Glory is a NASA/GSFC Earth climate and atmospheric monitoring mission. The overall objective is to respond to the US Climate Change Science Program (CCSP) by continuing and improving upon NASA's research of the forcings influencing climate change in the atmosphere. Specific mission goals are:
• To collect data on the properties and distributions of aerosols in the Earth's atmosphere
• To collect data on solar irradiance for the long-term Earth climate record.
Background: The Glory project has its origin in the VCL (Vegetation Canopy Lidar) mission, representing the first project in the newly created ESSP (Earth System Science Pathfinder) program of NASA, and selected in the spring of 1997. However, NASA cancelled the VCL mission in 2001 due to unresolved technical problems in lidar instrument development and due to budgetary pressures. As a consequence, the VCL bus was put into controlled storage in 2001.
In the summer of 2003 (after the first Earth Observation Summit in Washington D.C.), NASA announced that it would fly a new instrument, APS (Aerosol Polarimeter Sensor), under development for the NPOESS series, on an earlier dedicated mission called Glory. The objective was to get early aerosol data for the study of the causes and consequences of climate change. The planning was to use the cancelled VCL spacecraft for the Glory mission. Later on, a second observation instrument, TIM (Total Irradiance Monitor), was added to the Glory mission. In early 2005, the dedicated Glory mission was cancelled. The new direction was to wait with the APS observations until the launch of the NPOESS series. Then in July 2005, NASA reversed course to move ahead with the Glory mission (due to further possible delays of the NPOESS series well beyond 2010). 1) 2) 3) 4) 5) 6) 7) 8)
Figure 1: Artist's view of the deployed Glory spacecraft (image credit: OSC, NASA)
The Glory mission uses the refurbished bus of the cancelled VCL satellite. The Glory spacecraft is being built and integrated by OSC (Orbital Sciences Corporation) of Dulles, VA, using the LeoStar-1 bus with deployable solar panels. The S/C structure is comprised of two octagonal cylinders of 1.1 m max. diameter. The bus structure is based on the LeoStar design of OSC with three elements: core module, propulsion module, and separation system. The S/C structure consists of modular Al honeycomb plates with stringers. Eight stringers connect the propulsion deck to the core deck and 8 shorter stringers attach to the core deck and extend upwards to support the ISA (Instrument Structure Assembly). Thermal control is provided by a passive radiation system with backup heaters.
The S/C is three-axis stabilized (zero momentum technique) providing 23 arcsecond (3σ) pointing knowledge and 0.1º (3σ) attitude control. The onboard hydrazine propulsion system is feeding four 4 N thrusters with 45 kg propellant. Electric power of 750 W is provided by GaAs solar arrays of 3.8 m2 and by NiH2 batteries of 30 Ah capacity. The solar arrays are bi-axially articulated. Use of AstroNav GPS receivers. The S/C design life is 3 years with a goal of 5 years. The S/C launch mass is 525 kg. 9) 10)
Orbit: Sun-synchronous near-circular orbit, nominal altitude = 705 km, inclination = 98.5º, period = 99 minutes, local equator crossing at 13:34 hours on the ascending node. The repeat cycle is 16 days.
The Glory spacecraft is going to be added to the A-Train (afternoon) constellation of NASA.
Figure 2: The A-Train constellation with Glory addition (image credit: NASA/GSFC) 11)
Figure 3: Instrument module assembly (image credit: OSC)
Figure 4: Isometric view of the Glory spacecraft (image credit: OSC, NASA)
Figure 5: The Glory spacecraft in launch configuration (image credit: NASA)
Figure 6: Alternate view of the Glory spacecraft and sensor accommodation (image credit: NASA)
RF communications: An S-band transponder and an X-band transmitter are employed (use of CCSDS packetized data protocol). Commercial ground stations provide the space-to-ground RF communications with the Glory satellite. The primary ground station is located in Alaska (Poker Flats) and the backup is located in Norway.
Table 1: Overview of mission parameters
Figure 7: Photo of the fully integrated Glory spacecraft (image credit: NASA)
Launch: The Glory spacecraft was launched on March 4, 2011 on a Taurus-XL vehicle (referred to as T9) of OSC from the Space Launch Complex at VAFB (Vandenberg Air Force Base), CA. 12)
Unfortunately, the launch of the Glory spacecraft, along with the secondary payloads, failed to reach orbit and ended in the Pacific Ocean. Telemetry indicated the fairing, the protective shell atop the Taurus XL rocket, did not separate as expected about three minutes after launch.
NASA has begun the process of creating a Mishap Investigation Board (MIB) to evaluate the cause of the failure. — The launch failure represents to NASA the loss of two important investigations related to climate change: ongoing data collection to monitor the sun’s energy reaching Earth, and a study of how aerosols move through Earth’s atmosphere and may influence climate. 13)
The mission was delayed from Nov. 2010 due to problems with its solar arrays.
Secondary payloads on this flight are the CubeSats:
- Explorer 1 PRIME (E1P) of Montana State University
- KySat-1 (Kentucky Satellite-1) of Kentucky Space
- HERMES of the University of Colorado at Boulder.
Mishap Causes and Recommendations (public release version):
On February 20, 2013, NASA released a summary report on the findings from a panel that investigated the 2011 crash of the Glory spacecraft after it failed to reach orbit on board an Orbital Sciences Taurus XL rocket (referred to as T9) , falling into the Pacific Ocean. Early on, the problem was traced to the fairing – the clamshell nosecone that encapsulates the satellite as it travels through the atmosphere — which did not separate from the rocket, weighing the satellite down, preventing its flight toward orbit. 16)
Summary of failure report: Using information and evidence gathered from interviews, technical interchange meetings, documents, data reviews, and testing, the T9 MIB (Mishap Investigation Board) developed a timeline of events, performed a detailed Fault Tree Analysis, and a Root Cause Analysis. From these analyses, the T9 MIB identified the proximate cause and possible intermediate causes for the failure. As a matter of explanation, a proximate cause, also known as the direct cause, is the event or condition that directly resulted in the occurrence of an undesired outcome. In this case, the proximate cause was that the fairing did not separate from the launch vehicle. An intermediate cause is an event or condition that created the proximate cause and that if eliminated or modified would have prevented the proximate cause from occurring. 17)
Proximate Cause: Fairing separation failed. Flight telemetry data strongly suggested that the payload fairing did not separate as planned. The T9 MIB hesitated to declare this as the sole and obvious proximate cause and, instead, elected to conduct a comprehensive investigation of potential causes. Comparison of modeled structural behavior based on analysis and ground test to flight data provided through telemetry showed the only reasonable scenario was that the fairing system failed to separate. The detailed analysis pointed to a failure to fracture near the forward end of one of the fairing side rails, which prevented full separation. The extra pressure sensor added to the installation of the cold gas pressurization system, in place of the hot gas generator system as part of the response to the T8 MIB report, provided valuable data to the T9 MIB. This sensor data verified the cold gas pressurization system performance was satisfactory and that the T9 payload fairing’s base ring indeed had separated. As a result, the T9 MIB could eliminate the cold gas pressurization system and the base ring as a cause and focus on the scenario in which the forward end of the payload fairing side rail failed to fracture.
Root Cause Analysis: While the T9 MIB was able to identify the proximate cause and two possible intermediate causes for the T9 mishap, they were unable to identify the root cause for this failure. As a matter of explanation, an intermediate cause is between the proximate cause and the root cause in the causal chain. The root cause is the factor or set of factors that contributes to, or creates the proximate cause. Typically multiple root causes contribute to an undesired outcome.
The T9 MIB was unable to determine a root cause for the mishap mainly due to limited flight telemetry and the inability to recover the payload fairing hardware for analysis that would have enabled the determination of a definitive intermediate cause or causes. While the T9 MIB was unable to identify a root cause, they made several technical observations and findings which are summarized below:
• The T9 MIB determined that side rail charge holder slumping (compression) could possibly occur because of the following:
1) The Side Rail assembly’s susceptibility to (and the effects of) charge holder slump was not previously identified.
2) The temporal distribution of acceleration, vibration, shock environments had changed over time.
3) Orbital’s fairing joint buildup process variability could affect charge holder slumping susceptibility.
• In addition, the T9 MIB noted a large percentage of potential causes rated as “possible, but highly unlikely” involved frangible joint components and also observed that Orbital's manufacturing processes were not as tightly controlled as those applied by NASA in other pyrotechnic hardware designs. The possibility exists that manufacturing process controls could allow variation in material properties and hardware dimensions that may impact system performance.
• The T9 MIB did not find evidence that a detailed failure analysis of the frangible joint design used on the Taurus was performed at any point in its life cycle. Details of the design had evolved since the genesis of the base ring application for the Pegasus launch vehicle, and the effects of those evolutionary design changes, including the potential for charge holder slump, were not discovered.
• The T9 MIB also discovered that the qualification activity for the frangible joint system was generally performed at the subscale level, using industry practices for the qualification of pyrotechnic devices. The effects of all flight environments, either individual or combined, were not always considered. As with the system design, the flight environments also have evolved over time, and the effects of these changes on system performance margins should be understood. As a result of the analyses performed, the T9 MIB made the following technical recommendations:
1) Orbital should establish frangible joint system manufacturing process controls sufficient to assure that variability in materials properties and hardware component dimensions, within both maximum and minimum tolerances, will not invalidate design performance requirements.
2) An extensive failure analysis (for example, detailed fault tree or failure mode analysis) of the Taurus frangible joint design should be performed.
3) Design and implement a qualification and test activity for the Taurus frangible joint system based on the results of an extensive failure analysis (for example, detailed fault tree or failure mode analysis) and with consideration for the environments in which the joint is operated. “The MIB believes that if this recommendation were implemented, it could address all the possible frangible failure scenarios identified in the investigation.”
During the course of the mishap investigation, additional observations and recommendations related to Agency policy, special assessment procedures, and improved communications were noted by the T9 MIB. Although these items were not deemed causal to the launch failure by the T9 MIB, the MIB determined that they could be beneficial for future programs.
Sensor complement: (APS, CC, TIM)
APS (Aerosol Polarimeter Sensor):
APS is of the NPOESS series heritage. The instrument is a multi-spectral polarimetric sensor (spectrophotopolarimeter) with the capability to collect visible, near-infrared, and short-wave infrared (VNIR, SWIR) polarized radiometric data scattered from aerosols and clouds. The instrument is being developed by Raytheon SBRS. The overall objective is to retrieve aerosol and cloud parameters (operational products) for climate research using multispectral photopolarimetry (support for model development). The parameters to be measured are the along-track and multi-angle scene intensity as a function of wavelength and polarization to determine: aerosol optical thickness and aerosol particle size, aerosol refractive index, aerosol single-scattering albedo, aerosol shape (sphericity), and cloud particle size distribution.
In the framework of the Climate Change Research Initiative (CCRI), initiated in June 2001 to study areas of uncertainty about global climate change, research on atmospheric concentrations and effects of aerosols was specifically identified as a top priority. 18) 19) 20) 21)
Figure 8: Illustration of the APS instrument (image credit: NASA)
4) The global distribution of natural and anthropogenic aerosols (black carbons, sulfates, etc.) with an accuracy and coverage sufficient for reliable quantification of: a) the aerosol effect on climate, b) the anthropogenic component of the aerosol effect, and c) the potential regional trends in natural and anthropogenic aerosols.
5) The direct impact of aerosols on the radiation budget and its natural and anthropogenic components
6) The effect of aerosols on clouds (microphysics and coverage) and its natural and anthropogenic components
7) Investigate the feasibility of improved techniques for the measurement of black carbon and dust absorption to provide more accurate estimates of their contribution to the climate forcing.
The APS design is of RSP (Research Scanning Polarimeter) heritage, an airborne instrument of NASA, which laid the foundation for APS by successfully retrieving aerosol properties based on radiance in orthogonal polarizations at multiple wavelengths in the 0.4 to 2.4 µm region and demonstrated multi-angle along-track polarization measurements required of APS (polarization accuracy in the 0.2% range).
The APS measurement approach required to ensure high accuracy in polarimetric observations employs Wollaston prisms to make simultaneous measurements of orthogonal intensity components from exactly the same scene as shown in Figure 9. The field stop constrains the APS instantaneous field of view (IFOV) to 8 ± 0.4 mrad which, at the nominal A-Train altitude of 705 km, yields a geometric IFOV of 5.6 km at nadir. The spatial field is defined by the relay telescope and is collimated prior to the polarization separation provided by the Wollaston prism. This method guarantees that the measured orthogonal polarization states come from the same scene at the same time and allows the required polarimetric accuracy of 0.2% be attained.
To measure the Stokes parameters that define the state of linear polarization (I, Q, and U), APS employs a pair of telescopes with one telescope measuring I and Q and the other telescope measuring I and U (see Figure 11). This provides a redundant measurement set that increases the reliability of APS. The broad spectral range of APS is provided by dichroic beam splitters and interference filters that define nine spectral channels centered at the wavelengths λ = 410, 443, 555, 670, 865, 910, 1370, 1610 and 2200 nm. Silicon detectors are being used in the VNIR bands, while HgCdTe detectors, passively cooled to 160 K, are used in the SWIR bands.
Figure 9: RSP optical approach for polarization measurement adopted for APS (image credit: NASA)
Legend of Figure 9: The red markings show the orientations of the optical axes of the birefringent crystals forming the Wollaston prism. The orange lines show ray paths undergoing the split into orthogonal polarizations as indicated by the green and blue lines.
The critical ability to view a scene from multiple angles is provided by scanning the APS IFOV along the spacecraft ground track (Figure 10) with a rotation rate of 40.7 rpm with angular samples acquired every 8 ± 0.4 mrad, thereby yielding ~250 scattering angles per scene. The polarization-compensated scanner assembly includes a pair of matched mirrors operating in an orthogonal configuration and has been demonstrated to yield instrumental polarization < 0.05%. From the nominal A-Train altitude, the APS viewing angle range at the Earth is +60º / -80º with respect to nadir.
The scanner assembly also allows a set of calibrators to be viewed on the side of the scan rotation opposite to the Earth (Figure 11). The APS onboard references provide comprehensive tracking of polarimetric calibration throughout each orbit, while radiometric stability is tracked monthly to ensure that the aerosol and cloud retrieval products are stable over the period of the mission.
Figure 10: Along-track multi-angle APS measurements via 360º scanning from Glory orbit (image credit: NASA)
The APS provides scene radiance in orthogonal polarizations (Stokes parameters I, Q and U simultaneously) in nine spectral bands over a large range of viewing angles with the sensor scanning field-of-view (FOV) ±60º about nadir. APS performs an along-track scan to obtain its observational data for EDR (Environmental Data Record) generation. The instrument design comprises a polarimeter, electronics modules, and Raytheon's RTA (Rotating Telescope Assembly). 27) 28)
Even though the APS operates continuously, it nominally collects data only during sunlight, when the solar zenith angle is > 70º; no data is retrieved in the eclipse phase of the orbit. Approximately 490 Mbit of data is being collected per orbit.
Table 2: Overview of the APS instrument parameters
Table 3: APS spectral bands and required accuracies
Figure 11: The APS functional block diagram (image credit: NASA)
Figure 12: Components of the APS (image credit: NASA)
CC (Cloud Camera):
The instrument is designed and developed at BATC (Ball Aerospace Technology Corporation), Boulder, CO. The objective is to use two high spatial-resolution cloud cameras in conjunction with APS on the nadir-pointing side of the platform to assist in differentiating between clouds and aerosols. Over the ocean, the cloud camera will be used to determine aerosol load and fine mode fraction based on the aerosol microphysical model determined from APS measurements.
The cloud camera assembly consists of dual-band, visible imagers (412 & 865 nm). A non-scanning staring detector array is used that is analogous to a star tracker, but Earth-viewing. It consists of an optical imaging system that provides continuous cross-track coverage over a narrow swath (±125 km) centered on the APS along-track footprint. - The effective collection time is 44% of an orbit (~45 min). Approximately 584 Mbit of data (compressed onboard) is collected per orbit.
Table 4: Overview of the CC instrument parameters
Figure 13: Cross-section of the cloud camera assembly (image credit: NASA)
Figure 14: Photo of the Cloud Camera (image credit: NASA)
TIM (Total Irradiance Monitor):
The instrument is designed by LASP (Laboratory for Atmospheric and Space Physics) of the University of Colorado in Boulder, CO (TIM is of SORCE mission heritage, launched Jan. 25, 2003). The continued measurement of the TSI (Total Solar Irradiance) to determine the sun's direct and indirect effects on Earth's climate, at current state-of-the-art accuracy and without temporal gaps in the dataset, constitutes the solar irradiance requirement for the Glory mission and the objective of the TIM instrument. Note: The parameter TSI is also referred to as the “solar constant.”
TIM is an absolute active cavity radiometer of ACRIM heritage, but with significant improvements in sensor and electrical design, particularly in thermal control. It measures the TSI in SI units (Systeme International d'Units) with an absolute accuracy of 0.03% and relative accuracy of 0.001% per year. The complete wavelength range is covered. 29) 30) 31) 32) 33) 34)
Figure 15: Photo of the TIM instrument (image credit: LASP)
Figure 16: Schematic of an ESR (Electrical Substitution Radiometer) design (image credit: LASP)
The TIM instrument design features a cylindrical housing assembly containing four identical, right circular, 15º apex angle, conical cavity detectors cantilevered from a single support. The mount provides thermal impedance between the cavities and the heat sink which surrounds them. One cone, the measurement cavity, views the total solar disk through a precision aperture, and any of the other cones may be used as the reference cavity. The cones are made from 1 mm thick electro-deposited 99.99% pure silver. The cone interiors and the heat sink are coated with a Ni-P black and etched to produce a diffuse black surface which has a nominal visible reflectance of 0.3%.
Figure 17: Top and bottom view of cone housing assembly (image credit: LASP)
Figure 18: Illustration of a single conical cavity detector (image credit: LASP)
Six miniature (1 mm2) diamond-substrate chip heaters are soldered to the back of each cone. Gold electrode spinel chip thermistors mounted between the apex of each cone and its support structure are used as sensors for the electrical bridge circuit which controls the measurement cavity temperature. Baffles, located behind the primary aperture, shield the measurement cavity from earth albedo and off axis infrared radiation. Small Si and InAs photodiodes, located just behind the heat sink, view the measurement cavity at an oblique angle to measure the change in its reflectance. The heat sink is thermally isolated from the TIM channel case and its temperature is controlled to 0.001ºC.
The ESR (Electrical Substitution Radiometer) of TIM has dual bolometers with autobalancing through an AC, digital, feedback loop.
TIM incorporates four cavities (cones) and adheres to the basic concepts of Electrical Substitution Radiometers (ESRs), but employs modern, state-of-the-art electronics and materials. The four cavities provide multiple redundancy and added duty-cycling capability. TIM looks at the sun every spacecraft orbit. Each measurement consists of multiple samples taken over the course of a single orbit, providing 15 measurements per day. The TIM shutters, one for each radiometer, are driven by brushless torque motors. In its normal operational mode, the TIM shutter is cycled 50% open and 50% closed every 100 seconds throughout the orbit. Periodic (approximately once per week) field of view maps are obtained by offset pointing the S/C by ±15º about the sun vector. - Instrument mass = 7.9 kg, power = 14 W (average), size = 17.7 cm x 27.9 cm x 27.2 cm (H x W x D), nominal data rate = 539 bit/s.
Figure 19: Cutaway view of the TIM instrument (image credit: LASP)
Figure 20: The TSI record provided by various spaceborne instruments over a period of 25 years (image credit: BAMS)
Mission operations and control is being provided by the MOC (Mission Operations Center) of OSC, Dulles, VA. The APS SOC (Science Operations Center) is located at NASA/GISS (Goddard Institute for Space Studies), a NASA/GSFC facility at Columbia University (New York), while the TIM SOC is located at LASP (Laboratory for Atmospheric and Space Physics) of the University of Colorado in Boulder, CO. The NASA IONet (Internet protocol Operational Network) communication is being used between the varies elements of the ground segment.
Figure 21: Overview of the Glory mission elements (image credit: NASA)
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates.