ensure convergence.

The free system was computed for 260 vibration modes to 16 Hz. The modal gains were calculated from this system. There are two types of gains of interest, one for the line-of-sight and one for the rate-gyro assemblies. For the line-of-sight, a ray trace equation was used that included the HST10X NPM and NSM, and the OTA Primary Mirror (PM) and Secondary Mirror (SM), and terminated at the OTA focal point. Because the HST10X mirrors are flexible, a paraboloid of revolution was fitted to their deformed shapes to determine the best-fit least-squares displacement in translation and rotation. The equations for the two modal gains, called K_B, are shown in Table 5.1.3

The natural frequencies and modal gains of the assembled system are listed in Table 5.1.4. This data is for the solar arrays in their 00 orientations. Time precludes studying other orientations. The first six modes are for the free system as a rigid body which, by definition, have modal gains that sum to unity. The other modes are for elastic motion. Only the few modes with significant gains are listed. Above six Hz the table has been truncated because it is the low frequency modes that are of most concern to the control system. Modal damping is assumed to be 0.5% of critical. Unfortunately, because of the nature of the models, there are no plots of the system mode shapes. Other than the mass contingency, no uncertainty factors have been applied to these results.

Section 5.1. Structural Dynamics Assessment -- tables
 

5.2. HST10X Preliminary Thermal Systems Assessment

Conclusions: In general, the results of the thermal assessment are that thermal effects will be small or can be mitigated easily, with the exception of the solar arrays. The solar-array operating temperature will be elevated by about 20 degrees C due to the space-view reduction from the New Light-Shield (NLS). This can be mitigated by the addition of supplemental solar arrays on the bottom of the NLS which will, at least partially, offset this temperature rise. Adding reflective surfaces to channel the heat back into space will provide additional relief. Making the new arrays wider than currently planned will further reduce the temperature rise.

One of the thermal issues remaining to be addressed is the temperature stability of the New Primary Mirror (NPM). It is expected that most of the science will be near the ecliptic plane resulting in about 75% of normal viewing conditions directing the earth's albedo inside the NLS every orbit. The existing 2.4-m mirror is very heavy and relatively stable thermally. The NPM will be only 2 mm thick so when the earth's albedo is directed on it, at about 30% of sun energy levels, the telescope, depending on pointing, may see 3,600 watts or so; which has to go somewhere. Probably most of it will go into the NLS because the NPM will reflect most of what hits it. If we install a few hundred localized computer controlled heaters on the mirror with enough power to replace the suns energy (maybe 400 watts total) the New Primary Mirror shouldn't move. HST10X will probably have an aperture door for safety reasons but closing and opening the extra mechanism almost every orbit is not desirable. There are coatings available that will reject the albedo wavelength and still provide good light-baffle performance in the visible. This is unlikely to solve the whole problem but will reduce its effect. In addition, if the final mirror choice is Zerodur material, extending the NLS length to 15 meters is likely to be adequate.

5.2.1. HST10X Thermal Assessment Assumptions

The radiation view-factor to space will be reduced by approximately 1% (analytical result). An increased absorbed flux of 3% is assumed. Further analysis is needed to determine the range of increased absorbed flux through the entire range of HST10X pointing.

The historical external-surface degradation is assumed through 2010. Impacts beyond 2010, although expected to be minimal, need to be addressed

It is assumed here that HST10X pointing will be limited to +V3 (i.e., HST sun angle will be no less than 90° from +V1) based on the shortened NLS of nine meters. The longer sun shield of about 15 meters to allow pointing up to 45 degrees of the sun is expected to have minimal impact on this analysis but will have to be studied.

5.2.2. AFT Shroud Impacts

The Aft Shroud uses radiation shields on the ±V2 surfaces to dissipate instrument power but the increased absorbed flux will impact the environment.. The ±V3 surfaces are blanketed due to solar and earth views.

Instrument environments (including FGS and WFPC) will increase by 1^°C to 1.5^°C. The impact to instrument components will be on the order of 1^°C. FHST temperatures will increase by 1^°C. The External Cooling System Radiator temperatures will increase by less than 5^°C.

5.2.3. Forward Section Assessment

There will be small temperature impacts to the System Support Module (SSM) Equipment Section (ES) Bays and OTA ES Bays. The forward sun-avoidance angle for the optics will result in warmer attitudes for the Bays, however, the bay doors are mostly insulated. The space-view factor reduction in this case is not large (near 2%).

There will be a greater temperature impact on the Solar Arrays. There will be up to 25% view of the HST10X New Light-Shield Assembly (V1) resulting in reduced space view. This increase is estimated to be 20^°C or more for the Solar Arrays for -V1 sun point. The increase will be less for Sun toward +V3.

Mast Damper temperature will increase for -V1 sun point. The anti-sun side radiator space view decrease will cause a warmer range. The +V3 High Gain Antenna TAG temperature will also increase - the dish could be moved to the shield if this is a problem.

5.2.4. Thermal Summary

1. Thermal impacts to the aft shroud are less than 5^°C.
2. Internal instrument environments are impacted less than 1.5^°C.
3. SSM Equipment Bay impacts are small due to insulation, low blockage, and forward sun-point limit.
4. Solar-Array wing impacts are estimated at +20^°C or more for some attitudes. If this remains a problem it may be further mitigated by operational restrictions.
5. Other external component temperatures will increase due to reflections/reduced space view.
6. Although HST10X will not pose any extraordinary environmental challenges, a more detailed analysis is required for the optics and structure thermal control requirements.

5.3 FGS Impacts from HST10X Enhancement

This assessment was to investigate the impacts on the Fine Guidance Sensors (FGS) if the current HST were augmented with an 8-m primary mirror and a suitable secondary mirror. The issues assessed were the FGS encoder resolution for guide-star acquisition, the field-of-view reduction, and the ability to maintain fine-lock with the amplified jitter as measured by the FGS.

Conclusions: First, the FGS will still acquire its target in the usual sequence of Search, Coarse Track, and Fine Lock. Second, the FGS field of view will be reduced from 69 square arcminutes to approximately 6.3 square arcminutes, while the faint end of the guide- star candidates improves from 14.5 to 17 v-magnitudes. Third, jitter as measured by the FGS, will be amplified by a factor of 3.3 and requires additional consideration. Finally, a trade-off study of how the much improved faint guide-star capability will offset the reduced FOV needs to be addressed by the astronomy community.

5.3.1 Optical Configuration Considered

1. Two additional parabolic mirrors, an 8.0-m primary and a 2.4-m secondary.
2. The stop is the current primary mirror.
3. Entrance diameter is 8.0 m.
4. The focal length is 9.576 m.
5. The separation of the mirrors is 6.7032 m).

5.3.2 Reduction in the FGS field of view (FOV)
versus acquisition of fainter guide stars

The plate scale for the new telescope, HST10X, is reduced by a factor of 3.3. As a direct consequence, the FGS FOV will be reduced from 69 square arcminutes to 6.3 square arcminutes. Although the FOV decreases, the light collecting characteristics of the larger telescope allows the FGS to acquire 17 mv guide stars, 2.5 mv dimmer than present operations.

5.3.3 Guide Star Acquisitions

The increased angular magnification of the FGS results in a narrower S-curve. The distance between the peak and valley of the current S-curve is approximately 40 milliarcseconds (mas). With HST10X, this distance becomes 12 mas; this is an extremely narrow S-curve for this application. There was concern, initially, that the uplinkable parameter, KD, the Fine-Lock step size could not be made small enough to acquire either the peak or the valley of the S-curve. However, parameters like the FGS walk down step size will also be reduced by the 3.3 factor. The number of LSBís that it takes to generate a 6.5 mas step size for the current HST/FGS system will result in a KD of approximately 2 mas in HST10X. Hence, the instantaneous FOV of the FGS can move in small enough increments to ensure that the Fine Lock acquisition algorithm will be successful. Using the same argument, Search and Coarse Track modes will also operate with the HST10X. While the FGSís will work with the HST10X, it is likely that K-factors and Port 1 commands will require calibration for the new system.

5.3.4 Jitter and the PCS/FGS system.

FGS guide-star acquisitions will succeed with HST10X, however, vehicle jitter will be amplified by a factor of 3.3 when observed with HST10X and the FGSís. This item will require additional study by Science, FGS, and PCS specialists.
 

5.4. Pointing and Control

5.4.1. Slew and Control System Performance

Conclusions: For HST10X, it appears the control system will provide acceptable science and slew performance with minor changes to the controller.

The Moment-of-Inertia (MOI) increase of approximately a factor of two will increase the torque required of the control system proportionally. Because the Reaction-Wheel torque is fixed, control bandwidth must be decreased by one over the square root of the MOI change. This will increase response time and slow maneuvers. The requirement to slew 90 degrees in 18 minutes will have to be relaxed unless actuator torque can be increased.

With the New Secondary Mirror steerable, and only moving the 2-mm thick mirror surface weighing only about 50 Lbs. instead of the full 350 Lbs. Secondary Mirror Assemble, the ratio between the moving mirror and HST10X is of the order of 1000 to 1 so with 100 mas of jitter, HST10X would only react by 0.1 mas. HST's current normal rms jitter is 3.5 mas or better. The existing control law may have to be modified to suppress any new flex modes.

The increased inertia will also require an increased Bright-Object Detection (BOD) angle and/or decrease of the maximum slew rate. This operating restriction can be addressed by installing new Bright Object Detectors to provide continuous bright object coverage.

The Gravity Gradient Torque will increase by a factor of four and Aerodynamic Drag Torque by 10. These and the MOI shift will result in generally higher Reaction Wheel Assembly (RWA) speeds. This is likely to not be a problem since HST operates the RWA's far below their rated conditions but this should be analyzed more closely. And finally, the Momentum Management System (magnetic torquer bars) may require upgrading.

5.4.2 Accelerated Orbit Decay

Conclusions: The larger New Light-Shield (NLS) and the solar maximum in 2010 combine to accelerate the rate of orbit decay to the extent that either stationkeeping or a re-boost mission will be required. An Ion-Engine Stationkeeping system is described in Section 5.5.

The minimum science floor altitude coupled with the maximum density (minimum altitude) allowable for a 90 degree slew in 50 minutes require that something must be done to keep HST above approximately 300 nautical miles altitude out beyond year 2012. One option is adding an Ion-Engine Stationkeeping System (IESS) and, aside from the added solar array power needed, this appears to be a reasonable solution for the problem. To accommodate the power required for the IESS, a combination of turning off unneeded systems to minimize power and adding solar arrays to the New Light-Shield should provide adequate power margin for the 4,500 watts required to operate the ion thrusters. It is prudent to add solar arrays to the New Light-Shield anyway because of the increased operating temperature so this should provide adequate power for intermittent boosting, approximately 10 days per year.
 

5.5. Mission Analysis for an HST10X Ion Engine Stationkeeping System

Conclusions: There is a readily available Stationkeeping System that can provide sufficient boost for stationkeeping, acceptable science efficiency versus boosting ratio, and adequate fuel margin for end-of-life ascent. And, there are convenient attachment points that are structurally acceptable. There are large diameter threaded holes in the existing HST trunnion sockets that will make ideal mechanical mounts for the ion thrusters

5.5.1. Stationkeeping Analysis Issues

5.5.1.1. Is stationkeeping needed?

The addition of a large mirror and light shield to HST will increase the atmospheric drag resulting in faster decay of the orbit. The increased rate of orbital decay is likely to be in the range of a factor of two to three depending on the size and shape of the enhancement. This increase will likely require an interim re-boost mission (or missions) unless another solution is implemented. This analysis was to consider such an alternate solution to the accelerated orbital decay issue. In addition, a stationkeeping system may save a de-orbiting mission which would more than pay for itself.

The Analysis Issues, Design, and Discussion below are based on a readily available Ion Engine Stationkeeping System (IESS) currently in use on other missions.

5.5.1.2. Is the boost rate adequate

Is the boost rate of the proposed Ion Engine Stationkeeping System adequate to offset the decay rate for any arbitrary point in the solar cycle?

Model calculations show that the orbital-decay rate varies up to a maximum of 80 m/day at the nominal altitude of 600 km for the worst solar flux level, 250 J, whereas the boost rate capability of the proposed IESS is calculated to be between 500 and 700 meters per day when operating in the non-optimal-thrust mode described below. This provides adequate capacity to overcome significant altitude decay in a short period of time.

5.5.1.3. Is Science Efficiency adequate

Is the frequency of thrusting sufficiently low to allow reasonable science efficiency?

To maintain the spacecraft within 1% of its initial 600 km altitude, roughly 16 maneuvers will be required over the solar cycle (surprisingly invariant irrespective of the epoch chosen). At a boost rate of 600 meters per day, roughly 10 days are required for each maneuver while the interval between maneuvers varies from 28 months at solar cycle minimum to slightly less than 4 months at solar cycle maximum, thus providing an observing efficiency against stationkeeping always greater than 90%. For the assumed epoch of 2006, the first maneuver will be required in 27 months. This lead time quickly diminishes as the solar maximum approaches. For an epoch of 2008, the first maneuver is required after eight months.

5.5.1.4. Is there enough fuel

Is the fuel requirement reasonable to accomplish both mission phases; Stationkeeping and end-of-life ascent?

The fuel consumption was computed to achieve the assumed end-of-life altitude of 1000 km, which provides of the order of 1000 years of orbital "lifetime". Total ascent phase is of the order of 600 days. Approximately 120 Kgm of fuel is needed and, with approximately 2 Kgm required for each stationkeeping maneuver during the operational phase, a total fuel requirement of 160 Kgm is required for an active mission of 10 years.

5.5.2. Stationkeeping System Design

5.5.2.1. Problem

How to prolong the spacecraft mission lifetime with accelerated orbital decay, and to dispose of it safely at end-of-mission?

5.5.2.2. Concept

Use low-thrust ion engines to provide periodic stationkeeping for the spacecraft and, at end-of-mission, to boost the spacecraft to a sufficiently high orbit to meet the mandated debris removal requirement.

5.5.2.3. Assumptions

Schatten 2-sigma solar flux values
Spacecraft: Mass = 12,475 Kg

Cross sectional area = 90.5 sq. m

5.5.2.4. Conceptual Design

The concept for the proposed IESS is a modular design including two outboard plus back-up engines, their power supply, four pressurized Xenon vessels, power and fuel switching subsystems, and associated plumbing and wiring.

The structure on which the components are mounted attaches to the FSS mounting brackets on the aft shroud. The large diameter threaded hole in the existing HST trunnion sockets make ideal mechanical mounts for the ion thrusters. These trunnions are located and aligned to the HST vehicle center of mass, are structurally stiff, and are not otherwise occupied. P601 is co-located with the starboard socket and circumference harnessing could route power to the other ion thrusters, bussed off of the first installation.

Approximately 4.5 kW are required to drive the engines at maximum thrust, however, the actual levels can be less. A non-optimized thrusting profile has been assumed to provide a worst case result. To simplify the analysis, the spacecraft is assumed to maintain a quasi-inertial attitude during thrusting rather than attempting to maintain the thrust vector normal to the orbital radius vector by a combination of slewing and rolling the spacecraft. The inertial attitude is such as to provide normal incidence for the solar arrays (appropriately rotated) in order to maximize their power generation. In this idealized, albeit inefficient case, one thruster is used at a time over 120 degrees of an orbit while the second operates for additional 120 orbital degrees. The computational model used allows no thrusting in either of the following conditions: 1) during eclipses in order to conserve battery power and 2) during any portion of the orbit when the scalar product of the thrust and velocity vectors becomes negative.