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NASA NTRS2000

The Strategy for the Second Phase of Aerobraking Mars Global Surveyor

On February 19, 1999, the Mars Global Surveyor (MGS) spacecraft was able to propulsively establish its mapping orbit. This event followed the completion of the second phase of aerobraking for the MGS spacecraft on February 4, 1999. For the first time, a spacecraft at Mars had successfully employed aerobraking methods in order to reach its desired pre-launch mapping orbit. This was accomplished despite a damaged spacecraft solar array. The MGS spacecraft was launched on November 7, 1996, and after a ten month interplanetary transit was inserted into a highly elliptical capture orbit at Mars on September 12, 1997. Unlike other interplanetary missions, the MGS spacecraft was launched with a planned mission delta-V ((Delta)V) deficit of nearly 1250 m/s. To overcome this AV deficit, aerobraking techniques were employed. However, damage discovered to one of the spacecraft's two solar arrays after launch forced major revisions to the original aerobraking planning of the MGS mission. In order to avoid a complete structural failure of the array, peak dynamic pressure levels for the spacecraft were established at a major spacecraft health review in November 1997. These peak dynamic pressure levels were roughly one-third of the original mission design values. Incorporating the new dynamic pressure limitations into mission replanning efforts resulted in an 'extended' orbit insertion phase for the mission. This 'extended' orbit insertion phase was characterized by two distinct periods of aerobraking separated by an aerobraking hiatus that would last for several months in an intermediate orbit called the "Science Phasing Orbit" (SPO). This paper describes and focuses on the strategy for the second phase of aerobraking for the MGS mission called "Aerobraking Phase 2." This description will include the baseline aerobraking flight profile, the trajectory control methodology, as well as the key trajectory metrics that were monitored in order to successfully "guide' the spacecraft to its desired mapping orbit. Additionally, the actual aerobraking progress is contrasted to the planned aerobraking flight profile. (A separate paper will describe the navigation aspects of MGS aerobraking in detail.) Key to the success of the MGS mission is the delivery of the spacecraft to its final mapping orbit and the synergy the instrument complement provides to its scientific investigators when science data is returned from that orbit. The MGS mapping orbit is characterized as a low altitude, near-circular, near-polar orbit that is Sun-synchronous with the descending equatorial crossing at 2:00 AM local mean solar time (LMST).

Johnston, M. D.↗

NASA NTRS1961

Estimates of Minimum Energy Requirements for Range-Controlled Return of a Nonlifting Satellite from a Circular Orbit

Existing expressions are used to obtain the minimum propellant fraction required for return from a circular orbit as a function of vacuum trajectory range. trajectory are matched to those of the atmospheric trajectory to obtain a complete return from orbit to earth. The results are restricted by the assumptions of (1) impulsive velocity change, (2) nearly circular transfer trajectory, ( 3) spherical earth, atmosphere, and gravitational field, (4) exponential atmospheric density variation with attitude, and (5) a nonrotating atmosphere. The solutions for the parameters of the vacuum Calculations are made t o determine the effects of longitudinal and lateral range on required propeUant fraction and reentry loading for a nonrotating earth and for several orbital altitudes. the single- and two-impulse method of return is made and the results indicate a "trade off" between propellant fraction required and landing- position accuracy. A comparison of An example of a return mission from a polar orbit is discussed where the initial deorbit point is the intersection of the North Pole horizon with the satellite orbit. Some effects of a rotating earth are also considered. It is found that, for each target-orbital-plane longitudinal difference, there exists a target latitude for which the required propellant fraction is a minimum.

Jackson, Charlie M., Jr.↗

NASA NTRS1985

Satellite tour design for the Cassini mission

The mission design for the Cassini satellite tour of Saturn is described. The use of Titan encounters for gravity-assisted trajectory shaping to satisfy atmosphere, magnetosphere, and satellite science requirements is discussed. Three candidate satellite tours that illustrate the tradeoffs between different strategies to satisfy the science objectives are presented. Tour characteristics for a 4yr mission include at least 30 Titan encounters to provide a minimum of 90 deg orbit rotation for either a magnetotail petal orbit or noon petal orbit; evenly distributed Titan ground tracks for radar mapping coverage; at least 2 Iapetus and 1 Enceladus close flybys; inclined orbits during the first 3 yr to provide Earth and Sun occultations by Saturn and its rings; and Polar orbits during the fourth year.

Diehl, R. E.↗

NASA NTRS1991

Mars sample return mission: Two alternate scenarios

Two scenarios for accomplishing a Mars Sample Return mission are presented. Mission A is a low-cost, low-mass scenario, while Mission B is a high-technology, high-science alternative. Mission A begins with the launch of one Titan 4 rocket with a Centaur G' upper stage. The Centaur performs the trans-Mars injection burn and is then released. The payload consists of two lander packages and the orbital transfer vehicle, which is responsible for supporting the landers during launch and interplanetary cruise. Near Mars, the landers separate - one bound for a polar site and the other for an equatorial site. After descending to the surface, the landers deploy small, local rovers to collect samples. The rovers return these samples to the landers for loading on the direct return rockets, which return the samples directly to the Earth's surface. Mission B starts with four Titan 4 launches, used to place the components of the planetary transfer vehicle (PTV) into orbit. The fourth launch payload is able to move to assemble the entire vehicle by simple docking routines. Once complete, the PTV begins a low-thrust trajectory out from low Earth orbit, through interplanetary space, and into low Mars orbit. It deploys a communications satellite into a one half sol orbit and then releases the lander pad cage at 500 km altitude. The lander package contains the lander, the Mars ascent vehicle (MAV), two lighter-than-air rovers (called Aereons), and one conventional land rover. The entire package is contained within a biconic aeroshell. After release from the PTV the lander package descends to the surface, where all three rovers are released to collect samples and map the terrain. The Aereons attempt to circumnavigate Mars and collect samples from a wide variety of sites, while the land rover examines a local area more thoroughly. The Aereons are equipped with small sample return rockets that can return their samples to the lander in the event that an Aereon is incapable of returning to the lander itself. Once all samples have been collected they are loaded onto the MAV and launched into orbit. The PTV then collects the samples and returns them to Earth orbit for recovery.

Source record↗

NASA NTRS2021

Tracking aerosols and SO2 clouds from the Raikoke eruption: 3D view from satellite observations

The 21 June 2019 eruption of the Raikoke volcano (Kuril Islands, Russia; 48° N, 153° E) produced significant amounts of volcanic aerosols (sulfate and ash) and sulfur dioxide (SO2) gas that penetrated into the lower stratosphere. The dispersed SO2 and sulfate aerosols in the stratosphere were still detectable by multiple satellite sensors for many months after the eruption. For this study of SO2 and aerosol clouds we use data obtained from two of the Ozone Mapping and Profiler Suite sensors on the Suomi National Polar-orbiting Partnership satellite: total column SO2 from the Nadir Mapper and aerosol extinction profiles from the Limb Profiler as well as other satellite data sets. We evaluated the limb viewing geometry effect (the “arch effect”) in the retrieval of the LP standard aerosol extinction product at 674 nm. It was shown that the amount of SO2 decreases with a characteristic period of 8–18 d and the peak of stratospheric aerosol optical depth recorded at a wavelength of 674 nm lags the initial peak of SO2 mass by 1.5 months. Using satellite observations and a trajectory model, we examined the dynamics of an unusual atmospheric feature that was observed, a stratospheric coherent circular cloud of SO2 and aerosol from 18 July to 22 September 2019.

Eruption of the Raikoke volcano↗

NASA NTRS1999

Cloud Statistics for NASA Climate Change Studies

The Principal Investigator participated in two field experiments and developed a global data set on cirrus cloud frequency and optical depth to aid the development of numerical models of climate. Four papers were published under this grant. The accomplishments are summarized: (1) In SUCCESS (SUbsonic aircraft: Contrail & Cloud Effects Special Study) the Principal Investigator aided weather forecasters in the start of the field program. A paper also was published on the clouds studied in SUCCESS and the use of the satellite stereographic technique to distinguish cloud forms and heights of clouds. (2) In SHEBA (Surface Heat Budget in the Arctic) FIRE/ACE (Arctic Cloud Experiment) the Principal Investigator provided daily weather and cloud forecasts for four research aircraft crews, NASA's ER-2, UCAR's C-130, University of Washington's Convert 580, and the Canadian Atmospheric Environment Service's Convert 580. Approximately 105 forecasts were written. The Principal Investigator also made daily weather summaries with calculations of air trajectories for 54 flight days in the experiment. The trajectories show where the air sampled during the flights came from and will be used in future publications to discuss the origin and history of the air and clouds sampled by the aircraft. A paper discussing how well the FIRE/ACE data represent normal climatic conditions in the arctic is being prepared. (3) The Principal Investigator's web page became the source of information for weather forecasting by the scientists on the SHEBA ship. (4) Global Cirrus frequency and optical depth is a continuing analysis of global cloud cover and frequency distribution are being made from the NOAA polar orbiting weather satellites. This analysis is sensitive to cirrus clouds because of the radiative channels used. During this grant three papers were published which describe cloud frequencies, their optical properties and compare the Wisconsin FM Cloud Analysis to other global cloud data such as the International Satellite Cloud Climatology Program (ISCCP) and the Stratospheric Aerosol and Gas Experiment (SAGE). A summary of eight years of HIRS data will be published in late 1998. Important information from this study are: 1) cirrus clouds cover most of the earth, 2) they are found about 40% of the time globally, 3) in the tropics cirrus cloud frequencies are even higher, from 80-100%, 4) there is slight evidence that cirnis cloud cover is increasing in the northern hemisphere at about 0.5% per year, and 5) cirrus clouds have an average infrared transmittance of about 40% of the terrestrial radiation. (5) Global Cloud Frequency Statistics published on the Principal Investigator's web page have been used in the planning of the future CRYSTAL experiment and have been used for refinements of a global numerical model operated at the Colorado State University.

Wylie, Donald P.↗

NASA NTRS1995

Guidance, navigation, and control study for a solar electric propulsion spacecraft

A preliminary investigation of a lunar-comet rendezvous mission using a solar electric propulsion (SEP) spacecraft was performed in two phases.The first phase involved exploration of the moon and the second involved rendezvous with a comet. The initial phase began with a chemical propulsion translunar injection and chemical insertion into a lunar orbit, followed by a low thrust SEP transfer to a circular, polar, low-lunar orbit. After collecting scientific data at the moon, the SEP spacecraft performed a spiral lunar escape maneuver to begin the interplanetary leg of the mission. After escape from the Earth-moon system, the SEP spacecraft maneuvered in interplanetary space and performed a rendezvous with a comet.The immediate goal of this study was to demonstrate the feasibility of using a low-thrust SEP spacecraft for orbit transfer to both the moon and a comet. Another primary goal was to develop a computer optimization code which would be robust enough to obtain minimum-fuel rendezvous trajectories for a wide range of comets.

Kluever, Craig A.↗

NASA NTRS1995

Optimal lunar trajectories for a combined chemical-electric propulsion spacecraft

Spacecraft which utilize electric propulsion (EP) systems are capable of delivering a greater payload fraction compared to spacecraft using conventional chemical propulsion systems. Several researchers have investigated numerous applications of low-thrust EP including a manned Mars mission, scientific missions to the outer planets, and lunar missions. In contrast, the study of optimal combined high and low-thrust spacecraft trajectories has been limited. In response to the release of NASA's 1994 Announcement of Opportunity (AO) for Discovery class interplanetary exploration missions, a preliminary investigation of a lunar comet rendezvous mission using a solar electric propulsion (SEP) spacecraft was performed. The Discovery mission (eventually named Diana) was envisioned to be a two-phase scientific exploration mission: the first phase involved exploration of the moon and second phase involved rendezvous with a comet. The initial phase began with a chemical propulsion translunar injection and chemical insertion into a lunar orbit, followed by a low-thrust SEP transfer to a circular, polar, low-lunar orbit (LLO). After scientific data was collected at the moon, the SEP spacecraft performed a spiral lunar escape maneuver to begin the interplanetary leg of the mission. After escape from the Earth-moon system, the SEP spacecraft maneuvered in interplanetary space and performed a rendezvous with a short period comet. An initial study that demonstrated the feasibility of using EP for the lunar and comet orbit transfer was performed under the grant NAG3-1581. This final report is a continuation of the initial research efforts in support of the Discovery mission proposal that was submitted to NASA Headquarters in October 1994. Section 2 discusses the lunar orbit transfer phase of the Diana mission which involves both chemical and electric propulsion stages. Section 3 discusses the chemical lunar orbit insertion (LOI) burn optimization. Finally, section 4 presents the conclusions of this research effort.

Kluever, Craig A.↗

NASA NTRS2004

NASCAP-2K: An Overview

Nascap-2k is the modern replacement for the older 3-D charging codes NASCAP/GEO, NASCAP/LEO, POLAR, and DynaPAC. Built on the DynaPAC kernel and incorporating surface charging, environment and space potential models from the older codes, Nascap-2k performs charging calculations for a wide variety of space environments under control of a unified graphical interface. In this paper we illustrate the use of Nascap-2k for spacecraft charging calculations. We touch on some of the unique physical and mathematical models on which the code is based. Examples/demos include the use of Object Toolkit, charging calculations in geosynchronous substorm, solar wind, low earth orbit, and auroral environments, and display and analysis of surface potentials, space potentials and particle trajectories.

Mandell, M. J.↗

NASA NTRSDate not supplied

Design and Optimization of Steering Laws for Geocentric Solar Sailing

Solar sail technology has the potential to be a practical method of propulsion for future deep space Cubesats and small satellites. Although in general, deep space, heliocentric flight is the optimal regime for solar sailing, many near-term solar sail flight demonstrations have, or will occur in Earth-centered orbits. This is particularly true for low-cost, smallsat-based solar sail technology demonstration flights, such as the Planetary Society LightSail 1 and 2, and the upcoming NASA Advanced Composite Solar Sail System (ACS3). ACS3 will be the first spaceflight application of NASA’s newest smallsat deployable composite boom technology, and NASA’s first practical solar sail. Launch of ACS3 is currently anticipated for late 2021 or early 2022. In this paper, we will describe the challenges of operating solar sailing spacecraft in low Earth orbits, where, depending on initial orbital altitude and inclination, disturbances due to aerodynamic drag, eclipsing can be of the same order or greater than forces caused by solar radiation pressure. A notional smallsat-class solar sailing vehicle based on the ACS3 solar sail will be used as a reference for trajectory design and optimization studies. As a starting point we will assume initially circular orbits, and a locally optimal steering law designed to maximize the rate of change in semi major axis, similar to that described by McInnes. This will be generalized to orbits with arbitrary initial inclinations and non-zero eccentricities, and to account for atmospheric drag and eclipsing effects. Optimal steering laws for maximizing rate-of-change in orbit inclination and eccentricity will also be explored. Finally, we will examine potential for Earth-centered solar sailing mission applications beyond technology demonstrations, including: continuous observations from inside the Earth’s magnetotail; augmented geostationary capabilities from displaced geostationary orbits; and continuous polar observation for climate science investigations from pole-sitter orbits.

W. Keats Wilkie↗

NASA NTRS2002

New Approaches for Exploration Beyond Low Earth Orbit

Through the work of the NASA Exploration Team (NEXT), innovations in mission architectures have been developed. Missions beyond low Earth orbit have been dramatically simplified through new combinations in the application of advanced technologies and mission design. These concepts enable a stepping stone approach to discovery driven, technology enabled exploration. The architectures have evolved from the interest in easily reaching the poles of the moon, to take advantage of the water ice that is believed to exist in the permanently dark craters. The Polar Regions also have areas that are consistently sunlit, providing a constant source of power. By staging these missions from the Earth-Moon Lagrange point, L1, other sites on the lunar surface can be reached as easily as the polar sites. Through studying this approach to lunar exploration, and combining new trajectory approaches the staging of other missions from L1 has begun to look attractive. The possibilities of missions include the assembly and maintenance of arrays of telescopes for emplacement at the Earth-Sun L2, and travel to Mars. The resulting numbers and masses of vehicles required are greatly reduced. Vehicle design concepts are developed for proof of concept, to validate mission approaches and understand the value of new technologies. The stepping stone approach employs an incremental buildup of capabilities; allowing for decision points on exploration objectives. It enables testing of technologies to achieve greater reliability and understanding of costs for the next steps in exploration.

Cooke, D. R.↗

NASA NTRS2017

Juno Orbit Determination Experience During First Year At Jupiter

The Juno spacecraft successfully inserted into a polar orbit around Jupiter on 5- July-2016. Since the Jupiter Orbit Insertion (JOI) maneuver, Juno has completed six orbits around Jupiter. The mission plan at the time of JOI was for Juno to perform two 53.5-day capture orbits before executing a Period Reduction Maneuver (PRM) to place the spacecraft into its intended 14-day science orbit. This maneuver was canceled due to a concern with the propulsion system. As a result, the Juno spacecraft will remain in its longer orbit period for rest of its mission. This paper discusses the Navigation Team’s experience: the orbit determination strategy and how it changed due to the cancellation of the PRM, challenges fitting the data during perijove, and how we reconstructed the trajectory during Juno’s first year in orbit.

Thompson, Paul↗

NASA NTRS2010

Orbit Determination Toolbox

The Orbit Determination Toolbox is an orbit determination (OD) analysis tool based on MATLAB and Java that provides a flexible way to do early mission analysis. The toolbox is primarily intended for advanced mission analysis such as might be performed in concept exploration, proposal, early design phase, or rapid design center environments. The emphasis is on flexibility, but it has enough fidelity to produce credible results. Insight into all flight dynamics source code is provided. MATLAB is the primary user interface and is used for piecing together measurement and dynamic models. The Java Astrodynamics Toolbox is used as an engine for things that might be slow or inefficient in MATLAB, such as high-fidelity trajectory propagation, lunar and planetary ephemeris look-ups, precession, nutation, polar motion calculations, ephemeris file parsing, and the like. The primary analysis functions are sequential filter/smoother and batch least-squares commands that incorporate Monte-Carlo data simulation, linear covariance analysis, measurement processing, and plotting capabilities at the generic level. These functions have a user interface that is based on that of the MATLAB ODE suite. To perform a specific analysis, users write MATLAB functions that implement truth and design system models. The user provides his or her models as inputs to the filter commands. The software provides a capability to publish and subscribe to a software bus that is compliant with the NASA Goddard Mission Services Evolution Center (GMSEC) standards, to exchange data with other flight dynamics tools to simplify the flight dynamics design cycle. Using the publish and subscribe approach allows for analysts in a rapid design center environment to seamlessly incorporate changes in spacecraft and mission design into navigation analysis and vice versa.

Carpenter, James R.↗

NASA NTRS2013

In-flight Operation of the Dawn Ion Propulsion System Through Year One of Cruise to Ceres

The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt asteroids, 4 Vesta, and the dwarf planet 1 Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 rocket that placed the 1218-kg spacecraft into an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide an additional delta- V of approximately 11 km/s for the heliocentric transfers to each body and for all orbit transfers including orbit capture/escape and transition to the various science orbits. Deterministic thrusting to Vesta began in December 2007 and concluded with orbit capture at Vesta in July 2011. The transfer to Vesta included a Mars gravity assist flyby in February 2009 that provided an additional delta-V of 2.6 km/s and was the only postlaunch mission delta-V not provided by IPS. During the mission at Vesta the IPS was used for all orbit transfers which included six different near-polar science mapping orbits. Thrusting for departure from Vesta and the start of cruise to Ceres began on July 25, 2012 with escape from Vesta occurring on September 5, 2012. To date the IPS has been operated for approximately 31,000 hours, consumed approximately 300 kg of xenon, and provided a delta-V of approximately 8.3 km/s. IPS performance characteristics are very close to the expected performance based on analysis and testing performed prelaunch. Thrusting for cruise to Ceres will continue until the spring of 2015, with a planned arrival date at Ceres in April 2015. This paper provides an overview of Dawn's mission objectives and the results of Dawn IPS mission operations from Vesta departure through the first year of cruise to Ceres.

Hayabusa spacecraft↗

NASA NTRSDate not supplied

End-to-End Performance Optimization of a Crewed Lunar Landing Mission Staged from a Near Rectilinear Halo Orbit

A crewed lunar landing mission staged from a Near Rectilinear Halo Orbit (NRHO) presents a unique set of challenges not considered before for the Apollo-type direct missions. The NRHO does not have a fixed orbital plane and its period also changes with time, due to which the mission performance is very sensitive to the timing and location of the NRHO departure and arrival burns for a near-polar surface landing site. For such type of mission staged from the NRHO, an end-to-end optimization approach for the entire mission, from the NRHO departure to surface landing and from surface lift-off to the NRHO insertion, is well-suited for extracting the maximum performance out of this design. In this paper, an integrated approach for performing end-to-end optimization of a NRHO-staged crewed lunar landing mission is proposed that optimizes the in-space and powered flight segments simultaneously using Copernicus trajectory optimization tool. Copernicus is augmented with analytical approximations for the in-space, powered descent, and powered ascent segments in this work to improve the convergence of its multiple-shooting based direct optimization method. New analytical solutions are derived for the different phases of the powered flight segments by solving associated guidance problems. These analytical solutions act as initial guesses for starting the numerical optimization of the integrated mission in Copernicus. An example lunar landing mission is simulated to validate the proposed end-to-end optimization approach and the results are compared with the conventional patched trajectory approach.

Bharat Mahajan↗

DOE OSTI2020

Transformation of Spinors in Accelerators and Beam Transfer Lines

The stable spin direction and the spin tune are two important concepts used when polarized beams are transferred along beam lines or circulating in an accelerator. One of the methods to calculate the stable spin direction and the spin tune of the reference particle of a polarized beam bunch is by integrating simultaneously the Lorentz Force $\vec{F}$ = q($\vec{E}$ +$\vec{v}$ x $\vec{B}$) and the BMT equation of the reference particle as it is transported by the magnets of a beam line or accelerator. A modified version of the Raytrace code was used to integrate the Lorentz equation and the BMT equation. Using the spin motion which is governed by the BMT equation one can compute both, the stable spin direction of the reference particle at any point along the reference orbit of an accelerator or a beam line, and the spin tune of the closed orbit in an accelerator. The simultaneous integration of the Lorentz force and of the BMT equation is the most accurate method to calculate the trajectory and the spin motion of a charged particle moving in an electromagnetic field, however this method requires the accurate description of the electromagnetic fields of the elements in which the charged particles are moving. An alternative method of calculating the stable spin direction and the spin tune of spin ½ charged particles along the reference trajectory of a beam line or an accelerator is the transformation of the spinor which is the wavefunction of spin ½ particles. This method requires the knowledge of the stable spin direction of each element of the beam line or the accelerator. This technical note describes the “spinor-method” to calculate the stable-spin- direction and the spin-tune of a spin ½ particle moving along the reference trajectory of a beam line or an accelerator.

43 PARTICLE ACCELERATORS↗

NASA NTRS1989

EX Hydrae in outburst

Photometry and red spectroscopy of the intermediate polar EX Hya in its rare outburst state are presented. Photometry during the declining phase of the July-August 1986 outburst shows the 67-min (spin) modulation to be present with similar characteristics to that in quiescence. In contrast, photometry from near the peak of the 1987 May outburst shows little evidence of the 67-min modulation, while spectroscopy obtained nearly simultaneously is similarly lacking in such evidence, despite its presence in quiescent spectroscopic data. Near the beginning of the May 1987 outburst the H alpha emission line develops a broad, high velocity base component whose velocity is modulated with the orbital cycle. The velocity and phase of the broad base component suggest that it is produced near the magnetosphere of the white dwarf at a point along the projected trajectory of the gas stream from the companion. The feature disappears later in the outburst and is not present during quiescence. It is suggested that the outbursts in EX Hya are caused by an increase in the rate of mass transfer from the companion, and that part of this enhanced mass-transfer stream skims over the top of the accretion disk to strike the magnetosphere directly. The interaction of the stream with the magnetosphere gives rise to the broad-base component observed.

Hellier, C.↗

NASA NTRS1997

Geotail MCA Plasma Wave Investigation Data Analysis

The primary goals of the International Solar Terrestrial Physics/Global Geospace Science (ISTP/GGS) program are identifying, studying, and understanding the source, movement, and dissipation of plasma mass, momentum, and energy between the Sun and the Earth. The GEOTAIL spacecraft was built by the Japanese Institute of Space and Astronautical Science and has provided extensive measurements of entry, storage, acceleration, and transport in the geomagnetic tail and throughout the Earth's outer magnetosphere. GEOTAIL was launched on July 24, 1992, and began its scientific mission with eighteen extensions into the deep-tail region with apogees ranging from around 60 R(sub e) to more than 208 R(sub e) in the period up to late 1994. Due to the nature of the GEOTAIL trajectory which kept the spacecraft passing into the deep tail, GEOTAIL also made 'magnetopause skimming passes' which allowed measurements in the outer magnetosphere, magnetopause, magnetosheath, bow shock, and upstream solar wind regions as well as in the lobe, magnetosheath, boundary layers, and central plasma sheet regions of the tail. In late 1994, after spending nearly 30 months primarily traversing the deep tail region, GEOTAIL began its near-Earth phase. Perigee was reduced to 10 R(sub e) and apogee first to 50 R(sub e) and finally to 30 R(sub e) in early 1995. This orbit provides many more opportunities for GEOTAIL to explore the upstream solar wind, bow shock, magnetosheath, magnetopause, and outer magnetosphere as well as the near-Earth tail regions. The WIND spacecraft was launched on November 1, 1994 and the POLAR spacecraft was launched on February 24, 1996. These successful launches have dramatically increased the opportunities for GEOTAIL and the GGS spacecraft to be used to conduct the global research for which the ISTP program was designed. The measurement and study of plasma waves have made and will continue to make important contributions to reaching the ISTP/GGS goals and solving the significant problems of sun-earth connections. Plasma waves are involved in the energization and de-energization of plasma and energetic particles via numerous wave-particle interaction processes. Plasma waves in many instances are the source for the heating or cooling of the particles. They can cause particle precipitation by scattering particles into the loss cone. They move particles across boundaries in mass and energy dependent ways. Identifying the waves and the instabilities which produce them are thus crucial for understanding the plasma processes. Wave-particle interaction processes are especially important at various boundaries between the different regions of geospace including the bow shock, magnetopause, and interfaces in the geomagnetic tail between the magnetosheath, lobe, plasmasheet, boundary layers, and neutral sheet. In addition to identifying the characteristics of the instabilities and generation mechanisms encountered, plasma wave measurement are used in conjunction with other fields and particle measurements to identify the region of space the spacecraft is in or the boundary that is being crosed.

Anderson, Roger R.↗