Deep Space One Telecommunication Development M.I

• • SSC98-IV-l • Deep Space One Telecommunication Development M.I. Herman, S. Valas, W. Hatch, C.C. Chen, S. H. Zingales, R. P. Scaramastra, L.R. Amaro, • and M. D. Rayman Jet Propulsion Laboratory • California Institute of Technology 4800 Oak Grove Drive • Pasadena, CA 91109 m1s 161-213 818-354-8541 • martin.i.herman@jpl.nasa.gov • Abstract. Deep Space One (DS1) is the first of the New Millennium Program deep space technology validation missions, to be launched October 1998. This paper focuses on the • Telecommunication Subsystem architecture, technology developments, as well as the test results. Technical factors that influenced the subsystem architecture were the ability to command the • spacecraft and downlink telemetry data in cruise and emergency situations, and the need to provide radiometric data. Additional challenges included the requirement to demonstrate new • telecommunication technology, enable the validation of other system technologies (for example solar electric propulsion, autonomous navigation, and beacon monitor operation), and at the same time utilize a single string system design. From a programmatic perspective we had to accomplish • these goals within a budget and workforce load that was at least a factor 2 less than the Mars • Pathfinder Project. The Small Deep Space Transponder (SDST), a new technology developed by Motorola, is the • heart of the Telecommunication Subsystem and is a result of a JPL multimission sponsored competitive award. The SDST provides the functionality normally associated with 4-5 individual • subassemblies at less than half the mass (2.95 kg). Another new technology to be validated on DS1 is a 2.5W Ka-band solid state amplifier developed by Lockheed Martin (under their own funding). This technology not only extends the robustness of the system design (augmenting the • X-band downlink) but also provides the capability to characterize ~-band deep space • communication links. • Infusion of new technologies did not allow for the traditional subsystem integration and test program due to the additional lead time required for the technology developments. Working with the Project, we defined a test plan that was consistent with the spacecraft integration timeline while • still providing enough characterization to ensure confidence of the subsystem's functionality in • flight. Introductionl programs affordable, it is anticipated that • small spacecraft, launched on low-cost launch vehicles and with highly focused • Overview of New Millennium objectives, will be used for many of the missions. To prevent the loss of capability • that may be expected in making spacecraft "NASA's plans for its space and Earth smaller and less expensive, the introduction science programs in the early years of the of new technologies is required. • next century call for many exciting, scien­ • tifically compelling missions. To make such 1 • M. Hennan 12th AIAAlUSU Conference on Small Satellites • SSC98-IV-l •

With many spacecraft carrying out its Astro, Inc. as the industry partner for programs of scientific exploration, NASA spacecraft development. Planned for launch • would accept a higher risk per spacecraft; the in October 1998, DS l' s payload consists of loss of anyone spacecraft would represent a 12 technologies. The primary requirements of relatively small loss to the program. the mission are the validations of four of • Nevertheless, the use of new technologies in these: space science missions forces the first users • solar electric propulsion (SEP), • to incur higher costs and risks. The implemented on DS 1 as the ion concomitant diversion of project resources propulsion system (IPS); from the focused objectives of the science • solar concentrator arrays, supplied • missions can be avoided by certification of to DS 1 as solar conCentrator arrays the technologies in a separate effort. using refractive linear element • technology (SCARLET); • autonomous on-board navigation • The principal goal of the New Millennium (AutoNav); and program (NMP) is to validate selected high­ • an integrated panchromatic visible risk, high-benefit technologies in order to imager and infrared and ultraviolet • reduce the risks and the costs future missions imaging spectrometers, would experience in their use. NMP implemented on DS 1 as the • comprises dedicated deep space and Earth miniature integrated camera orbiting missions focused on the validation of spectrometer (MICAS). these technologies. As each mission is • flown, the risk of using the technologies that formed its payload should be substantially To assist in the validation of SEP, DS 1 • reduced, both because of the knowledge includes IPS diagnostic sensors (IDS), gained in the incorporation of the new composed of instruments to quantify • capability into the spacecraft, ground system, magnetic and electric fields, ion and electron and mission design as well as, of course, the densities, and surface contamination. quantification of the performance during the • mission. DS 1 also has level 1 goals which include the • validation of eight more technologies: By their very nature, NMP missions are high • a small deep space transponder risk. The key technologies that form the (SDST); • basis for each mission are the ones which • a Ka-band solid state power require validation to reduce the risk of future amplifier (KAPA) and associated • missions. Still, the failure of a new experiments in ~ -band technology on an NMP mission, even if it telecommunications; • leads to the loss of the spacecraft, does not • an integrated ion and electron necessarily mean the mission is a failure. If spectrometer, known as plasma the nature of the problem with the technology experiment for planetary • can be diagnosed, the goal of preventing exploration (PEPE); future missions from accommodating the risk • a remote agent experiment (RAX) • can be realized. Showing that a technology architecture for autonomous needs modification before it is appropriate for onboard planning and execution; use on science missions would be a useful • a beacon monitor operations • result of an NMP flight." experiment (BMOX) for autonomous onboard health and • status summarization and request Overview of DSl1 for ground assistance; • • a set of low power electronics (LPE); "Deep Space 1 (DS 1) is the first mission of • a high-packaging-density smart • NMP, It is being led by JPL, with Spectrum power switch, known as a power • 2 1 M. Herman 12 " AIAAIUSU Conference on Small Satellites • • • • SSC98-IV-l activation and switching module Table 1. New Millennium Deep Space 1 • (PASM); and Telecom Hardware Summary • a multifunctional structure (MFS) experiment combining electronics DC Power Notes ITEM Mass Dissipation • and thermal control in a structural (kg) (W) • element. SDST 2.95 New .... ····T1:·6 ...... ~~~~g,~lY·· .. ' .. ~.. 9.~.. y. .... Details on each technology (and further ...... 1'3':'3...... RX, X- • background on the project) are presented Exciter elsewhere." 2.3 ...... ·1·5:·2····· .. · • "RX~"'K~'''~' ...... T6:·9·· .... ·· Exciter The level 1 goals specify that DS 1 encounter "RX~"""'X~' • & Ka - an asteroid (1992 KD). In one extended mission scenario, DS 1 could encounter the Exciters • comet Borrelly. X-SSPA 1.61 59 MPF Spare • Ka-SSPA 0.66 20 New DSI Telecom Architecture Design Diplexer 0.37 MPF • Spare The telecommunication subsystem for DS 1 is DAM 0.20 0.30 New • single string (see Figure 1) as mandated by Design the project. The primary communication link WTS 0.78 Cassinil • is on Channel 19 at X -band for both uplink (2 of MPF and downlink ( 7.168 GHz and 8.421 GHz, them) Heritage respectively). As part of the technology MiscHW 5.25 (wiring • demonstration we have an auxiliary Ka -band harness, downlink (32.155 GHz). A major departure wvgd, • from a science driven mission to a technology brackets, driven one is that we are capabilities versus J..lwave requirement driven. This played a major role • comps in how components were selected as well as etc.) what our communication capabilities would • be. RFE 11.82 TOTAL

• MPF Table 1 presents a detailed summary of the HGA 1.2 telecommunication subsystem mass and Spare • power for the NM DS 1 mission. LGA L08 Cassinil (3 of MPF • them) HeritaKe HGA 0.04 New • Radome Design Ka-band 0.80 New Hom Design • Antenna 3.11 • TOTAL TOTAL 14.9 Allocated • TELECOM mass= • 16.5 kg 3 • M. Herman 12th AIAAlUSU Conference on Small Satellites • SSC98-IV-l •

Radio Frequency Electronics (RFE) This single unit which replaces 4 components used in past designs provides new • capabilities never considered before as part of The Radio Frequency Electronics (includes a transponder. Details about this unit will be all components except the antennas) was a described later in this paper. • combination of new technology insertion and the use of low-cost space heritage • components. From the Mars Pathfinder Without a functional SDST, the spacecraft (MPF) program we procured a flight spare would not be able to communicate with High Gain Antenna (HGA), diplexer, and Earth. In addition, other key DS 1 • 12.5 W RF X-band Solid State Power technologies depend on the SDST Amplifier (XPA). Implicit in this strategy functionality and include: • was that MPF would not require their spare hardware before launch in December 1996. • Ka -band communications: The SDST • Following the MPF launch, if DSI provides the Ka -band RF drive as experienced any failures (of these well as the modulation/encoding and components) there would not be any spares ranging for Ka -band. • to replace them. In addition, many other critical components on the spacecraft did not • Beacon Monitor: The SDST can • have spares readily available. Another generate the beacon tones. Each tone strategy to decrease cost and delivery time represents a desired ground response was exemplified in our procurement of (e.g. contact sic immediately, • waveguide transfer switches (WTS). Our everything is nominal, had problem statement of work mirrored Cassini' s but all is nominal, have data to send • requirements. This enabled the vendor to use in the next 2 weeks) established high reliability processes which • were more than adequate for DS 1 and • Autonomous Optical Navigation: The allowed for quick low-cost tum around of the radiometrics from the SDST are units. critical in the calibration and • verification of the optical navigation experiment. • A major challenge for our team was to develop a subsystem around a transponder that was concurrently under development Many components in the RFE have dual • with the Radio Frequency Electronics. roles. As an example, the Ka -band Solid Although the use of existing space-qualified State Power Amplifier (KAPA) is a • parts would seem logical and low risk, the technology demonstration which by itself is diplexer was a major area of concern. The important. It also provides a secondary • diplexer's function is to allow a single downlink to earth in the event that we lose physical connection at the antenna interface our prime X-band link. while separating the uplink from downlink. It • also provides the necessary 100 dB isolation required to prevent the transponder from Also included as part of the technology • locking to a spurious signal coming out the validation of the Ka -band SSPA are two RF XPA. By working with the vendor and detectors at the input and output. Even performing detailed analysis, we concluded without a Ka -band link to Earth, we can • that the MPF spare diplexer would work. monitor the health of the KAPA through This was, of course, verified during telemetry on the X -band downlink. In the • subsystem integration and test. event that we observe a degradation on the Ka-band downlink, the detectors will play an important role in diagnosis. We will have the • The heart of the RFE is the Small Deep Space ability to determine if the SDST Ka -band Transponder (SDST) which was one of the exciter had any degradation in performance or • major technologies to be validated on DS 1. if a failure occurred between KAPA output • 4 M. Hennan 12th AIAAlUSU Conference on Small Satellites • • • • SSC98-IV-l detector and the Ka-band horn antenna. This Antenna • information is extremely useful for the fault protection algorithms. The antenna portion of the Telecom '. subsystem has been greatly affected by A subtle feature of the Telecom architecture is limited resources and has impacted what our • that a majority of our telemetry flows through downlink capability performance will be. As the SDST versus being monitored directly by mentioned above, the HGA X -band antenna the spacecraft. The ability to have our own is a flight spare from MPF. This unit has a • temperature sensors and analog to digital physical aperture of 28 cm. The gain for the converters allows us the ability to provide a HGA was 20.1 dBi and 24.6 dBi for uplink • self-contained subsystem with engineering and downlink (right hand circularly telemetry. Therefore, analog signals from polarized), respectively. RF detectors monitoring the X-and K -band • • a excIter output power levels and the KAP A output power level are fed into a new The HGA is a microstrip array fabricated on • component developed for this mission, the duroid. Based on the spacecraft mechanical Detector Amplifier Module (DAM). It configuration, the HGA would be facing the • amplifies, buffers, and transmits the detector sun a majority of the time and the low signals along with a DAM health monitor emissivity of the material would create a voltage to the SDST which then does an thermal problem. A thermal radome was • analog to digital conversion and incorporates therefore needed. The solution was to place a the data onto its 1553 bus engineering status beta cloth cover around the HGA itself. This • word. This telemetry has proven invaluable material is normally used in the thermal in the course of subsystem functional testing. subsystem; however, at X-band the loss of • the material is approximately 0.2 dB and the gain numbers quoted above did include the Finally, as mentioned above, the X -band cover affects. The beta cloth option was the • SSPA (developed at IPL) is a 12.5 WRF unit lowest-cost, lowest mass, highest-reliability with 22% power added efficiency and a gain solution to this problem. However, no matter • of 38 dB. This unit has built in temperature how we solved the radome problem we compensation and redundant power supplies. needed to address the issue of electrostatic Since the output power of the SSPA is a charge buildup on the surface of the beta • function of the bias voltage applied, this cloth. Based on a worst case discharge design permits an RF output as high as 20W scenario and the fact that no sensitive • (with similar efficiency). Since a fully electronic components are near the HGA, the qualified MPF unit with 12.5WRF was risk was minimal for the mission. • available and the prime power was a limited project resource, we decided to utilize the unit as originally built without modification. For initial acqUISItIon and emergency • coverage we integrated the Cassini waveguide low gain antenna (LGA) design • Figures 2 - 4 show the flight panels with the with commercial polarizers and circular-to­ telecom Radio Frequency Electronics rectangular WR112 adapters used on MPF. integrated onto them. In a following section Boresite gain is 8.8 dBi and a 3 dB • we will describe the rationale of how the beamwidth of 70°. For this mission we will components were distributed on the employ three LGAs (±Z and +X direction, • spacecraft bus. where the HGA faces +X and the top of the • spacecraft is +Z, see Figure 5). The validation of the Ka-band SSPA in flight • is a prime objective; however, it is extremely • valuable to demonstrate a fully functional Ka- 5 • M. Hennan 12th AIAAlUSU Conference on Small Satellites • • SSC98-IV-l • band telemetry and radiometric link in this New Technology Impact mission. Therefore, we were challenged to • implement an antenna which was low-mass, flight worthy (not requiring substantial Two new technologies on DS 1 have very modification for flight acceptance), low-cost, significant impact on the Telecom subsystem • and provides substantial gain over an LGA architecture. The first is the solar electric The solution was a corrugated waveguide propulsion system (IPS). The most obvious • horn antenna with a gain of 28 dBi (right­ concern is potential EMI effects from the hand-circularly polarized). thruster and its associated power electronics.4 • Many ion propulsion systems in the past have been compromised by this. Throughout Supported Mission Data Rates solar-thermal vacuum testing (the only time in • which the telecom and SEP operated in a flight like environment) we carefully • The most important telemetry data for this monitored our subsystem performance. We mission are engineering and health status of were looking for degradation of receiver the spacecraft. Therefore, extremely high data carrier threshold and false lock conditions. • rates, although desirable, are not essential for No effects from the SEP were observed. a technology validation mission (as stated • previously, we are capabilities driven versus requirements driven). Technologists must be Plasma effects from the SEP on • able to obtain enough data in order to communications is the next area of concern. diagnose any potential problems which the For Ka -band the effects are hardly noticeable. new technology may present during the At X -band, losses were measured to be less • course of the mission. than 0.2 dB. Static testing indicates that we should not expect to see any major effects on • the link. We have not characterized a weak Due to limited project resources (mass, signal through a plasma. For prime power, and funding), the spare X-band MPF communication passes, the HGA will be • transmitter and high gain antenna predefine pointed at Earth and the engine will not be what our maximum EIRP would be. For near thrusting. However, during thrusting we do • earth communication we will support data plan to have the ability to communicate rates as high as 19.9 kbps (even through the through the LGAs and at this time we do not • low gain antenna). Keeping a conservative 3 foresee any problems. dB link margin for the asteroid flyby we can support -700 bps using a 34m ground • station, 2800 bps using a 70m ground Communications while thrusting has a major network. At the comet Borrelly the data rates impact on the Telecom architecture. It is not • which can be support via the HGA are -400 for the reasons listed above but a very bps and 1600 bps for the 34m and 70m practical one: Earth-Spacecraft-Thruster ground networks, respectively. angle. Up to now the LGAs as stated were • for initial acquisition and emergency. Now a new set of angles must be considered when • The downlink data rate capability of the Ka­ selecting the number of LGAs required and band link (2 dB link margin) is 40 bps at the where they are placed on the spacecraft. For • asteroid using a 34m ground station. For the DS 1, the -Z LGA is for initial acquisition and comet, the data rate goes down to 10 bps. As near Earth communications. For a majority of stated earlier the Ka -band links provide an the mission the +X LGA will be used to • emergency backup for X-band. communicate through while thrusting. For periods where the +X LGA will not be • sufficient we can employ the +Z LGA To complicate matters the trajectory selection process is dynamic and being redefined as • this paper is written. The earlier the • 6 M. Herman 12th AIAAJUSU Conference on Small Satellites • • • • SSC98-IV-l trajectory can be defined and finalized, the all subsystem components was to place them • better the chance that a good LOA strategy on the outside of the spacecraft structure. The can be implemented. propellant tanks and a portion of the telecom electronics were allowed to reside inside the • bus. This was the only way which we could Autonomous optical navigation has presented implement our subsystem with minimal RF • a new challenge for the Telecom subsystem losses and minimal mass. in two ways. The first (which affect\) the entire spacecraft) is that an op-nav maneuver • may occur as often as once every other day The antenna locations were continually (for DS 1, the average will be once per week). repositioned to the point where they almost • This means that the spacecraft wi1l slew were not attached to the spacecraft. around to look at targets and then slew back Originally, the HOA was to be mounted in • to its regular position. During this the center of the +X panel. However, due to approximate 2 hour period the spacecraft will mass properties requirements, the batteries be thermally cycled. Over the course of a were moved from the back to the front of the • 10.5 month prime mission, this translates to spacecraft on the +X panel. This in turn over 42 cycles. This is not common for most forced the X -band SSP A radiator area to • deep space missions and can result in extra move to the top half of the+X paneL The reliability requirements for future hardware HOA was forced to the top of the spacecraft whose mission durations are longer and and the field of view of the +X LOA was • whose op-nav maneuvers are more frequent. compromised for the early portion of the The second area is that when we try to mission (this was part of the justification of • communicate with a spacecraft just coming adding the -Z LOA). Naturally, all this was out of an op-nav maneuver, the subsystem is occurring as we were trying to implement our • not at thermal equilibrium. This makes baseline design. By the time the spacecraft thermal predicts difficult and creates a bus structure was finally defined we had only challenge for locking up the receiver at the 9 months to design, fabricate, and test the • proper frequency. For DS 1 we have Telecom subsystem. characterized the best lock frequency (BLF) • as a function of temperature and then developed a ground sweep strategy that Due to the fact that the HOA was now located would guarantee that we would sweep on the +Z panel of the spacecraft, it had to • through BLF and lock the uplink. We have stay outside the field of view (FOY) of found for the SDST that for a range of ±30 PEPE. It also had to stay outside the FOY of • kHz and sweep rate of 250 Hz/sec we can the +X LOA. This forced the +X LOA always lock the receiver. towards the + Y axis which in turn pushed the • front portion of the Ka-band horn off the bus. We had to design a special bracketing scheme Mechanical Implementation that went over 2 panels just to support the • Ka-band horn. • As shown in Figures 2-4, the telecom RFE was placed on two separate spacecraft Without careful consideration, mechanical panels. Naturally, our goal from an electrical and thermal requirements can severely impact • perspective was to minimize RF losses by the telecom performance. placing components as close together as • possible. The thermal design was a passive radiative approach which required significant New Technology • amount of panel space for radiation. This necessitated putting the X-band SSPA on the + X panel while the rest of the radio The Pluto Fast Flyby Advanced technology • frequency electronics went on the +X,+ Y Initiative5 (A TI) was the first step toward a • panel. The overall mechanical architecture for new vision of a lower-mass higher- 7 • M. Hennan 12th AIAAlUSU Conference on Small Satellites • SSC98-IV-l •

performance telecom subsystem. The work details which may be of importance for future initiated back then was the stepping stone for users. • the technologies which are going to be validated on DS 1. The Lockheed Martin Ka­ band 1.5W output power module was a The basic capabilities which the SDST • major A TI development and was used to help provides are receiving an uplink command develop the DSI 2.SW Ka-band SSPA signal, downconverting and then • (which includes the driver stages, power demodulating it. It takes spacecraft supply, as well as output power module) . engineering and science data, and modulates and upconverts the information for • transmission. The SDST is also important for New Millennium pursued the Pluto ATI radiometric purposes (Doppler and ranging). • HGA development; however, due to limited The unit can be run in coherent as well as funding a flight spare MPF unit was selected. non-coherent modes, and provide coherent • The Mars '98 program has continued this (to the carrier) Differential One-Way Ranging development with Boeing and wil1 be flying (DOR) tones. In fact, the SDST allows for this technology. Another major push with regenerative ranging (via an extemal interface • the Pluto ATI was the development of a from the receiver) if desired. digital receiver. The schedule and funding • was limited but it did result in the competition for a new full transponder The downlink from the SDST can be X-band development (including a digital receiver). As and/or Ka -band. The clock and data streams • will be described below, the Small Deep for these frequencies can be independent. For Space Transponder by Motorola is a major DS 1 we hard wired together the clock and • technology validation for the New data for simplicity since we would receive a Millennium Program. single clock and data stream from the • spacecraft computer. Small Deep Space Transponder • (SDST) To summarize some of the key features of the SDST: • The Small Deep Space Transponder is the • X -band receiver! downconverter heart of the telecommunication subsystem capable of carrier threshold of at least • and many of the new technologies to be -156 dBm. validated on DS 1 depend on this development • Command Detector Unit function • (one new technology enabling others). • Telemetry Modulation function Although the Pluto A TI effort started the • X- and Ka-band exciters • development of what would become the • Beacon Monitor Operation SDST, it became clear that the cost of • Coherent and non-coherent operation developing high technology components may • X- and Ka-band two-way ranging • be too expensive for future low-cost missions • Differential One-Way Ranging (DOR) to absorb; therefore, a consortium of JPL for both X- and Ka-bands • pre-projects from the Mars Exploration, • C&DH communication via 1553, Space and Earth Sciences, Technology And 1773, or RS422 Applications Program and the Data interface via RS422 • • Telecommunication and Mission Operations • External ports for temperature Directorates joined together to fund this sensors • ambitious technology task. In prior sections • External ports for analog telemetry of this article we have provided the reader signals • with the power states for the SDST, and the • Uplink and downlink radio science advantages of additional temperature capability sensors and analog-to-digital converters for • subsystem telemetry. Next we will focus on • 8 M. Herman 12th AIAA/uSU Conference on Small Satellites • • • • SSC98-IV-l The SDST mass is 2.95 kg and its volume is available as well (Manchester, NRZ-L etc. • less than 2681cc. Key technologies which all of which are already designed into the enabled this high functionality in a compact ASIC in the SDST). single unit include: GaAs Monolithic • Microwave Integrated Circuits (MMICs), advanced high frequency MultiChip Modules The desired subcarrier is also a user defined • (MCMs) using Low Temperature Cofired quantity. In fact, the beacon tones is just a Ceramic (LTCC) technology and Si designated subcarrier frequency (for DS 1 Application Specific Integrated Circuits they are 15 kHz, 20 kHz, 25kHz and 30 • (ASICs). This single unit replaces the kHz) with the carrier suppressed. These combination of a Deep Space Transponder tones can be on both X-band and Ka-band • (which is composed of an exciter and downlinks if desired. Also the number of receiver), Command Detector Unit (CDU), choices is not limited to four (the limit is in Telemetry Modulation Unit (TMU), Ka-band the numerically controlled oscillator- giving • over 900 frequencies to chose from). Exciter and provides performance never incorporated into a transponder ever before: The SDST power supply can be synched to • Beacon tones, external temperature an external source and is capable of operation monitors, external ports for analog-to-digital from +21V to +36V. • signal conversion, radio science I1Q output port, and external continuity monitors (e.g. sense waveguide switch position) just to list a Another facet of the SDST is the ability to • few. apply an external Ultra Stable Oscillator (USa) for uplink and/or downlink radio • science experiments. Future SDST users expressed the need to have design diversity to meet their unique • Another set of desirable features of the SDST requirements. In the design of the SDST we have tried to account for these needs. As an are the watchdog timers and various reset • example, some missions which are extremely mechanisms. The SDST has the ability to power constrained will desire an RS422 monitor when the last valid 1553 • interface versus a 1553 bus. Another mission communication was sent. If after a minute may use a fiber optic bus network and require there isn't any valid communication to the a 1773 interface (using a 1553 protocol). The unit, a reset would occur (if the RT timeout is • SDST was designed to accommodate all 3 enabled). In addition a reset can be achieved interface requirements. At the time of via the 1553 bus command or by an external • procurement, the desired interface can be opto-isolated reset switch. Although we do specified. not want to have to reset the transponder, like many advanced components (e.g. your • personal computer) sometimes it is The CDU functionality of the SDST necessary. During the DS 1 spacecraft • conforms to the CCSDS standard. While the integration and test phase we employed the telemetry modulation functionality is reset feature to quickly put the unit into a • extremely flexible due to the fact that the predefined known state. spacecraft provides the data clock and data. The data can be encoded in a Reed-Solomon • format before being sent to the SDST and The SDST development was initiated just then concatenated within the SDST with a before the formation of the DS 1 project. • choice of 7 112, 15 114, 15 112, or 15 1/6 Electing to use the SDST for DS 1 was a convolutional encoding. Also the downlink high-risk high-payoff baseline for NMP. The telemetry encoding can be bypassed to ease use of a Cassini-based transponder would • ground support testing or to take advantage have been low risk but it would not have of a higher performance encoding scheme if been cheaper and the ability to enable the • one becomes available in the near future. validation of other technologies (e.g., beacon • Various pulse code modulation schemes are monitor operation etc.) would not have been 9 • M. Herman 12th AIAAlUSU Conference on Small Satellites • SSC98-IV-l •

possible. Without the SDST the mass of the Unique features include built-in input/output subsystem would have been much greater. A isolators and engineering telemetry monitors • key to our success was by working closely (two gate currents, output drain voltage, and with the SDST leader and becoming part of internal unit temperature). Due to the short his (one-man) team. The DS 1 Telecom team development time for this unit, Lockheed • helped to test and evaluate the engineering Martin did not hermetically seal it. After model SDST as well as the flight unit. This deli very, some accelerated testing on other • true partnership was rewarding for our team. similar power devices has shown no major degradation after an initial bum-in. The flight • unit after 250 hours of operation (both in Ka -band Solid State Power Amplifier vacuum and atmosphere) has not shown any (KAPA) degradation of operation. We are very • cautious not to operate the unit too much in an open atmosphere and will not allow the • This unit will be the highest power solid state unit to go below dew point. Ka -band amplifier ever used for deep space communications. With the future • improvement of ground facilities as well with The key technology for KAPA is the use of spacecraft hardware, Ka-band holds a 0.25 micron GaAs Pseudomorphic High • potential four fold increase in data rate in Electron Mobility Transistors (PHEMT). The comparison with X-band. This is extremely efficiency could have been optimized further • important since a faster data rate translates to with the use of 0.15 micron devices; less required ground resources/mission however, time and financial resources operation support and this means reduced defined what our final product would be. • project cost. Another benefit of going to Ka­ band is the availability of greater bandwidth. • NASA as well as commercial programs Radio Frequency Electronics Test recognize this and are developing the StrateeY and Ground Support technology to move beyond microwave Equipment (GSE) • bands which are becoming crowded due to PCS and other emerging information • technology ventures. The RFE test strategy for DS 1, as with many aspects of the project, was a significant • departure from the traditional test programs at KAPA is a flyable engineering model JPL. At the system level, environmental developed by Lockheed Martin using their requirements were a particularly good • own funding. Its mass is 0.66 kg (this example of this. Because of the uniqueness includes input/output isolators, power of the RFE layout on multiple spacecraft • supply, telemetry circuitry and RF panels, there was no project level requirement electronics), with a RF output power of for vibration of the RFE subsystem as a 2.5W and a gain of 37 dB. whole before its delivery to the spacecraft. • Instead, the only requirement was that major subassemblies within the RFE had to be • The unit was qualified to DS 1 requirements either vibration tested or reviewed for which include: acceptability based on similar testing (as in • the case of MPF spare hardware). At the • Random Vibration- subsystem level we had to take the initiative 20 Hz 0.0322 G2/Hz and perform a detailed mechanical analysis • 50-500 Hz 0.2 G2/Hz concurrently with the layout of the RFE to 2000 Hz 0.0126 G2/Hz ensure that all mounting and bracketing • Overall 13 qms techniques could withstand the worst case • Thermal Vacuum cycling from -14°C vibration loads. When the spacecraft was to +40°C later vibration tested in all three axes, the • • Full EMC testing to MIL SPEC 461 • 10 M. Hennan 12th AIAAIUSU Conference on Small Satellites • • • • SSC98-IV-l telecom subsystem passed with no validate the required functionality of the RFE • mechanical or electrical anomalies. with the knowledge that in later phases of testing we could perform additional characterization and also track the consistency • A similar approach was taken for thermal of data through each phase. vacuum (TVAC) and EMC testing. As with • vibration, we performed both of these tests only on the major individual subassemblies. The need to track the performance of the RFE Those that were not tested were analyzed for through all phases of testing required us to • potential problems; this was especially develop an efficient yet inexpensive method important in the case of TVAC where of data acquisition. For the SDST being • multipaction becomes a significant concern developed at Motorola, a system built around with high power levels and the small the commercially available HPVEE program • geometries at X- and Ka-band. In many was conceived; it was installed on a standard cases, good upfront designs that factored in PC platform and included the ability to poll TVAC and EMC issues early on precluded (as fast as twice a second) either the primary • any problems. Later spacecraft testing in communication bus or RS-422 direct access both of these environments confirmed the path to retrieve engineering telemetry data and • robustness of the telecom subsystem. display it in a specially designed panel on the computer screen. However, Motorola did not incorporate any utility to monitor and control • Test planning for the DS 1 RFE was also a external instrumentation or to store data in challenging task in adapting to the new way memory for later extraction and review. • of business. It began in the early stages of Since we decided that this type of capability the project with a detailed test matrix. This was key to a successful test program, we • matrix listed all of the required tests against initiated the task to enhance Motorola's the various phases of testing throughout the telemetry system not only for the SDST but duration of the project from initial the entire RFE subsystem. When the • characterization of the RFE to spacecraft software was finally completed, we could launch. For each phase, a subset of the monitor and control all of the RFE Ground • required tests would be checked off to Support Equipment, access a variety of validate performance in that configuration. display panels to view important subsets of As the project progressed, mechanical and data, store data on a daily basis for later • thermal issues set the pace for defining the review, and plot data trends using specially telecom architecture and soon the allotted time designed macros. • for testing became shorter and shorter. This forced us to reconsider our test plan several • times with multiple iterations of the test The ability to store and extract data for later matrix until we were left with only two review and analysis proved to be an weeks of test time (before subsystem delivery indispensable feature of our test strategy. • to the spacecraft). With few options left we Because the data acquisition operated at twice decided to eliminate testing over temperature, a second, we possessed great ability in • scrub the functional and Deep Space Network correlating the condition of the RFE compatibility testing to the bare minimum, subsystem at a specific time with any kind of and work two ten hour shifts a day, seven external event; this was especially valuable • days a week. We accepted the risk of not when tests were being conducted at the temperature testing because all of the major spacecraft level. We had a tremendous • subassemblies had been previously advantage over the spacecraft telemetry temperature tested and the RFE subsystem system in that our data acquisition ran up to • would eventually undergo temperature twenty times faster. This proved invaluable cycling when the spacecraft went into Solar while performing spacecraft level vibration Thermal Vacuum (STV) testing. In and EMC tests where it's vital to have almost • streamlining the DSN compatibility testing, instantaneous visibility into the health of your • we performed the minimum necessary tests to subsystem. But it was also valuable in 11 • M. Hennan 12th AIAA/USU Conference on Small Satellites • SSC98-IV-l • correlating changes in the RFE with signal and reception of a downlink signal commands and sequences from the spacecraft from the spacecraft. It also has ability to • computer, with testing being performed on downconvert the downlink signal (at X-band) other subsystems, or with work in general to baseband and demodulate, bit sync., and being done on or near the spacecraft In finally decode it. The TRC also contains • countless number of cases our data two data acquisition units which monitor the acquisition confirmed that problems were not outputs of the commercial instrumentation. • due to the RFE subsystem. • The SDST support equipment consists of two When we considered how to implement our racks. One of them houses the computer that Ground Support Equipment, cost was as commands the SDST into different modes • always a major factor and our requirement and also controls the data acquisition for the was to be able to support component, entire RFE. The racks also contain the uplink • subsystem and system level testing. At the signal generator, variable attentuator to set time Mars Pathfinder was at Cape Canaveral uplink power level, ranging tone generator, preparing for launch and had a couple of spectrum . analyzer, oscilloscope, and a • racks of telecom GSE that we could retrofit frequency counter. for DS 1. We decided to purchase these racks • with a promise of immediate delivery after launch so we would have time to make the The final rack is the Command Detector Unit necessary modifications and perform some • (CDU) rack. This is another computer functional tests. By doing this we saved the controlled rack used to provide a number of project considerable cost. The other GSE uplink command related functions. One of • that would be required were two racks which these functions is to provide uplink subcarrier Motorola used to operate the SDST and one modulation to the SDST with either pseudo­ • more half-rack , the Remote Calibration random (PN) data or data from an external Module (RCM), to multiplex signals from the source. The PN data are used to characterize various telecom antennas and provide Ka­ the uplink bit error rate of the SDST receiver • band downlink monitoring functions. The while the external data input is used to cost of the developing the SDST GSE racks perform end to end (uplink to downlink) • were covered as part of the overall contract tests. The CDU rack can also perform with Motorola. So the only rack that we had probability of acquisition and de-acquisition to develop from scratch was the RCM. tests which determine the SDST's threshold • Though the cost of this rack was not for acquiring command data lock or unlock in insignificant due to the extensive Ka-band the required number of bit times. • functionality, as a whole we reduced our overall GSE cost enormously by making use • of currently existing hardware. . Acknowledaements • A block diagram of the DS 1 telecom GSE is We wish to thank Leslie Livesay (DS1 shown in Figure 6. As described before, the Spacecraft Manager) and Dave Lehman (DS 1 • RCM is designed to be positioned near the Project Manager) for their trust and support. spacecraft to select the proper antenna path Robert Hughes and Gerry Gaughen provided for the X-band signals which are then carried guidance and expert technical input which • by one RF cable to and from the test facility helped enable an aggressive plan to be up to 600 ft. away. The RCM also has successfully implemented- we owe them a • equipment to monitor the Ka-band downlink great deal of thanks. Steve Burkhart, Keith and downconvert it to X-band for lower loss Pfleiger, and Tom Cooper, whom are no • transmission and ease of processing in other longer are with JPL, provided us an racks. The RCM is controlled by the excellent start for DS 1 via their work on the TransmitJReceive Controller (TRC) which MPF telecommunication subsystem· • allows simultaneous output of an uplink development. • 12 M. Herman 12tb AIAA/USU Conference on Small Satellites • • • • SSC98-IV-l We wish to thank both the Lockheed Martin 3. World Wide Web address: • Ka-band SSPA and Motorola SDST teams: http://nmp.jpl.nasa.gov/dsl/ Since Web sites can be dynamic, • Lockheed Martin Ka -band SSPA team: another approach is to go to the M. Karnacewicz, W. Taft, T. Renna, S. main JPL site and then look at new • Conway, M. Hirokawa, S. Valenti and and/or on going missions at: • L. Newman. http://www.jpl.nasa.gov Motorola SDST team: 4. Sovey, J.S., L. M. Carney, and S.C. • S. Jackson, C Knuckolls, K. Seimsen, Knowles," Electromagnetic Emission L. Carson, D. Altum, D. Burgess, Experiences Using Electric R. Hurkes, D. Andersen, K. Newman, Propulsion Systems," AIAA Journal • of Propulsion and Power, Vol. 5, and many more .... No.5, September-October 1989, pp • 534-547. Contributions by Paul Kahn for his HPVEE • programming, Steve Petree for GSE support, Joe Vachionne for his antenna 5. Herman, M., S. Burkhart, R. Crist, C. contributions, and Andrew Makovsky for Hornbuckle, W. Hoffmann, D. • system analysis were greatly appreciated. Smith, J. Vacchione, R. Hughes, Plus many more who have helped us achieve K. Kellogg, A. Kermode, D. • our goal. Rascoe," Microtechnology in Telecommunication for Spacecraft Cost and Mass Reduction", IAA • International Conference on Low­ The research described in this paper was carried out by the Jet Propulsion Laboratory, Cost Planetary Misssions, Laurel • California Institute of Technology, under a MD,1994. contract with the National Aeronautics and • Space Administration.

• References • L Rayman M.D., P.A. Chadbourne, J.S. Culwell, and S.N.Williams, • "Mission Design For Deep Space 1: A Low-Thrust • Technology Validation Mission," Third IAA International • Conference on Low-Cost Planetary Missions, Pasadena, • CA, April 1998. • 2. Rayman, M. D. and D. H. Lehman, "Deep Space One: NASA's First Deep -Space Technology Validation Mission," • 48th International Astronautical Congress, Turin, Italy, • October, 1997. • 13 • M. Herman 12th AIAAlUSU Conference on Small Satellites • • SSC98-IV-l • • • Hardwired POR • TELEMETRY • Telemetry Data, Telemetry Clock ------l. fromffiM • Command Data, ....f---­ LGA+Z Command Lock, Command Clock • toIEM LGA·Z Control and • Health Status viaIEM • IBM is the iniegrared electronics To Det.Mod. moduie, which houses the sIc computer and other key electronics • • TELEMETRY c=I Hetitage DeSign New Design/Procurement • Figure 1. DS 1 Telecom Subsystem block diagram. • • • • • • • • • Figure 2. This is the interior view ofthe +X panel (length is -1m). Attached (to the left side) • is the X-band solid state power amplifier. This unit was developed at JPL and has an output power of 12.5W (RF) and a gain of 38 dB. {The left side of the panel • corresponds to the +Z axis of the spacecraft.} • 14 M. Hennan 12th AIAAIUSU Conference on Small Satellites • • • • SSC98-IV-l • • • • • • • • • Figure 3. This is the exterior view of the +X, +Y panel. Attached are the waveguide transfer • switches (WTSs), diplexer and waveguide connections to the antenna. {The right • side of the panel corresponds to the +Z axis of the spacecraft.} • • • • • • • • • • Figure 4. This is the interior view of the +X,+ Y panel. A major portion of the active telecom subsystem electronics resides here. Key components include (from right to left) • SDST, Detector Amplifier Module, and KAPA {The left side of the panel • corresponds to the +Z axis of the spacecraft.} 15 • M. Hennan 12tb AIAAlUSU Conference on Small Satellites • SSC98-IV-l • • • A Z • • • A • X • -2m • • • • • • • • Figure 5. View of the DSI spacecraft featuring the telecom panels (+X to the left and +X,+Y to the right of center). Note that the HGA , 2 LGA and the Ka-band antenna are • visible on top of the spacecraft. The -Z LGA was not attached to the service boom at the time of this photo. • • • • • • • 16 M. Herman 12th AIAAlUSU Conference on Small Satellites • • • • SSC98-IV-l • TRC SDST CDU Digital Multimeters RS-422 Downlink Frequency Frequency Counter Interface Drawer Counter • Spectrum Analyzer Subcarrier Signal Power Meters Generator • Spectrum Analyzer Oscilloscope Oscilloscope • RF Interface Panel Oscilloscope

RCM Control Uplink Synthesized Digital Multimeter • Signal Generator RCM Computer Monitor • S-band Receiver Bit Error Rate Uplink Variable Attenuator Electronics • Data Acquisition Antenna Select ion Computer Ranging Tone Subcarrier Computer • Generator Demodulator Ka-band Spectrum • Analyzer Bit Sync. & 10 MHz Reference Uplink Frequency Viterbi Decoder Counter Power Supplies Sweep Oscillator Distribution • Amplifier • Power Supplies Power Supplies • • Figure 6. DS 1 Telecom support equipment block diagram. • Biollraphies Sam Valas: B.S. University of California, Los Angeles (,86), M.S. from the University • Martin Herman: B.S. Drexel University of Southern California ('91) both in Electrical ('82), M.S. and PhD. from University of Engineering. Mr. Valas was with Hughes • Michigan ('83, '87) all in Electrical Microwave Products Division ('87-'91) Engineering. Dr. Herman was with the where he worked on high power millimeter­ • Hughes Microwave Products Division ('87- wave amplifiers, transceivers, and advanced '92) where he worked on advanced MMIC technology. From 1991-92 he was at automotive radar and GaAs MMIC Hughes Space & Communications Group • technology. At the California Institute of where he designed microwave amplifiers and Technology's Jet Propulsion Laboratory other components for commercial satellites. • ('92-present) he has been involved in Since 1992 he has been with the Jet advanced packaging for high frequency Propulsion Laboratory, where he is currently components, Telecom lead for the Pluto a Senior Engineer in the Spacecraft • Flyby study team and for the past 3 years he Telecommunications Equipment Section and has lead the Telecorrimunication the Radio Frequency Subsystem Cognizant • Development for Deep Space 1. Currently he Engineer on New Millennium Deep Space is co-lead for the Advanced Micromachined One. Prior to his work on DS 1, he was RF Front End Technology Development at involved in design and development of a Ka - . • JPL. • band transmitter for the Cassini mission to 17 • M. Herman 12th AIAAlUSU Conference on Small Satellites • • SSC98-IV-l •

Saturn and led the development of an X -band Sam Zingales worked at JPL from 1966 to transmitter for Mars Pathfinder. Besides his 1978. He was involved in Spacecraft • duties on DS 1 he is also co-leading the telecommunications equipment for the Advanced Micromachined RF Front End Mariner-Mars 1969, Viking and Voyager • Technology Development. programs. Upon returning in 1991, he managed the design and implementation of two 34 meter antennas for tracking low Earth • William A. Hatch: B.S. in Electrical and orbit spacecraft. He then led a feasibility Electronic Engineering from California State study for an autonomous airplane powered • Polytechnic University in Pomona, CA in by a microwave beam from the ground. 1982. From 1979 until 1992, he worked for Flying in a circular pattern 70,000 feet above the General Dynamics Air Defense Systems the station, the aircraft platform was designed • Division, where he was involved in the to provided various types of telecomm. component, subsystem and system level services including cellular phone access and • Microwave design for use in Missile and direct TV. In 1995 he was assigned Ground based RADAR systems. In addition, responsibility for the development of the • as part of his research activities, he developed Small Deep Space Transponder being single-chip GaAs subassemblies. In 1992 he developed by Motorola. He spent much of went to work at Technology Applications, his time in Phoenix, AZ where he was • where he designed a LTCC Multi-Chip intimately involved in the detailed design and module Transceiver. In 1994 he came to the programmatic aspects of the effort. • work for the Spacecraft Telecommunications Equipment Section at the Jet Propulsion Laboratory, where he has designed Mr. Zingales holds a BSE from Cleveland • subassemblies for the Primary and Auxiliary State University. transmitters for the Mars Pathfinder • Spacecraft. He is presently the lead engineer responsible for the Integration and Test of the Rocco Scaramastra is a graduate of West • Telecommunications Subsystem for the New Coast University, BSEE. He is also a Millennium Deep Space 1 Spacecraft. Mr. registered professional electrical engineer in Hatch is a member of Eta Kappa Nu and the state of California. He has an extensive • IEEE Microwave Theory and Techniques background in the design and development of Society and is a Registered Professional telecommunications and RADAR equipment • Engineer. for aerospace and commercial applications. • Chien-Chung Chen: is a communications His recent experiences at JPL include systems engineer at JPL's Communications leadership for the development of the Mars • Systems Research Section. He received his Pathfinder Radio Frequency Subsystem, lead Ph.D. in electrical engineering from the engineer for the DSI Ka-band electronics and • University of illinois at Urbana-Champaign detector system, and he was a key engineer in 1987, and joined JPL shortly after. At for the Cassini Radio Subsystem JPL, he has worked on various . development. • communications systems from RF to optical wavelengths. He is currently serving as the • telecommunications systems engineer on the Luis R. Amaro. Received his BSEE from New Millennium DS 1 project, and for the California State Polytechnic University, X2000 Optical Communications Subsystem. Pomona in 1984. He received his MS degree • Dr. Chen has authored or co-authored over from California State University, Northridge 25 technical publications in laser comm. and in May of 1998. From 1984 to 1987 he was • related fields. a Member of the Technical Staff for the Metrology Dept. at Rockwell Inti. Autonetics • Marine Systems Division. From 1987 to 1991 he was a Member of the Technical Staff • 18 M. Herman 12th AIAA/USU Conference on Small Satellites • • • • SSC98 .. IV .. 1 with the Microwave Components Dept. at • TRW Inc. He has designed various waveguide components such as attenuators, filters and multip1exers for use on flight • programs. He received a best paper award at the 1987 Measurement Science Conference • for his work on power sensor calibrations. From 1986 to 1989 he served in the IEEE as Membership Chairman and Chairman of the • Instrumentation and Measurement and Industrial Electronics Societies, Los Angeles • Chapter. He is a Fellow of the Institute for the Advancement of Engineering. From • 1991 to the present he has been a Member of the Engineering Staff with the Spacecraft Antennas Group, Jet Propulsion Laboratory. • While at JPL he has developed a novel waveguide hinge for use on the NSCAT • project and recently developed a dual mode dual channel waveguide rotary joint for the • SeaWinds project. • Mr. Amaro's interests include EM simulation and numerical analysis of antennas and • microwave components. • Marc D. Rayman received his A.B. in physics from Princeton University and his • M.S. and Ph.D. in physics from the University of Colorado in Boulder. Since joining the Jet Propulsion Laboratory in 1986 • he has worked on a wide variety of projects including deep-space optical comm., optical • interferometry space missions, Mars sample return, and the Space Infrared Telescope • Facility. In 1994 he he1ped formulate the New Millennium program and when Deep Space 1 began in 1995 he became the Chief • Mission Engineer. He will be the Deputy Mission Manager when the project enters • operati ons. • • • • • 19 • M. Herman 12th AIANUSU Conference on Small Satellites •