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I. IntroductionTHE Demonstration and Science Experiments (DSX) flight program, is AFRL’s fourth space science technologyexperiment (SSTE4). It was originally conceived by AFRL researchers in 2003 to conduct physics basedexperiments, and was selected as an AFRL mission in 2004. With primary funding from AFRL, DSX enjoysadditional support from DARPA and NASA, and is comprised of elements provided by AFRL, NASA, academia,and many contractors.The DSX spaceflight experiment comprises three major research areas, which together will pave the way for newDoD capabilities in space surveillance, small satellites with significant operational capabilities, and protection ofspace assets from natural and enhanced radiation environments. This mission will advance the warfighter’scapabilities in communication, surveillance, and navigation. The DSX experiments include research in three majorexperiment categories.The physics of Very Low Frequency (VLF, 3-50 kHz) electromagnetic wave injection from space and groundbased transmitters, propagation, and wave-particle interactions in the magnetosphere comprises the first category,Wave Particle Interaction Experiments (WPIx). Equipment on DSX will transmit and receive VLF waves andquantify their effect on the trapped electron populations in the magnetosphere. Detailed measurement and mappingof high and low energy particle and plasma distributions; radiation dose rates, local magnetic fields and pitch angledistributions in the poorly characterized MEO environment and slot region make up the DSX Space WeatherExperiments (SWx). The third major experiment category, Space Environmental Effects (SFx), involvescharacterization of the effects of space weather and the space environment on materials and electronics. SFxconsists of NASA’s Space Environment Testbed, as well as AFRL developed photometers and radiometers.Coupling these experiments into a single platform provides a lower-cost opportunity for the AFRL due to theircommon requirements and goals. All three experiments need a 3-axis stabilized spacecraft bus, a suite of radiationsensors, and extended duration in a MEO orbit.II. Flight Experiment ConceptThe functional baseline configuration for the DSX flight experiment is shown in Figure 1. The core is theEvolved Expanded Launch Vehicle (EELV) Secondary Payload Adapter (ESPA) Ring [1,2], which is used tomaximize launch opportunities on both Space Test Program (STP) and operational DoD launchers. ESPA wasdeveloped and designed by CSA Engineering. Every EELV launch is a potential ride for an ESPA ring, and thusalso for DSX. The ESPA ring comprises the primary structure for DSX, and is upgraded to provide host spacecraftfunctions (avionics, TT&C, ADCS, C&DH, and power management and distribution) by the addition of componentspackaged on an avionics module (AM). The DSX payloads (including deployable booms) are mounted on anidentical structure, the payload module (PM), attached to the ESPA ring opposite the avionics module. The AM andPM together comprise the DSX Host Spacecraft Bus (HSB). The entire assembly is designed to be stowed within a4-m diameter EELV fairing.(4)Figure 1: DSX baseline deployed configuration.2American Institute of Aeronautics and Astronautics

The design intent for the ESPA ring was for it to remain attached to the launch vehicle upper stage to facilitatedeployment of up to six microsatellites. However, unlike the traditional ESPA approach, the DSX avionics andpayload modules do not separate; they remain attached. After the primary satellite is deployed, the DSX ESPAseparates from the EELV 2nd stage booster to become a free-flyer spacecraft. The ESPA ring serves as the DSXspacecraft’s primary structure, with an upper separation interface to the primary payload adapter and a lowerseparation interface to the EELV upper stage adapter. Figure 2 shows DSX and identifies its major components in astowed configuration. The experiment is baselined for flight in the radiation belts with a nominal orbit of 6,000 kmx 12,000 km elliptical, mid-inclination, and 1-year of mission operations.PRIMARY SEPARATION I/FEELV UPPER STAGE SEP I/FAvionics Module (AM)Payload Module (PM)(AM)YXZResidualSeparationHardwareESPAFigure 3: DSX stackedconfiguration fordedicated launchersFrom the earliest DSX conceptual planning, it was understoodthat the key to a successful program execution was maximizingDSX’s compatibility with numerous launch vehicles. While theESPA-based approach makes every EELV a potential ride, not allco-manifest opportunities possess sufficient excess launch vehicleperformance needed to inject DSX into its MEO orbit. Therefore,flexibility in the design will be maintained for as long as possibleto allow repackaging DSX for flight on dedicated launchers.Inherent in the modular approach shown in Figure 2 is the abilityto separate the two modules from the ESPA ring, and directlystack them in a vertical assembly for a dedicated launcher. Thestacked HSB configuration is shown in Figure 3, and fairingencapsulated views of DSX are shown for EELV, Minotaur IV/V,and Falcon 5 in Figure 4. This modular architecture not onlymakes DSX reconfigurable for various launch vehicles, but it alsogreatly simplifies payload integration.Primary Payload(placeholder)Figure 2: DSX baseline stowed configurationFigure 4: DSX EncapsulatedFrom left to right - Atlas V EELV (baselineESPAsat), Minotaur IV/V, and Falcon 5III. DSX Science ExperimentsA. Wave-Particle Interaction Experiment (WPIx)The Earth's radiation belts exhibit considerable dynamic behavior ranging from the creation and destruction ofwhole new belts in the outer zone ( 6,000 km) occurring on time scales of minutes to days and the slower diffusionof the innermost regions ( 6,000 km) occurring on timescales of months to years. A major cause of the energeticelectron belt dynamics are wave-particle interactions driven by electromagnetic waves in the VLF range and below.Properly directed, this scattering will increase the electron velocity parallel to the local magnetic field by altering the"pitch-angle" (i.e. the angle between the particle's parallel and perpendicular velocity) and lower the altitude atwhich the electron magnetically reflects along the field line, an effect due to the conservation of magnetic moment(Figure 5). If the magnetic reflection is lowered to altitudes within the upper atmosphere, collisions with theplentiful neutral particles will lead to a depopulation of the radiation belt along that magnetic flux tube.3American Institute of Aeronautics and Astronautics

Figure 5: VLF wave particle interaction processes.Natural sources of VLF in the magnetosphere include magnetospheric hiss generated by space weather processesand lightning-induced whistler waves propagating along field-aligned ducts through the magnetosphere. To matchcurrent models of radiation belt dynamics to the observed behavior of electrons, it is necessary to postulate that manmade VLF leaking into the magnetosphere from ground-based submarine communications systems is also asignificant source. Direct measurements of VLF power have been sparse because scientific satellites with therequired capability (e.g. the USAF CRRES [3] and NASA IMAGE [4] missions) have been in highly elliptic orbitswith apogee 6 Earth radii and consequently spend most of their time outside the inner belt and slot region. Placinga VLF receiver on the DSX satellite will allow for a quantitative determination of the current levels of natural andman-made VLF waves in the inner magnetosphere. Models of ground-based VLF injection and globalmagnetospheric propagation will also be validated in a controlled manner. The VLF transmitter on DSX will providethe capability to conduct active experiments to quantify space-based VLF wave-injection efficiency and determinethe details of the wave-particle interactions. An electron loss cone detector on DSX will allow direct correlation ofchanges in energetic particle distributions with injected wave power. Accurate models of the VLF injection,propagation, and wave particle interaction processes will be developed and validated with DSX data. Such modelsare critical components of global radiation belt nowcast and forecast models required for space situational awarenessand mission planning. The WPIx payloads consist of four major components, further described below.1. Wave-induced Precipitation of Electron Radiation (WIPER) Broad-band Receiver (BBR)The WIPER broadband receiver (BBR) is a passive receiverdesigned to measure the natural and man-made ELF, VLF andLF signals in the inner magnetosphere in the range 3-50 kHz.The BBR (shown in Figure 6) is electrically connected to allfour deployed booms previously shown in Figure 1, with twochannels of the BBR measuring electric fields via the y and zaxis VLF antennas, and three more channels measuring Figure 6: WIPER BBRFigure 7: TASCmagnetic field about threes axes of the Triaxial Search CoilAntennas (TASC). TASC (shown in Figure 7) is mounted on the tip of the Z boom and the BBR is containedwithin the payload module. BBR measurements are also coordinated with high power transmissions resulting fromthe WIPER transmitter, transmissions from ground based facilities, as well as naturally occurring sources ofmagnetospheric VLF (i.e., lightening). The WIPER BBR is being developed by Stanford University.2. WIPER Transmitter and Narrow-band Receiver ElectronicsThe WIPER transmitter electronics portion consists of the transmitter controllerunit (TCU) with an integral narrow-band receiver (NBR) as well as two transmitteramplifier tuner units (TATU). Each TATU (shown in Figure 8) electrically connectsto a pole of a center-fed dipole antenna supported by the Y booms shown in Figure1. The transmitter electronics include dynamic tuning circuits, embedded within theTCU. The VLF transmission system will broadcast up to the kilowatt level in therange 1-50 kHz and at much smaller power levels from 50 kHz - 750 kHz by charging4American Institute of Aeronautics and AstronauticsFigure 8: WIPER TATU

conductive antenna elements that run along the Y-axis booms to levels up to 10 kV. The TCU/NBR and bothTATUs are housed within the payload module. The WIPER transmitter electronics are being developed by theUniversity of Massachusetts-Lowell, benefiting from significant heritage resulting from work on a similar payloadflown on NASA’s IMAGE satellite.3. Loss Cone Imager (LCI)The Loss Cone Imager (LCI) is an electron loss-cone particle detector that will provide a measurement of threedimensional energetic particle distributions with emphasis on the measurement of the fluxes of energetic electronsalong the direction parallel and anti-parallel to the local geomagnetic field vector. The LCI consists of two scanningheads, with 180 motorized articulation capability, and each projected a 180 field of view fan, with each scanhead position, so that together the full 4π unit sphere may be imaged. LCI depends on data from the DC VectorMagnetometer (VMAG) for its motor pointing control loop, and for later ground data correlation.A separate Solid State Detector (SSD) telescope called the High SensitivityTelescope (HST) will be mounted on the DSX in order to obtain fluxes of energeticelectrons along the geomagnetic field vector direction. This telescope is designed tohave a geometric factor of 0.1 cm2 steradian with sufficient shielding to permit thedetection of 100 particles/cm2-sec-steradian in the loss cone. The LCI is underdevelopment by Boston University, and also benefits from heritage obtained in thedevelopment and flight of similar instruments such as the Imaging ElectronSpectrometer (IES) on the ESA-NASA Cluster mission. The LCI scanning heads andHST are shown in Figure 9.Figure 9: Loss Cone Imager4. DC Vector Magnetometer (VMAG)A DC fluxgate vector magnetometer (VMAG) will be used todetermine the direction of the magnetic field to better than onedegree at all points in the orbit. This performance will allow themapping of locally measured particle distributions to globaldistributions. VMAG sensors will measure the DC B-field over a100-10000 nT range with 0.1 nT accuracy at 20 Hz. VMAG dataFigure 10: VMAGis also used in real-time by the LCI, as previously discussed.VMAG will operate continuously to provide DC magnetic field and ULF wave environment data required by theSWx MEO space particle modeling experiment, thus it supports both WPIx and SWx experiment objectives.The VMAG electronics will be mounted within the payload module and the fluxgate sensor will be deployed onthe tip of the –Z boom, opposite TASC. VMAG is under development for DSX by UCLA, and is similar to otherfluxgate magnetometers (shown in Figure 10) developed by UCLA and flown on several missions.B. Space Weather Experiments (SWx)With an orbit between 6,000-12,000 km, DSX will explore a large swath of the inner magnetosphere, inparticular the outer region of the inner proton belt the “slot region”, and inner regions of the outer electron belt. Thisdomain has remained largely unexplored due to the understandable tendency of military, commercial, and scientificsatellite systems to be located outside the intense regions of radiation, most often at LEO or GEO. However,increasing demands for uninterrupted coverage of the Earth from space are driving planners to consider puttingsatellite constellations in orbits spending significant time in the inner magnetosphere. Current standard space particlemodels (e.g. AE-8 and AP-8) can be off by as much as 50 times or more in estimating the time taken to reachhazardous dose levels in the MEO regime. It is imperative that measurements be made of the plasma and energeticparticles so that adequate climatological, situational awareness, and forecast models can be developed to enable thesuccessful design and operations of systems in these new and desirable orbit regimes. Accurate environmentdetermination is important for DSX so that quantitative correlation with the performance of the spacecraft and itspayloads over the course of the mission may be performed.Deficiencies of current standard radiation belt models in the inner magnetosphere include the lack of (a)spectrally resolved, uncontaminated measurements of high energy protons (10-400 MeV) and electrons (1-30 MeV)and (b) accurate mid to low energy ( 1000 keV) measurements of the energetic particle and plasma environment.Not surprisingly, most of the space weather data to-date has been accumulated in the LEO and GEO regimes, asillustrated in Figure 11 with data from dosimeters aboard the TSX5 and DSP satellites in LEO and GEO,respectively. The Space Weather instruments onboard and the unique orbit of DSX will address these deficiencies.5American Institute of Aeronautics and Astronautics

Included will be electron and proton detectors measuring both the spectral content and angle of arrival of bothspecies over broad energy ranges. An on-board magnetometer (VMAG, see section A.4) will allow for thetransformation of angle-of-arrival measurements into estimates of the flux distribution with respect to the localpitch-angle (i.e. the angle between the particle velocity and magnetic field). Local pitch-angle distributions can thenbe used to estimate global particle distributions by mapping techniques using the well-known equations of motionand magnetic field models tagged to the local measurements. The nominal space weather sensor suite capabilitiesare described in detail in the following sub-sections.Figure 11: Space weather data from TSX and DSP shows lack of data in MEO.Figure 12: Space weather instruments CEASE, LEESA, HIPS, LIPS, and HEPS.1. Compact Environment Anomaly Sensor (CEASE)Composed of two dosimeters, two particle telescopes and a Single Event Effect detector, CEASE has thecapability to monitor a broad range of space hazards from surface damage and charging (keV electrons) to SingleEvent Effects resulting from 100 MeV cosmic rays and solar protons [5,6]. The angular field-of-view for CEASEis relatively large and will not allow for pitch angle resolved measurements. CEASE will be mounted on an exteriorpanel of the payload module. One change for DSX is that CEASE will capture and downlink the full dose spectrafrom each dosimeter, whereas prior versions only captured six reduced data points (two for low LET data and fourfor high LET data). The CEASE instrument was developed by Amptek, Inc. and has already flown on severalspacecraft. CEASE and the other SWx instruments are shown in Figure 12.2. Low Energy Electrostatic Analyzer (LEESA)LEESA will measure energy fluxes and energy spectra for low energy electrons and protons (100 eV to 50 keV).These low energy particles are responsible for surface electric charging and damage to thin films such as thin-filmphotovoltaics, conventional solar cell cover glasses, and coatings. LEESA will be mounted on an exterior panel ofthe payload module. It is under development for DSX by Amptek.3. High Energy Imaging Particle Spectrometer (HIPS)HIPS will measure electrons with energies between 1 and 10 MeV and protons with energies between 30 and300 MeV. These high energy particles are responsible for microelectronics damage, displacement and total dosedamage, SEEs, and deep dielectric charging. HIPS will be mounted on an exterior panel of the payload module. Itis under development for DSX by Physical Sciences, Inc (PSI).4. Low Energy Imaging Particle Spectrometer (LIPS)LIPS is designed to measure the ring current particles that are important in the energy flow processes in themagnetosphere. This particle population plays an important role in spacecraft charging and surface damage due tosputtering. The instrument, built by Physical Sciences Inc, uses specially designed combinations of scintillatormaterials to detect electrons and protons with energies between 30 keV and 2 MeV. Eight 10 degree by 8 degreefield-of-view windows will provide pitch angle resolution. LIPS will be mounted on an exterior panel of thepayload module.6American Institute of Aeronautics and Astronautics

5. High Energy Particle Spectrometer (HEPS)HEPS is a two box system that will measure protons with energies between 15 and 440 MeV and electrons withenergies between 1.5 and 20 MeV. These high energy particles are responsible for microelectronics damage,displacement and total dose damage, SEEs, and deep dielectric charging. The HEPS sensor head will be mounted onthe AM battery enable bracket and the electronics unit will mount below it to an exterior panel of the AM. It wasdeveloped for DSX by Amptek, Inc. and is currently being completed by Assurance Technology Corporation(ATC).C. Space Environmental Effects1. Space Environment Testbed (SET)The SET carrier (shown in Figure 13), under development by NASA GSFC provides standard mechanical,electrical, & thermal interfaces for a collection of small flight investigations. The general areas of investigation ofthe instruments carried on SET fall in the following categories:1. Characterization of the space environment;2. Performance improvement for microelectronics used inspace;3. Accommodation and/or Mitigation of SpaceEnvironment Effects for detectors & sensors4. Accommodation and/or mitigation ofcharging/discharging effects on spacecraft & spacecraftcomponents;5. Definition of mechanisms for materials’ degradationand performance characterization of materials designed Figure 13: SET-1 (the red plate is non-flight).for shielding from ionizing radiation.The first SET mission, SET-1 consists of several specific investigations which are summarized on the followingpage.7American Institute of Aeronautics and Astronautics

Credance. Cosmic Ray Environment Dosimetery and Charging (Credance) measures the radiation environment(total dose), and the charging environment for normal background and during space weather events. Credancespecific measurements include:1. Proton flux 40 MeV per unit solid angle2. Charge deposition in large silicon diodes arranged in telescopes. Pulse height analysis is used to obtain ion linearenergy transfer (LET) spectra of heavy ions in the 100 MeV cm2/g to 25000 MeV cm2/g range3. Threshold voltage shift as a function of time to measure total ionizing dose in silicon at two different shieldingdepths4. Charging current at three different shielding depths which provides energetic electron flux measurements at threeenergiesCredance measurements overlap in part with the CEASE space weather instrument, and will provide for mutualvalidation and data correlation of both instruments. Credance has been developed by QinetiQ/UK.DIME. Dosimetry Intercomparison and Miniaturization (DIME) will measure space radiation environmentswhich are detrimental to space system reliability using novel dosimetry techniques. DIME occupies two 3U cardslots on the SET-1 carrier, and is being developed by Clemson University. Specific DIME measurements include:1. Radiation-Sensing Field-Effect Transistor (RADFET) - threshold voltage shift as a function of time; converted tototal ionizing dose.2. Erasable Programmable Read-only Memory (EPROM) - threshold voltage shift as a function of time; convertedto total ionizing dose. Also number of single event upsets as a function of time to measure rates as a function ofradiation level.3. Static Random Access Memory (SRAM) - hold devices at different voltages and measure single event upsets.Change in voltage at which an error occurs is converted to dose. Also number of single event upsets as afunction of time to measure rates as a function of radiation level.4. Linear Energy Transfer (LET) Spectrometer - pulse height spectra as function of time; converted to LET tomeasure ions including protons.5. Optically Stimulated Luminescent (OSL) Films - visible emission spectrum as a function of time to measuremicro-dose.ELDRS. Focused on characterization of Proton Effects and Enhanced Low Dose Rate Sensitivity (ELDRS) inBipolar Junction Transistors (BJT), ELDRS measures space radiation induced total ionizing dose and displacementdamage on linear devices sensitive to enhanced space low dose rate effects. ELDRS will measure base and collectorc

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