Launch Window Research Articles

An Analysis of the Lifetimes of Large Elliptical Orbits of Space Debris

The lunisolar perturbation is the principal factor that causes variation in the orbit of a GTO (geosychronous transfer orbit) space debris. In the present article the analytical expression of the long-periodic term of the altitude of the GTO orbital perigee caused by the lunisolar perturbation is derived, and the basic characteristics of the variation of the perigee caused by the lunisolar perturbation are analyzed. A method of selecting a launch window that will help to improve the space debris situation is given, and its correctness is tested and verified through simulation calculation.

NASA Mars mission delayed to 2011

NASA's Mars Science Laboratory has been rescheduled to launch in the fall of 2011, two years later than previously planned. The decision to delay, announced on 4 December, was made following an assessment that more time is needed for testing and to resolve hardware challenges. The length of the delay is because the 2009 launch window ends in October, with the next launch opportunity in 2011.The mission will study the early environmental history of Mars, exploring a site where there were wet conditions in the past that may have been favorable for supporting microbial life.

Locating the solar source of 13 April 2006 magnetic cloud

Abstract. Using Advanced Composition Explorer (ACE) in situ data we identify and describe an interplanetary magnetic cloud (MC) observed near Earth on 13 April 2006. We also use multi-instrument and multi-wavelength observations from the Solar and Heliospheric Observatory (SOHO), the Transition Region and Coronal Explorer (TRACE) and ground-based solar observatories to determine the solar source of this magnetic cloud. A launch window for the MC between 9 and 11 April 2006 was estimated from the propagation time of the ejecta observed near Earth. A number of large active regions (ARs) were present on the Sun during this period, which were initially considered to be the most likely candidate source regions of the MC. However, it was determined that the solar source of the MC was a small, spotless active region observed in the Northern Hemisphere. Following an eruption from this region on 11 April 2006, the ACE spacecraft detected, 59 h later, the passage of the MC, preceded by the arrival of a weak, forward fast shock. The link between the eruption in this active region and the interplanetary MC is supported by several pieces of evidence, including the location of the solar source near to the disk centre and to the east of the central meridian (in agreement with the spacecraft trajectory through the western leg of the magnetic cloud), the propagation time of the ejecta, the agreement between the amount of flux in the magnetic cloud and in the active region, and the agreement between the signs of helicity of the magnetic cloud and the active region (which differs from the sign of helicity of each of the other active regions on the Sun at this time). In addition, the active region is located on the boundary of a coronal hole, and a high speed solar wind stream originating from this region is observed near Earth shortly after the passage of the magnetic cloud.

Launch windows for libration point missions

The four upcoming ESA missions to the collinear Lagrange points in the sun–Earth system, Herschel and Planck, GAIA, and LISA Pathfinder, will be launched from French Guiana by different Rockets, namely Ariane 5, Soyuz, and VEGA. Herschel and Planck will be launched together in 2008 on an Ariane 5. The launcher injects Herschel directly into the stable manifold of a large amplitude quasi-Halo orbit at L 2 . Departing from this launch orbit, Planck performs a three manoeuvre transfer to a Lissajous orbit, on which the sun–spacecraft–Earth angle remains below 15 ∘ . The daily and seasonal launch window for Herschel/Planck will be constrained by sun aspect angle conditions, by eclipse conditions, and by the propellant allocation on the Planck satellite. GAIA will be launched by Soyuz/Fregat end 2011. The target orbit at L 2 is similar to that of Planck, with a constrained sun–spacecraft–Earth angle. Initially, to reach a full year launch window, a launch scenario was preferred using a nearly circular parking orbit at 19 ∘ inclination, before igniting the Fregat upper stage to enter into the transfer. Recently alternative launch scenarios have been introduced, to guarantee a controlled re-entry of the Soyuz third stage, but leading to a reduction of the launch window. LISA Pathfinder will be delivered 2010 by the VEGA launcher to a low inclination elliptical orbit, with 1624 km apogee height. From there an apogee raising manoeuvre sequence, using a propulsion module connected to the spacecraft, reaches L 1 transfer conditions. Launch window calculations for these projects combine the choice of lift-off time and launch orbit conditions, with the selection of the target orbit at the Lagrange points and the minimisation of propellant to reach that orbit. For Herschel and LISA Pathfinder the target orbits have been chosen such that a transfer without deterministic manoeuvres is possible. Planck and GAIA both use a multiple manoeuvre fast transfer strategy reaching the Lissajous orbit at L 2 within 50 days.

Optimal Lunar Launch Trajectories toSun-Earth L1 Vicinity

The sun/Earth-moon L1 point is a growing target for space missions. This paper assumes a need to place an object directly between the Earth and the sun. Possible purposes for such a mission include a type of shield to intercept some portion of the sun's rays, or conversely a lens or mirror to focus them. In addition, this paper requires a launch directly from the moon instead of a traditional orbit-to-orbit transfer. A detailed case using accurate positional and velocity data for all planetary bodies is used in this paper. Instead of focusing solely on trajectories defined as optimal by minimum fuel consumption, we looked for trajectories with low sensitivity to initial conditions to additionally minimize the frequency and magnitude of course corrections. We first analyzed the reverse problem to identify solution areas of low sensitivity to launch conditions. Using the determined restricted solution space, we were successfully able to determine a launch site on the moon that could reach L1 with the greatest number of low-cost trajectories and launch windows within the resolution of the solution space.

Microbial Rock Inhabitants Survive Hypervelocity Impacts on Mars-Like Host Planets: First Phase of Lithopanspermia Experimentally Tested

The scenario of lithopanspermia describes the viable transport of microorganisms via meteorites. To test the first step of lithopanspermia, i.e., the impact ejection from a planet, systematic shock recovery experiments within a pressure range observed in martian meteorites (5-50 GPa) were performed with dry layers of microorganisms (spores of Bacillus subtilis, cells of the endolithic cyanobacterium Chroococcidiopsis, and thalli and ascocarps of the lichen Xanthoria elegans) sandwiched between gabbro discs (martian analogue rock). Actual shock pressures were determined by refractive index measurements and Raman spectroscopy, and shock temperature profiles were calculated. Pressure-effect curves were constructed for survival of B. subtilis spores and Chroococcidiopsis cells from the number of colony-forming units, and for vitality of the photobiont and mycobiont of Xanthoria elegans from confocal laser scanning microscopy after live/dead staining (FUN-I). A vital launch window for the transport of rock-colonizing microorganisms from a Mars-like planet was inferred, which encompasses shock pressures in the range of 5 to about 40 GPa for the bacterial endospores and the lichens, and a more limited shock pressure range for the cyanobacterium (from 5-10 GPa). The results support concepts of viable impact ejections from Mars-like planets and the possibility of reseeding early Earth after asteroid cataclysms.

On the design of low‐thrust transfers through L1 for the ESMO mission

AbstractThe following paper summarises the mission analysis studies performed for the phase A of the solar electric propulsion option of the European Student Moon Orbiter (ESMO) mission. ESMO is scheduled to be launched in 2011, as an auxiliary payload on board an Ariane 5. Hence the lack of influence on the launch date. Two methods were devised to assess efficiently wide launch windows for the Earth‐Moon transfer, they are presented below. (© 2008 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)

New Horizons Mission Design

In the first mission to Pluto, the New Horizons spacecraft was launched on January 19, 2006, and flew by Jupiter on February 28, 2007, gaining a significant speed boost from Jupiter’s gravity assist. After a 9.5-year journey, the spacecraft will encounter Pluto on July 14, 2015, followed by an extended mission to the Kuiper Belt objects for the first time. The mission design for New Horizons went through more than five years of numerous revisions and updates, as various mission scenarios regarding routes to Pluto and launch opportunities were investigated in order to meet the New Horizons mission’s objectives, requirements, and goals. Great efforts have been made to optimize the mission design under various constraints in each of the key aspects, including launch window, interplanetary trajectory, Jupiter gravity-assist flyby, Pluto–Charon encounter with science measurement requirements, and extended mission to the Kuiper Belt and beyond. Favorable encounter geometry, flyby trajectory, and arrival time for the Pluto–Charon encounter were found in the baseline design to enable all of the desired science measurements for the mission. The New Horizons mission trajectory was designed as a ballistic flight from Earth to Pluto, and all energy and the associated orbit state required for arriving at Pluto at the desired time and encounter geometry were computed and specified in the launch targets. The spacecraft’s flight thus far has been extremely efficient, with the actual trajectory error correction ΔV being much less than the budgeted amount.

Orbit concepts at [formula omitted] for Soyuz launches from Kourou

Many of the currently planned astronomy missions will go into orbits around the L 2 libration point in the Sun–Earth system. Two ESA projects at L 2 , GAIA (2012) and DARWIN (2016) will use a Soyuz launch from the currently constructed launch facility at Kourou. Soyuz will launch into a low circular parking orbit from which the Fregat upper stage will inject to the transfer. The nearly parabolic transfer orbits can be reached at inclinations of up to 30 ∘ with little mass penalty compared to low inclination orbits. This allows to reach a Lissajous orbit with a maximum Sun–spacecraft–Earth angle of 15 ∘ (GAIA baseline) with an insertion maneuver of less than 130 m/s for a full year launch window. Or by selecting the parking orbit near the lunar orbit plane, a lunar gravity assist can be used to reach the stable manifold of an even smaller amplitude (e.g. 7 ∘ ) Lissajous orbit without any deterministic maneuver during the transfer. Another interesting case of transfers to orbits around L 2 has been studied for the DARWIN mission. DARWIN consists of four spacecraft. One studied option is to launch them in pairs, on two Soyuz launches from Kourou. The launch dates have to be at least 20 days apart, but the spacecraft have to rendezvous in an orbit around L 2 . It has been shown that without any phasing orbits for a given separation of the launch dates of up to 60 days, a target orbit can be defined in which the two stacks of spacecraft can be brought to the same position with a small Δ V .

KSLV-III를 이용한 한국형 화성 탐사 임무의 설계

Mission opportunities and trajectory characteristics for the future Korean Mars mission have designed and analyzed using KSIV-III(Korea Space Launch Vehicle-III). Korea`s first space center, `NARO space center` is selected as a launch site. For launch opportunities, year 2033 is investigated under considering the date of space center`s completion with KSLV series development status. Optimal magnitude of various maneuvers, Trans Mars Injection (TMI) maneuver, Trajectory Correction Maneuver (TCM), Mars Orbit Insertion (MOI) maneuver and Orbit Trim Maneuver(OTM), which are required during the every Mars mission phases are computed with the formulation of nonlinear optimization problems using NPSOL software. Finally, mass budgets for upper stage (launcher for KSIV-III and spacecraft are derived using various optimized maneuver magnitudes. For results, daily launch window from NARO space center for successful Korean Mars mission is avaliable for next 27 minutes starting from Apr. 16. 2033. 12:17:26 (UTC). Maximum spacecraft gross mass which can delivered to Mars is about 206kg, with propellant mass of 109kg and structure mass of 97kg, when on board spacecraft thruster`s Isp is assumed to have 290 sec. For upper stage, having structure ratio of 0.15 and Isp value of 280 sec, gross mass is about 1293kg with propellant mass of 1099kg and structure mass of 194kg. However, including 10% margins to computed optimal maneuver values, spacecraft gross mass is reduced to about 148kg with upper stage`s mass of 1352kg. This work will give various insights, requiring performances to developing of KSIV-III and spacecraft design for future Korean Mars missions.

The MaCWAVE program to study gravity wave influences on the polar mesosphere

Abstract. MaCWAVE (Mountain and Convective Waves Ascending VErtically) was a highly coordinated rocket, ground-based, and satellite program designed to address gravity wave forcing of the mesosphere and lower thermosphere (MLT). The MaCWAVE program was conducted at the Norwegian Andøya Rocket Range (ARR, 69.3° N) in July 2002, and continued at the Swedish Rocket Range (Esrange, 67.9° N) during January 2003. Correlative instrumentation included the ALOMAR MF and MST radars and RMR and Na lidars, Esrange MST and meteor radars and RMR lidar, radiosondes, and TIMED (Thermosphere Ionosphere Mesosphere Energetics and Dynamics) satellite measurements of thermal structures. The data have been used to define both the mean fields and the wave field structures and turbulence generation leading to forcing of the large-scale flow. In summer, launch sequences coupled with ground-based measurements at ARR addressed the forcing of the summer mesopause environment by anticipated convective and shear generated gravity waves. These motions were measured with two 12-h rocket sequences, each involving one Terrier-Orion payload accompanied by a mix of MET rockets, all at ARR in Norway. The MET rockets were used to define the temperature and wind structure of the stratosphere and mesosphere. The Terrier-Orions were designed to measure small-scale plasma fluctuations and turbulence that might be induced by wave breaking in the mesosphere. For the summer series, three European MIDAS (Middle Atmosphere Dynamics and Structure) rockets were also launched from ARR in coordination with the MaCWAVE payloads. These were designed to measure plasma and neutral turbulence within the MLT. The summer program exhibited a number of indications of significant departures of the mean wind and temperature structures from ``normal" polar summer conditions, including an unusually warm mesopause and a slowing of the formation of polar mesospheric summer echoes (PMSE) and noctilucent clouds (NLC). This was suggested to be due to enhanced planetary wave activity in the Southern Hemisphere and a surprising degree of inter-hemispheric coupling. The winter program was designed to study the upward propagation and penetration of mountain waves from northern Scandinavia into the MLT at a site favored for such penetration. As the major response was expected to be downstream (east) of Norway, these motions were measured with similar rocket sequences to those used in the summer campaign, but this time at Esrange. However, a major polar stratospheric warming just prior to the rocket launch window induced small or reversed stratospheric zonal winds, which prevented mountain wave penetration into the mesosphere. Instead, mountain waves encountered critical levels at lower altitudes and the observed wave structure in the mesosphere originated from other sources. For example, a large-amplitude semidiurnal tide was observed in the mesosphere on 28 and 29 January, and appears to have contributed to significant instability and small-scale structures at higher altitudes. The resulting energy deposition was found to be competitive with summertime values. Hence, our MaCWAVE measurements as a whole are the first to characterize influences in the MLT region of planetary wave activity and related stratospheric warmings during both winter and summer.

Optimization of a Broad Class of Ephemeris Model Earth-Mars Cyclers

Recent interest in cycler trajectories has led to the discovery of many new solutions. Most, however, are valid only in a circular-coplanar solar system. The current study presents a method to optimize accurate model cyclers based on simple model solutions and catalogs the results for a broad class of ephemeris model Earth‐Mars cyclers. An efficient constrained optimization problem is defined to accommodate long-duration, ballistic, patched conic trajectories. A continuation algorithm is then developed to transition infinitely continuous circular-coplanar solutions to finite-duration accurate ephemeris solutions. Finally, the algorithm is applied to 203 promising circular-coplanar parent cyclers for 21 launch windows each, equaling 4263 cases. In total, there are nine parent cyclers (one of them is the well-known Aldrin cycler) that have at least one finite-duration cycler with a total maneuver requirement of less than 1 m/s over seven full cycles. Additionally, 39 and 74 parent cyclers have at least one launch date with a maneuver requirement of less than 10 and 300 m/s, respectively. Several of the most promising cyclers have consistently low requirements for all launch windows considered. In general, the study demonstrates the broad feasibility for accurate ephemeris cyclers.

Reduction of Lunar Landing Fuel Requirements by Utilizing Lunar Ballistic Capture

Ballistic lunar capture trajectories have been successfully utilized for lunar orbital missions since 1991. Recent interest in lunar landing trajectories has occurred due to a directive from President Bush to return humans to the Moon by 2015. NASA requirements for humans to return to the lunar surface include separation of crew and cargo missions, all lunar surface access, and anytime-abort to return to Earth. Such requirements are very demanding from a propellant standpoint. The subject of this paper is the application of lunar ballistic capture for the reduction of lunar landing propellant requirements. Preliminary studies of the application of weak stability boundary (WSB) trajectories and ballistic capture have shown that considerable savings in low Earth orbit (LEO) mission mass may be realized, on the order of 36% less than conventional Hohmann transfer orbit missions. Other advantages, such as reduction in launch window constraints and reduction of lunar orbit maintenance propellant requirements, have also surfaced from this study.

The effect of schedule constraints on the success of planetary missions

Previous studies have shown that planetary missions have a significantly higher failure rate than earth-orbiting missions. One contributor to this higher failure rate appears to be schedule restrictions imposed on planetary missions by their fixed launch dates as opposed to the more relaxed launch schedules enjoyed in many cases by Earth-orbiting missions. Data will be presented showing the failure rates, planned vs. actual schedules, and other factors of planetary vs. Earth-orbiting missions that will test this hypothesis. Specific examples of this relationship will also be shown to further illustrate the correlation between fixed launch windows and an increased chance of mission failure.

USV test flight by stratospheric balloon: Preliminary mission analysis

The Unmanned Space Vehicle test flights will use a 7 m 1300 kg aircraft. The first three launches will take place at the Italian Space Agency ASI base in Trapani–Milo, Sicily, through a stratospheric balloon that will drop the aircraft at a predefined height. After free fall acceleration to transonic velocities, the parachute deployment will allow a safe splash down in the central Mediterranean Sea. The goal of this article is to show the preliminary analysis results for the first USV flight. We carried out a statistical study for the year 2000–2003, evaluating the typical summer and winter launch windows of the Trapani–Milo base. First, in the center Mediterranean, we define safe recovery areas. They cannot be reached during the balloon ascending phase so, after a sufficiently long floating part able to catch the open sea, the balloon will go down to the release height (24 km). The simulation foresees a 400,000 m 3 balloon and 3 valves for the altitude transfer. A safe splash down must occur far enough from the nearest coast: the minimum distance is considered around 25 km. The vehicle should be released at a distance, from the nearest coast, greater than this minimum amount plus the USV model maximum horizontal translation, during its own trajectory from balloon separation to splash down. In this way we define safe release areas for some possible translations. Winter stratospheric winds are less stable. The winter average flight duration is 7 h and it is probably too long for the diurnal recovery requirement and its scheduled procedures. Comparing past stratospheric balloons flights and trajectories computed using measured meteorological data (analysis), with their predictions made using forecast models and soundings, we obtain the standard deviation of the trajectory forecast uncertainty at the balloon–aircraft separation. Two cases are taken into account: predictions made 24 and 6 h before the launch. Assuming a Gaussian latitudinal uncertainty distribution for the prediction 6 h before the launch, we are able to identify the forecast trajectories that have a probability greater than 97% to reach the safe release areas. Simulating the summer windows trajectories for the years from 2000 to 2003 and for the favorable ground wind days, we obtain the number of trajectories with the desired forecast probabilities.

Solutions for Space Operations with a Star Finder

This paper shows a possible new use of a common or computerised star finder for a rapid and easy solution to some space problems such as: – launch windows;– satellite fly-over;– the rise and setting of a satellite;– Earth or Moon satellite surveys;– the eclipse of a satellite.

Mars express orbit design

This paper gives a survey of the orbit design, mission analysis aspects for Mars Express, the ESA mission to Mars. The emphasis is on those features which are unconventional. These are: the launch window with only 9 flight programs on the Fregat launcher upper stage for 30 launch days, and the re-targeting and Planetary protection strategy during the cruise; the delivery of the Beagle 2 lander to its entry corridor and the spacecraft re-targeting after lander separation; the capture strategy which reaches an orbit around Mars also in the case of a main engine failure; a plane turn manoeuvre during the injection sequence, and the selection of a nearly resonant eccentric orbit, using the J 2 effect of Mars to create global coverage from low altitudes at balanced sun illumination conditions for optical and radar payload, and to provide regular communications relay contacts to the lander.

Trajectory optimization design of a lunar polar probe

This article presents the optimized design, analysis and calculation concerning the trajectory of a lunar polar probe. Firstly, the trajectory design principles and constraints are determined. The preliminary design and analysis of the circumlunar orbit, transfer orbit to the moon and earth parking orbit are carried out separately and some computations for the flight trajectory concept have been made too. To reduce the fuel needed for error in orbital maneuver efficiently and satisfy the requirements on the launch window, some detailed design and analysis for the rather advanced phasing earth‐moon transfer orbit are given here and also the strategy for optimum orbit correction.

Launch vehicles for low-cost planetary missions

Delta has been synonymous with the launch of planetary missions for several decades. To make launches of all missions, including planetary missions, more affordable, launch vehicle evolution is necessary. The history of the Delta expendable launch vehicle has been one of constant evolution for almost 40 years and continues to evolve. Delta expendable launch vehicles include three basic variants: Delta II, Delta III, and Delta IV. Each vehicle is described and performance capabilities given for escape missions, GTO missions, and LEO missions. Injection accuracy for the Delta II has been better than 3 sigma. Actual injection accuracy performance figures for several planetary missions are presented. Data on launch environments are presented as well as the cleanliness environment available in the fairing and mobile service tower. Since many planetary missions depend on rather restricted launch windows, data are presented showing Delta's launch record within the launch window requirements. Delta has launched approximately 30 secondary missions over the past several years. Boeing is developing dual-payload capabilities for Delta II, Delta III, and Delta IV Heavy. Boeing also is evaluating auxiliary launch concepts for Delta IV. Over the past 40 years, Delta has significantly increased overall capability while reducing launch costs. Delta IV will continue this tradition with dramatic cost reductions while providing a more robust expendable launch vehicle.

On the comet rendezvous exploration and its potential extension via solar electric propulsion (Comet Nucleus Sample and Return Mission)

The Comet Nucleus Sample and Return (CNSR) mission has been of great interest for many years. The mission objective is very clear and it provides the clues to the solar system and the origin of life. The paper presents the feasible CNSR scenario taking the advantage of Solar Electric Propulsion to the comet Wilson-Harrington (WH). The plan is found feasible, and the launch window is in 2005 and the spacecraft can return the sample back to the Earth in only 8 years. Another window that is open in 2009–2010 is also provided. They are making use of the ion engine propulsion accumulating the orbital energy at Earth distance, followed by the Earth gravity assist. This is the unique point enabling even the smaller spacecraft to accomplish this high energy mission. This strategy accelerates the spacecraft orbital energy up to 36– 40 km 2/ s 2 in C3 that is adequate for the flight to the comet WH. The idea shown here also includes: (1) the rendezvous point to be at Earth distance so that the electric propulsion is utilized at maximum and (2) the target comet to be extinct so that the sampling and in situ measurement scenario robust. The trajectory synthesis developed here and tactical target selection make this impressive mission to be envisioned even by smaller missions. This is not an endorsed mission but ISAS may consider collaborating with the foreign space organization on the mission.