Well, I'm still at page 50 of this space epic, but I really wanted to congratulate the writers. This is indeed cool stuff!
Hmm. It sounds like for the Planetary Science Teams could be dealing with either a massive defeat, or a substantial boost to their studies. But more likely what I'm seeing right now - given the budgetary focus on Artemis - is something between the two points.
The methane out-gassing on Mars, how long can methane be sustained in the current Mars Atmosphere before it's broken into Carbon and Hydrogen? I do know though, that the Martian Interior does still have some heat left in it, not enough for a liquid core, but its still there.
The methane out-gassing on Mars, how long can methane be sustained in the current Mars Atmosphere before it's broken into Carbon and Hydrogen? I do know though, that the Martian Interior does still have some heat left in it, not enough for a liquid core, but its still there.
According to this Planetary Society post, about 300 years. That's short enough that, in geological terms, there'd have to be an ongoing source. In essence, they've done exactly what Mars Express did IOTL about the same time (that part was a coincidence, discovering methane was not, though obviously it was there to be found and they just needed the right instruments in orbit).
I wrote to Winchell Chung about our recent updates on the wiki. Hopefully he'll update the two links he has to the pages over at Atomic Rockets.
Update: Got an e-mail from him yesterday, and he's already updated it. What a prompt guy !
Morning all. With the historic discovery of a potential bio-indicator on the Red Planet, this week let's take a look at the MACO probe.
Part IV, Post 9: Commercial reactions to StarLaunch Thunderbolt, including Northrop TransOrbital Services
Hello, everyone! So, when we left off, the planetary science teams were dealing with the fallout from the absence of a new Pioneer selection in 2006 and the struggle for approval of larger missions. However, they aren't the only ones facing troubles, as the commercial market wrestles with the implications of the StarLaunch Thunderbolt...
Eyes Turned Skyward, Part IV: Post #9
The arrival and initial success of Paul Allen’s StarLaunch Thunderbolt on the commercial market was a dramatic shock to an industry already in turmoil. The traditional leader in the commercial launch market, Lockheed, had recently retired its Titan family, having cited rising safety, processing, and propellent costs related to its use of hypergolic propellants as the cause. This left only the ex-McDonnell-Douglas Delta 5000 for future customers. Since the Delta was capable of only 4 tons to a geosynchronous transfer orbit (GTO) instead of the more than 6 tons of Titan and the even greater payload of new competition like the Europa 5 and the Russian Neva and Vulkan, Lockheed-McDonnell had been robbed of its capacity to handle the largest of geosynchronous communications satellites--the 6-ton “super”-class busses that Lockheed themselves had introduced. To fill this capability gap, Lockheed had initially planned on a combination of a commercial derivative of the Starclipper program they had worked on with NASA and the new generation of electrically-propelled satellite busses that was emerging from their satellite plants. Reusable launchers, after all, could dramatically drop their costs, while the transition from chemical fuels to electric ion engines for initial circularization and station-keeping in geosynch could allow the equivalent of the old “super”-class bus to fit into the 4-ton capacity of the Delta 5000 as an interim solution.
While the interim “band-aid” of the electric satellite bus had some success with sales to customers impressed by the cheaper launch costs available by launching a bus of similar capacity to an old 6-ton “super” on a 4-ton LV keeping Delta 5000 sales steady, Lockheed’s long-term plan to transition to reusable SSTO launchers was crippled when the X-33 program intended to demonstrate the feasibility of Starclipper hardware for eventual SSTO use instead did everything but. Despite a number of successful test flights, the most profound takeaway for Lockheed and NASA engineers was that SSTO was not possible with current technologies. Instead, they had doubled down on the trying to compete with the same conventional rockets as their competition, further trimming Delta 5000’s operational costs, and offering discounts to the major operators on joint flights to bolster immediate sales. However, the dramatic entrance of StarLaunch into the market would strike like a Thunderbolt.
Thunderbolt’s entrance into the world launch market, even in its “L1” semi-reusable form, was disruptive enough to make the already-existent uncertainties in the market seem inconsequential by comparison. While Europa 5, Neva, and Vulkan clamored for a shot at the long-term Lockheed dominance of the market, their prices were not far below those of the Delta 5000 on which Lockheed had standardized. Thunderbolt’s costs, however, were in a whole different class--less than half as costly per-kg as its competitors, and with the promise of further reductions should they be able to fund development of the L2 reusable upper stage. While the putative competition had been able to ignore them as over-ambitious and likely doomed to failure--as they had ignored other ambitious startups in the past--once StarLaunch began regular full-up flight simulations on the path to commercial operations that was no longer an option. Within a year of its first test flights, reservations of Thunderbolt had begun to nearly monopolize the “small” launcher class, leaving sales of competitors like ALS Carrack and Europaspace’s Europa 2-HE to slump. While Europaspace could fall back for the moment on Europa 5 launches, the loss in business for ALS’ sole product was devastating. Dramatic action on their part would be necessary to maintain profitability.
However, while Thunderbolt--even without fully reaching operational status--was already revolutionizing the “small” launch market, the traditional geosynchronous payload market was insulated from immediate effects. With a payload of roughly a ton to conventional GTO, Thunderbolt was simply unable to loft the kinds of multi-ton satellites commercial operators were purchasing for launch. While the competitive advantages--and feasibility--of reuse, even partial reuse, were being strongly demonstrated by StarLaunch, it would take a larger vehicle to be able to actually compete for traditional commercial satellite launches--and that need gave a window for those traditional launch companies to react before they faced the same threats ALS was already struggling with, though many would be unwilling or otherwise unable to make use of it.
In the eyes of European advocates of reusability, StarLaunch’s success with Thunderbolt was a validation of much of what they had been trying to prove to Europe’s space ministers for almost a decade--reusability was a key capability, but one that was technically feasible, and crucial to pursue. However, the vertical-takeoff and conventional-rocket-propelled Thunderbolt was also such a radical departure from the air-breathing booster and winged upper stage that had been pursued under Sanger/Horus that some of the more traditional advocates in France and the UK were now saying that the German-lead efforts had “wasted” Europe’s chance to beat the American state of the art by aiming for using immature, beyond-state-of-the-art technologies instead of “good-enough” and more mature systems. Both charges were roughly true, but the conflicting views of past efforts would shape the debate over how to react to Thunderbolt. A multitude of paths were open for the development of such a European RLV, and the advocates of each within Europe’s space industry and spaceflight agencies broke down largely on the grounds of where individuals had stood before Thunderbolt’s development.
The first option was an adaption of the work already performed for Sanger/Horus, which had successfully validated many of the structural and aerodynamic models for a winged, reusable hydrogen spaceplane/upper stage before being stymied by the continued failures of the engine demonstrations necessary to validate the first stage air-breathing design. Several factions within ESA now saw the chance to adopt a model with a conventional rocket-powered, Horus-derived booster as well as the existing Horus upper stage. Studies of both vertical and horizontal liftoff variants began to circulate within the existing Sanger/Horus offices in Germany. However, this faced opposition from a second option, put forward by French designers aligned with Europaspace who had been involved with Europa 5’s development not long past. In this proposal, a similar (though smaller) Horus upper stage would be paired with reusable boosters to be directly based on the Europa 5 core stage, potentially fitted with winglets and jet engines for a flyback to launch site, or with landing legs for a rocket-propelled boostback maneuver like that carried out by StarLaunch Thunderbolt. This approach, they claimed, would offer cheaper development and a closer continuity with the existing (and reliable) history of European launcher development, while also continuing the flexibility of capacity Europa had enjoyed since Europa 4.
Yet a third, though much smaller group, was led by Alan Bond, a member of the British development team assisting with Sanger’s (unsuccessful) turborocket first stage. Before the final selection of the Sanger design for the reusability testing, Bond and his team had been involved with the early definitional studies of the German reusability program, involving their own concept, a multi-mode engine capable of transitioning from a pre-cooled air/LH2 mode in atmosphere to a fully-internal “conventional” LH2/LOX rocket engine. This engine, combined with suggested refinements to traditional structural techniques, was enough in Bond’s calculations to enable a fully-reusable single-stage vehicle. However, the engine and the vehicle to which it were attached attracted little attention after the repeated failures of the (unrelated) Hypersonic Engine Demonstrators, and Bond eventually resigned from Rolls-Royce along with some of the rest of his team to pursue his technologies independently.
In the end, neither of the two “mainstream” European proposals was able to capture solid support either. Ironically, in debates over which mode of reuse was preferable, both major factions pointed to the success of Thunderbolt as evidence to support their respective arguments: the largely-German “high-tech” team traced their suggestion of adapting Horus as both a booster and orbiter to StarLaunch’s planned similarity between Thunderbolt L1 and the Thunderbolt L2 upper stage, while the largely-French reusability plan from Europaspace pointed instead at the more “traditional” layout of the Thunderbolt vehicle. British, Italian, and other national space ministries similarly fractured, and the lack of immediate pressure from Thunderbolt on Europa 5’s launch market left just enough breathing space for a complete stalemate in the short term. While Europa 5 remained focused on competing with Neva and Vulkan for who could capture Titan’s old market share, the question of European reuse remained effectively tabled--while it was generally agreed that having such a plan was critical, no one could agree on which path forward was to be followed. It would take more than Thunderbolt’s early commercial flights to shock the complex political system of European spaceflight out of its stalemate.
Europe’s program wasn’t the only one unable to immediately react to StarLaunch’s dramatic emergence. The Russian program depended heavily for cashflow on the portions of the comsat market it had captured following the retirement of Titan or wrestled from Europa and Delta 5000 since, a success largely contingent on the fact that development of Vulkan and Neva rockets had been fully paid previously. Thus, it was a matter of merely matching their lower labor costs against the higher costs of American or European lunch suppliers, a strategy which had resulted in some success. However, investing billions of dollars into development of their own reusable vehicle was no more an option for the Russians in the new millenium than it had been for Chelomei, the last of the Chief Designers, in the waning days of the USSR more than a decade earlier. Despite a willingness to seize the initiative and public statements about an intention to investigate their own RLV, public relations was essentially all these efforts were--the development money for a program intended to move beyond simple mockups simply wasn’t an option even with the program continuing to recover slowly for its nadir in the early 90s.
The Japanese had never made an extensive play to enter the satellite launch market, focusing their N and H-series rockets more on internal needs, such as the JLV (Japanese Logistics Vehicle) which was to be their new contribution to Freedom. This freed them from much of the urgency of the need to respond to Thunderbolt, and meant that much of the Japanese industry cultivated a “wait-and-see” approach to the development of reusable launch vehicles--while a program could potentially be afforded, unlike Russia, it would still monopolize much of the spaceflight budget--a singular risk that the Japanese program managers were hesitant to take. Instead, their official stance was that the HOPE development work they had conducted in the 90s would allow them to easily catch up any lead other RLV operators might gain, and thus they would allow StarLaunch and any followers to probe the market before committing to catch up to whatever approach worked best. If the Thunderbolt and followers failed, there might be room in an upset commercial market to finally make a play with their own H-I rocket. For the moment, Japan would watch and wait.
This “wait-and-see” attitude wasn’t unique to the Japanese, however. Lockheed-McDonnell, like many others in the geosynchronous launch market, had been observing Thunderbolt’s emergence with a mix of trepidation and Japan’s cool calculation. However, unlike Japan, whose HOPE project was quite literally just getting off the ground, Lockheed had the benefit of their X-33 flight test program to draw on with regard to reusable vehicle design and operations. While the X-33’s performance shortcomings had ultimately killed the original plan to build an SSTO Starclipper, as early as 1997 Lockheed’s engineers had begun to consider adaptations of the design into part of a multi-stage vehicle, much like Thunderbolt. Moving cautiously to avoid committing the firm to a government-sized development program, Lockheed began to reconsider these earlier studies, analyzing whether they could be adapted to fit with a two-stage reusable vehicle. While they were not yet investing in a new vehicle, they were positioning themselves to be ready to pull the trigger if necessary, with key factor in deciding if it made sense to try to “leap-frog” Thunderbolt in the race to seize the lost Titan market share being the scale of the vehicles involved. While Thunderbolt was seeing success in the smallsat market, its performance to the geosynchronous transfer orbits critical in the commercial world was too little to liftany standard comsat bus. To serve this market, either the prospective two-stage Starclipper would have to be several times the size, or Lockheed would have to find another way to serve the market which had nurtured commercial spaceflight.
The most dramatic result of Thunderbolt’s introduction, though, didn’t come from a competing launch provider at all. Instead, Northrop (manufacturer of the Centaur upper stage in use on almost all US launchers, including Thunderbolt itself) sought to offer a service to create synergy with StarLaunch--Northrop TransOrbital Services. The reason that Europe and others had been able to delay their reactions to Thunderbolt largely stemmed from its inability to carry large commercial satellites to GTO. Thus, its low launch costs were not an option for most commercial satellite users. It had been assumed that it would take a second-generation LV to loft large enough payloads to compete directly with other existing offerings. Northrop TransOrbital would change that calculation entirely. The proposal drew on the already well-advanced work that Northrop was doing for NASA on a Centaur-based cryogenic depot, which included work on the technologies needed for extended sub-1%-per-day boiloff, studies of orbital fuel transfer technologies and planned ground-based trials of such systems, and some study of RL-10 adaption for relight many times in space over extended periods--all technologies NASA was aiming at a potential Pegasus-base full-size depot which could aid in cutting launch costs for Artemis missions.
However, Northrop saw a chance to adapt this already-in-progress work at minimal cost to a unique niche: orbital transfer of commercial payloads, hence the “TransOrbital” program name. The business model involved using an operational Centaur-Depot (based directly on the unit already in work for a NASA launch in 2006) to tend to a Centaur-Tug drawing on many of the same technologies but fitted with the extended-life RL-10 variant also under development. Purchasing fuel from commercial suppliers (StarLaunch had an obvious lead, but Northrop was willing to entertain bids for supply from any company which could match price) would allow the depot to be refilled, while the tug could make rendezvous with geosynch-bound payloads in LEO, dock using a CADS port, and conduct a burn to carry them to GTO or (for a higher cost) directly to GEO. The tug could then undo the job, and return itself to LEO to refill at the depot. The 6-ton payload of Thunderbolt was already enough capacity for most of the largest comsats (ironically, partially thanks to Lockheed’s advocacy of their “super efficiency” electric propulsion bus, which let a 4-ton bus carry the same power and transmitters as an all-chemical 6-ton “super” bus), and the Centaur-based tug and depot could easily deliver such a payload to high orbits. Northrop had worked with StarLaunch (and particularly Don Hunt) to develop the concept ahead of its announcement, which set off waves at a joint press conference in late 2003 as Northrop announced the service, and also immediately announced contracting StarLaunch for fuel delivery services, while StarLaunch in turn announced that they were pleased to partner to offer the TransOrbital transfer to customers as soon as 2009--an aggressive schedule Northrop planned to be able to meet thanks to more than two years of existing work on the demonstration depot for NASA.
With the debut of TransOrbital, the shock of Thunderbolt’s original entry into the market was amplified. If TransOrbital functioned as planned, Thunderbolt would become direct competition for the prized geosynchronous comsat market, turning it from a curiosity heralding future competition into a potentially mortal threat. As ALS’ growing losses on their income statements and dwindling launch manifests proved, trying to compete against a reusable launcher with a conventional expendable was an uphill battle. Even though even the conceptual function of depots had yet to be tested in flight and wouldn’t be for more than two years, the potential risk was too large to totally ignore. Many within the commercial spaceflight arena had to once more re-evaluate their plans--and their schedules--in the wake of TransOrbital’s announcement.
Eyes Turned Skyward, Part IV: Post #9
The arrival and initial success of Paul Allen’s StarLaunch Thunderbolt on the commercial market was a dramatic shock to an industry already in turmoil. The traditional leader in the commercial launch market, Lockheed, had recently retired its Titan family, having cited rising safety, processing, and propellent costs related to its use of hypergolic propellants as the cause. This left only the ex-McDonnell-Douglas Delta 5000 for future customers. Since the Delta was capable of only 4 tons to a geosynchronous transfer orbit (GTO) instead of the more than 6 tons of Titan and the even greater payload of new competition like the Europa 5 and the Russian Neva and Vulkan, Lockheed-McDonnell had been robbed of its capacity to handle the largest of geosynchronous communications satellites--the 6-ton “super”-class busses that Lockheed themselves had introduced. To fill this capability gap, Lockheed had initially planned on a combination of a commercial derivative of the Starclipper program they had worked on with NASA and the new generation of electrically-propelled satellite busses that was emerging from their satellite plants. Reusable launchers, after all, could dramatically drop their costs, while the transition from chemical fuels to electric ion engines for initial circularization and station-keeping in geosynch could allow the equivalent of the old “super”-class bus to fit into the 4-ton capacity of the Delta 5000 as an interim solution.
While the interim “band-aid” of the electric satellite bus had some success with sales to customers impressed by the cheaper launch costs available by launching a bus of similar capacity to an old 6-ton “super” on a 4-ton LV keeping Delta 5000 sales steady, Lockheed’s long-term plan to transition to reusable SSTO launchers was crippled when the X-33 program intended to demonstrate the feasibility of Starclipper hardware for eventual SSTO use instead did everything but. Despite a number of successful test flights, the most profound takeaway for Lockheed and NASA engineers was that SSTO was not possible with current technologies. Instead, they had doubled down on the trying to compete with the same conventional rockets as their competition, further trimming Delta 5000’s operational costs, and offering discounts to the major operators on joint flights to bolster immediate sales. However, the dramatic entrance of StarLaunch into the market would strike like a Thunderbolt.
Thunderbolt’s entrance into the world launch market, even in its “L1” semi-reusable form, was disruptive enough to make the already-existent uncertainties in the market seem inconsequential by comparison. While Europa 5, Neva, and Vulkan clamored for a shot at the long-term Lockheed dominance of the market, their prices were not far below those of the Delta 5000 on which Lockheed had standardized. Thunderbolt’s costs, however, were in a whole different class--less than half as costly per-kg as its competitors, and with the promise of further reductions should they be able to fund development of the L2 reusable upper stage. While the putative competition had been able to ignore them as over-ambitious and likely doomed to failure--as they had ignored other ambitious startups in the past--once StarLaunch began regular full-up flight simulations on the path to commercial operations that was no longer an option. Within a year of its first test flights, reservations of Thunderbolt had begun to nearly monopolize the “small” launcher class, leaving sales of competitors like ALS Carrack and Europaspace’s Europa 2-HE to slump. While Europaspace could fall back for the moment on Europa 5 launches, the loss in business for ALS’ sole product was devastating. Dramatic action on their part would be necessary to maintain profitability.
However, while Thunderbolt--even without fully reaching operational status--was already revolutionizing the “small” launch market, the traditional geosynchronous payload market was insulated from immediate effects. With a payload of roughly a ton to conventional GTO, Thunderbolt was simply unable to loft the kinds of multi-ton satellites commercial operators were purchasing for launch. While the competitive advantages--and feasibility--of reuse, even partial reuse, were being strongly demonstrated by StarLaunch, it would take a larger vehicle to be able to actually compete for traditional commercial satellite launches--and that need gave a window for those traditional launch companies to react before they faced the same threats ALS was already struggling with, though many would be unwilling or otherwise unable to make use of it.
In the eyes of European advocates of reusability, StarLaunch’s success with Thunderbolt was a validation of much of what they had been trying to prove to Europe’s space ministers for almost a decade--reusability was a key capability, but one that was technically feasible, and crucial to pursue. However, the vertical-takeoff and conventional-rocket-propelled Thunderbolt was also such a radical departure from the air-breathing booster and winged upper stage that had been pursued under Sanger/Horus that some of the more traditional advocates in France and the UK were now saying that the German-lead efforts had “wasted” Europe’s chance to beat the American state of the art by aiming for using immature, beyond-state-of-the-art technologies instead of “good-enough” and more mature systems. Both charges were roughly true, but the conflicting views of past efforts would shape the debate over how to react to Thunderbolt. A multitude of paths were open for the development of such a European RLV, and the advocates of each within Europe’s space industry and spaceflight agencies broke down largely on the grounds of where individuals had stood before Thunderbolt’s development.
The first option was an adaption of the work already performed for Sanger/Horus, which had successfully validated many of the structural and aerodynamic models for a winged, reusable hydrogen spaceplane/upper stage before being stymied by the continued failures of the engine demonstrations necessary to validate the first stage air-breathing design. Several factions within ESA now saw the chance to adopt a model with a conventional rocket-powered, Horus-derived booster as well as the existing Horus upper stage. Studies of both vertical and horizontal liftoff variants began to circulate within the existing Sanger/Horus offices in Germany. However, this faced opposition from a second option, put forward by French designers aligned with Europaspace who had been involved with Europa 5’s development not long past. In this proposal, a similar (though smaller) Horus upper stage would be paired with reusable boosters to be directly based on the Europa 5 core stage, potentially fitted with winglets and jet engines for a flyback to launch site, or with landing legs for a rocket-propelled boostback maneuver like that carried out by StarLaunch Thunderbolt. This approach, they claimed, would offer cheaper development and a closer continuity with the existing (and reliable) history of European launcher development, while also continuing the flexibility of capacity Europa had enjoyed since Europa 4.
Yet a third, though much smaller group, was led by Alan Bond, a member of the British development team assisting with Sanger’s (unsuccessful) turborocket first stage. Before the final selection of the Sanger design for the reusability testing, Bond and his team had been involved with the early definitional studies of the German reusability program, involving their own concept, a multi-mode engine capable of transitioning from a pre-cooled air/LH2 mode in atmosphere to a fully-internal “conventional” LH2/LOX rocket engine. This engine, combined with suggested refinements to traditional structural techniques, was enough in Bond’s calculations to enable a fully-reusable single-stage vehicle. However, the engine and the vehicle to which it were attached attracted little attention after the repeated failures of the (unrelated) Hypersonic Engine Demonstrators, and Bond eventually resigned from Rolls-Royce along with some of the rest of his team to pursue his technologies independently.
In the end, neither of the two “mainstream” European proposals was able to capture solid support either. Ironically, in debates over which mode of reuse was preferable, both major factions pointed to the success of Thunderbolt as evidence to support their respective arguments: the largely-German “high-tech” team traced their suggestion of adapting Horus as both a booster and orbiter to StarLaunch’s planned similarity between Thunderbolt L1 and the Thunderbolt L2 upper stage, while the largely-French reusability plan from Europaspace pointed instead at the more “traditional” layout of the Thunderbolt vehicle. British, Italian, and other national space ministries similarly fractured, and the lack of immediate pressure from Thunderbolt on Europa 5’s launch market left just enough breathing space for a complete stalemate in the short term. While Europa 5 remained focused on competing with Neva and Vulkan for who could capture Titan’s old market share, the question of European reuse remained effectively tabled--while it was generally agreed that having such a plan was critical, no one could agree on which path forward was to be followed. It would take more than Thunderbolt’s early commercial flights to shock the complex political system of European spaceflight out of its stalemate.
Europe’s program wasn’t the only one unable to immediately react to StarLaunch’s dramatic emergence. The Russian program depended heavily for cashflow on the portions of the comsat market it had captured following the retirement of Titan or wrestled from Europa and Delta 5000 since, a success largely contingent on the fact that development of Vulkan and Neva rockets had been fully paid previously. Thus, it was a matter of merely matching their lower labor costs against the higher costs of American or European lunch suppliers, a strategy which had resulted in some success. However, investing billions of dollars into development of their own reusable vehicle was no more an option for the Russians in the new millenium than it had been for Chelomei, the last of the Chief Designers, in the waning days of the USSR more than a decade earlier. Despite a willingness to seize the initiative and public statements about an intention to investigate their own RLV, public relations was essentially all these efforts were--the development money for a program intended to move beyond simple mockups simply wasn’t an option even with the program continuing to recover slowly for its nadir in the early 90s.
The Japanese had never made an extensive play to enter the satellite launch market, focusing their N and H-series rockets more on internal needs, such as the JLV (Japanese Logistics Vehicle) which was to be their new contribution to Freedom. This freed them from much of the urgency of the need to respond to Thunderbolt, and meant that much of the Japanese industry cultivated a “wait-and-see” approach to the development of reusable launch vehicles--while a program could potentially be afforded, unlike Russia, it would still monopolize much of the spaceflight budget--a singular risk that the Japanese program managers were hesitant to take. Instead, their official stance was that the HOPE development work they had conducted in the 90s would allow them to easily catch up any lead other RLV operators might gain, and thus they would allow StarLaunch and any followers to probe the market before committing to catch up to whatever approach worked best. If the Thunderbolt and followers failed, there might be room in an upset commercial market to finally make a play with their own H-I rocket. For the moment, Japan would watch and wait.
This “wait-and-see” attitude wasn’t unique to the Japanese, however. Lockheed-McDonnell, like many others in the geosynchronous launch market, had been observing Thunderbolt’s emergence with a mix of trepidation and Japan’s cool calculation. However, unlike Japan, whose HOPE project was quite literally just getting off the ground, Lockheed had the benefit of their X-33 flight test program to draw on with regard to reusable vehicle design and operations. While the X-33’s performance shortcomings had ultimately killed the original plan to build an SSTO Starclipper, as early as 1997 Lockheed’s engineers had begun to consider adaptations of the design into part of a multi-stage vehicle, much like Thunderbolt. Moving cautiously to avoid committing the firm to a government-sized development program, Lockheed began to reconsider these earlier studies, analyzing whether they could be adapted to fit with a two-stage reusable vehicle. While they were not yet investing in a new vehicle, they were positioning themselves to be ready to pull the trigger if necessary, with key factor in deciding if it made sense to try to “leap-frog” Thunderbolt in the race to seize the lost Titan market share being the scale of the vehicles involved. While Thunderbolt was seeing success in the smallsat market, its performance to the geosynchronous transfer orbits critical in the commercial world was too little to liftany standard comsat bus. To serve this market, either the prospective two-stage Starclipper would have to be several times the size, or Lockheed would have to find another way to serve the market which had nurtured commercial spaceflight.
The most dramatic result of Thunderbolt’s introduction, though, didn’t come from a competing launch provider at all. Instead, Northrop (manufacturer of the Centaur upper stage in use on almost all US launchers, including Thunderbolt itself) sought to offer a service to create synergy with StarLaunch--Northrop TransOrbital Services. The reason that Europe and others had been able to delay their reactions to Thunderbolt largely stemmed from its inability to carry large commercial satellites to GTO. Thus, its low launch costs were not an option for most commercial satellite users. It had been assumed that it would take a second-generation LV to loft large enough payloads to compete directly with other existing offerings. Northrop TransOrbital would change that calculation entirely. The proposal drew on the already well-advanced work that Northrop was doing for NASA on a Centaur-based cryogenic depot, which included work on the technologies needed for extended sub-1%-per-day boiloff, studies of orbital fuel transfer technologies and planned ground-based trials of such systems, and some study of RL-10 adaption for relight many times in space over extended periods--all technologies NASA was aiming at a potential Pegasus-base full-size depot which could aid in cutting launch costs for Artemis missions.
However, Northrop saw a chance to adapt this already-in-progress work at minimal cost to a unique niche: orbital transfer of commercial payloads, hence the “TransOrbital” program name. The business model involved using an operational Centaur-Depot (based directly on the unit already in work for a NASA launch in 2006) to tend to a Centaur-Tug drawing on many of the same technologies but fitted with the extended-life RL-10 variant also under development. Purchasing fuel from commercial suppliers (StarLaunch had an obvious lead, but Northrop was willing to entertain bids for supply from any company which could match price) would allow the depot to be refilled, while the tug could make rendezvous with geosynch-bound payloads in LEO, dock using a CADS port, and conduct a burn to carry them to GTO or (for a higher cost) directly to GEO. The tug could then undo the job, and return itself to LEO to refill at the depot. The 6-ton payload of Thunderbolt was already enough capacity for most of the largest comsats (ironically, partially thanks to Lockheed’s advocacy of their “super efficiency” electric propulsion bus, which let a 4-ton bus carry the same power and transmitters as an all-chemical 6-ton “super” bus), and the Centaur-based tug and depot could easily deliver such a payload to high orbits. Northrop had worked with StarLaunch (and particularly Don Hunt) to develop the concept ahead of its announcement, which set off waves at a joint press conference in late 2003 as Northrop announced the service, and also immediately announced contracting StarLaunch for fuel delivery services, while StarLaunch in turn announced that they were pleased to partner to offer the TransOrbital transfer to customers as soon as 2009--an aggressive schedule Northrop planned to be able to meet thanks to more than two years of existing work on the demonstration depot for NASA.
With the debut of TransOrbital, the shock of Thunderbolt’s original entry into the market was amplified. If TransOrbital functioned as planned, Thunderbolt would become direct competition for the prized geosynchronous comsat market, turning it from a curiosity heralding future competition into a potentially mortal threat. As ALS’ growing losses on their income statements and dwindling launch manifests proved, trying to compete against a reusable launcher with a conventional expendable was an uphill battle. Even though even the conceptual function of depots had yet to be tested in flight and wouldn’t be for more than two years, the potential risk was too large to totally ignore. Many within the commercial spaceflight arena had to once more re-evaluate their plans--and their schedules--in the wake of TransOrbital’s announcement.
a surprising chapter
first on European
seems they got into same labyrinth of studies like OTL
for Ariane 5 and 6 they study allot RLV version
like 1982 CNES using a huge shuttle like craft with Lox/kero and upper stage inside huge bomb bay in rocket
in same time Germany MBB Offenbrunn very busy a group under W.Kleinau
they proposed Ariane Reusable AR-X
based on H45 Tanks cluster and HM60 Engine but inside a Aerodynamic hull with Heatshield
AR X-1 Demonstrator
Stage one
4xH45 and 9xHM60 (8 for Launch, 1 for Landing)
Stage two (inside the payload shroud)
1xH45 with 1xHM60
the finale AR
has a Second stage build from 2XH45 Tanks piece and 5xHM60 engine
(4 for launch, 1 for landing) inside Aerodynamic hull with Heatshield
the Frist stage lands 1500 km away in Atlantic, while second stage made several orbit and land near Space Port in in Atlantic.
Now that could be adapted on Europa 5 by putting the stages inside a Aerodynamic hull with Heatshield.
my point of critic on Lockheed-McDonnell
that they have to abandon the Titan is obvious: Toxic, complex and hell of cost to launch it and mend the launch pad after each takeoff.
now McDonnell got in TL the Douglas company,
so why they not use the Saturn S-IVC stage with UA1205 boosters or use reusable Booster ?!
i mean with production for Saturn rockets, the S-IVC stage must be low-cost
first on European
seems they got into same labyrinth of studies like OTL
for Ariane 5 and 6 they study allot RLV version
like 1982 CNES using a huge shuttle like craft with Lox/kero and upper stage inside huge bomb bay in rocket
in same time Germany MBB Offenbrunn very busy a group under W.Kleinau
they proposed Ariane Reusable AR-X
based on H45 Tanks cluster and HM60 Engine but inside a Aerodynamic hull with Heatshield
AR X-1 Demonstrator
Stage one
4xH45 and 9xHM60 (8 for Launch, 1 for Landing)
Stage two (inside the payload shroud)
1xH45 with 1xHM60
the finale AR
has a Second stage build from 2XH45 Tanks piece and 5xHM60 engine
(4 for launch, 1 for landing) inside Aerodynamic hull with Heatshield
the Frist stage lands 1500 km away in Atlantic, while second stage made several orbit and land near Space Port in in Atlantic.
Now that could be adapted on Europa 5 by putting the stages inside a Aerodynamic hull with Heatshield.
my point of critic on Lockheed-McDonnell
that they have to abandon the Titan is obvious: Toxic, complex and hell of cost to launch it and mend the launch pad after each takeoff.
now McDonnell got in TL the Douglas company,
so why they not use the Saturn S-IVC stage with UA1205 boosters or use reusable Booster ?!
i mean with production for Saturn rockets, the S-IVC stage must be low-cost
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Ouch. Thunderbolt's really causing quite the stir in the established Launch Providers with their far lower $:Kg costs, and if that Propellant Depot works it'll go right up to Eleven.
For for the traditional providers responses? With NASA, it's like herding cats (or so I've heard), but with ESA, said cats are running on a mix of Caffeine & Taurine, making it that much harder. The contractors aren't having it easy either, with only the heavy payloads to GTO offering a lifeline for their businesses which is looking decidedly fragile at the moment.
This, is gonna have some ripples down the line.
For for the traditional providers responses? With NASA, it's like herding cats (or so I've heard), but with ESA, said cats are running on a mix of Caffeine & Taurine, making it that much harder. The contractors aren't having it easy either, with only the heavy payloads to GTO offering a lifeline for their businesses which is looking decidedly fragile at the moment.
This, is gonna have some ripples down the line.
Kindof like the SASSTO? Only using solids to provide initial lift, rather than using slush hydrogen?
Depending on the number of solids, the payload could very well be in the 4 to 6 tonnes range for LEO, where they might use TransOrbital for delivery to GEO.
Thanks, we try.
To avoid serious weight gain, you'd need a lot of redesign of the tanks to integrate them structurally, and that's expensive--likely not much less expensive than a clean-sheet flyback first stage. heavier than a fully-redesigned , and
Well, to start with, they have Delta 5000, and with that a certain assured DoD and NASA market. That's one reason not to change radically but instead to try and sell their electric 4-ton busses launched on Delta--saves a big development project that may not be nesecary.
However, if they were to pull the trigger on that massive effort, there's two ways to take that the way you wrote it, and I'll address both. The first is to strap solids around the S-IVC, and essentially make sort of an Ariane 5-type rocket. For that, the Delta 5000 CCM-46 is probably a better choice, with variable numbers. You end up with slightly better numbers than the equivalent Delta 5000, but not likely a big enough jump to get back to the 6-ton bus market, and the costs will be comparable. Stack a Centaur on top, and you're now up into the 10-ton-to-GTO range, but same cost problems--you'll be competitive with Europa 5 and Neva, but not with Thunderbolt/TransOrbital. And, of course, you'll have to radically overhaul the S-IVC to turn it into a first stage core with solid attachment points, not a second stage (kind of the inverse of the one ATK was talking about pulling with Ariane 5's stage for Liberty before that idea died the early death it so well deserved).
The second way is to essentially recreate the INT-5 or INT-16, make the S-IV still an upper stage with some kind of new first stage. In that case, you probably favor the smaller S-IVB rather than the stretched S-IVC. In any case, it tends to drive the size of the LV to the large end. It wouldn't surprise me if LockMac had considered something of the like if the period, if one dug through TTL's equivalent of Astronautix, but the problem is that what you end up with if you don't go reusable with the booster is essentially an off-brand Saturn M02, only without Boeing's assured minimum launch rate from DoD and NASA flights. If you do try reuse...well, you need a stage in the ballpark of several hundred tons, which is a pretty big vehicle to invest in developing from scratch. Until X-33 shows them SSTO isn't feasible, that's Lockheed's preferred path, then until Thunderbolt starts seeing solid commercial success a couple years later in 2003, they're not looking as hard at reuse. Then, before they can really react to that, TransOrbital comes along, and maybe you don't need 16-odd tons IMLEO and thus hundreds of tons GLOW to get 6 tons to GTO and that starts looking at Starclipper-based TSTO and such, but it's a gamble to start work on it before TransOrbital proes it can work...basically, LockMac's behind the power curve on reuse.
Well, as noted, LockMac would probably favor a Starclipper-derived solution to a S-IVB modification--they've been working on those kinds of things themselves quite a bit more recently. It's a question of what size vehicle they need, what fraction they decide they're willing to invest development to make reusable, and the like.
Um, I'm confused--isn't the whole Saturn Multibody series, all its liquid stages anyway (not so sure about the solid boosters) owned and made entirely in-house by Boeing? What does McDonnell or Douglas have to do with it?
Why Boeing doesn't get into the commercial launcher biz--is something we've discussed here before, IIRC:
1) Multibodies are huge by commercial launcher standards;
2) Boeing is comfortable with DoD and NASA orders, and since the former and even to some extent the latter is jealous of foreign operations, doesn't want to risk its plum deal with the US Government by selling to other operators (even purely US corporations--after all in the era of globalization, what company is pure US anyway?)
3) The Multibodies are designed to be launched from NASA's and the Air Force's particular facilities, at Canaveral and Vandenberg, and these sites tend to be kept busy by NASA and Air Force missions. It is more in Boeing's interest to keep these two customers busy ordering their own scheduled launches and if a lull were to occur it could be awkward arranging to use these two USG sites for commercial launches. Were they to find a market for their big rockets in defiance of point 1 and risk USG displeasure (or get a clear dispensation) to address point 2, they might still need to develop a third launch site, one with facilities, if not quite as elaborate as the gargantuan VAB at Canaveral, whatever monster hangar they have at Vandenberg. Building such facilities from scratch anywhere then staffing them is going to be a big budget item, not to be considered for a marginal market share. I did once have enthusiastic hopes of there being a branch site at Kourou, but that runs straight into issues of international competition, on both sides.
So Boeing is out of the picture and I am very confused by all this talk of other companies horning in on Saturn derivatives at this late date.
------
Alan Bond, again at this late date (if I have not forgotten some earlier mention of him being involved with Horus) is a bit of an Easter Egg, but I fear possibly one who will never be found before going rotten. OTL I'm a big fan of Skylon, but that's partly from being a romantic who roots for underdogs. Mind I'm sold on the idea that his singular approach to SSTO can work, and at a stroke come much closer to the dream of an airplane-like mode of operations (land, refuel, load in the new payload whilst doing cursory routine diagnostics on the engines and other systems that generally confirm, yep, they are all ready to go again as they are with minor tweaking and oiling, then take off and go, deliver the next cargo and possibly pick up a small down cargo, and so on for tens of sorties). Also it looks as elegant as all hell, and its concept of operations is also elegant. But REI is an underdog because it doesn't have the backing of any of the big players, only just now attracting serious investment and scrutiny from HMG and the ESA.
So that can give me hope--TTL's Bond et al are no more (if no less) in the wilderness than they were OTL at this date. Perhaps. We don't know if he's where OTL Bond was in the same year, behind or just maybe perhaps a bit ahead. But as per OTL he's a long shot coming out of left field.
---
I note though that there are really not one but two schemes to achieve transfer to GEO by secondary means. Lockheed need not panic nor let Northrop's "TransOrbital" reusable vehicle steamroller them. They can repurpose their own electric busses into a similarly reusable inter-orbital shuttle, I'd think. True, at the moment they seem to be thinking of using these as stopgap kludges, to be launched with the payload on a one-shot basis. But if it can make sense to design a Centaur-based hydrogen-oxygen transstage that ferries payloads from LEO to GEO and then returns to an orbiting base (or meets a rocket sent up for the one-shot purpose of refueling it) and then picks up another load for GEO---well, the electric buses seem even better suited to that purpose! Half the mass, or nearly, of the TransOrbital missions will be chemical fuel ferried up from Earth; every time you want to send 3 tons to GEO, you still also need to launch another 3 tons of the propellants, or so. It would help if the latter were part of a pipeline of routine prop replenishment missions launched to a storage station, but one way or another the customer is still going to be paying for that propellant, albeit perhaps with some modest economies of scale helping cheapen it a bit.
The electric buses, on the other hand, while I suppose if they are solar powered will be quite slow (taking months or more to bring their loads to target--actually looking around the Web the past year or so I'd think years would not be inaccurate) will use really tiny amounts of propellant. Well, that depends on exactly what sort of ISP we are talking about--bearing in mind a spiraling low-thrust trajectory is far less efficient of delta-V than a single high-thrust impulse (I gather perhaps double the impulse, due to foregoing the benefits of the Oberth effect) if the ISP is say 2000, and the delta-V say 5000, we might need to use 1/3 the total orbited mass, and that's including the bus mass, and before considering that we want to bring the bus back and use it again. Still, even this would cut the propellant budget in half, and the mass needed to be launched for the total mission significantly lowered. Whereas the propellants used for electric propulsion are I believe more easily stored than hydrogen, and anyway having less mass are that much less of a problem.
So, in this moment of challenge, Lockheed might look into competing directly with Northrup in the orbital transfer biz.
That doesn't save their bacon in the basic launch to LEO business because what they are doing is reinforcing StarLaunch's competitiveness by offering a competitive means of rendering Thunderbolt's small payloads equivalent to the big payloads SL cannot as yet match, undercutting their own launch service all the more. But it keeps them operating on some basis or other while others might go under completely.
I suppose the slow timeframe of the electric orbital transfers might prove a major hitch, to be sure. After all by my recommendation they want their old bus back to use again--but that multiplies the effect of the time lag. Spiraling transfers are something I'd love to see a rigorous analysis of, but by my own methods of estimation, the inner orbital part recedes from or approaches LEO at an acceleration proportional to thrust to mass--coming back from GEO the bus alone will mass less than bus plus payload and so come in faster than it went out.
I figure the constraint on speed is that first of all electric thrusters of various types tend to have very low thrust-to-mass rates, and second to get them thrusting harder might not only require larger thruster masses but more importantly, very high power inputs. The general tradeoff in rocketry is, the higher the ISP, the higher the specific energy per unit of thrust obtained--thus, although a hydrogen-oxygen flame delivering ISP in the ballpark of 440 or so is already very powerful indeed, to multiply that ISP by 5 or more would then require 5 or more times the already prodigious power. Even nuclear reactors cannot be expected to deliver such powers. If we are OK with low thrust, then we need the same high energy but it is spread out over more time, meaning lower power--but to get 1/5 the thrust at 5 times the ISP we still need the same power. Solar power has many virtues but it can't get the payload where you want it to go fast; modern operations with electric propulsion operate on timescales of months or years, even decades.
Meanwhile the Northrup TransOrbital thing will be looping back to LEO in a matter of hours. If the refueling and mating to a new payload required little enough time it could put two cargoes into GEO a day, approaching three! Each time it uses as much propellant mass as the payload masses and then some to be sure.
So I guess I see why Lockheed doesn't want to jump into the orbital tug business despite the high ISP of their own tug. They'd need a power source orders of magnitude greater than solar panels to cut the transfer time down by factors of ten, if not more.
Oh well it was a thought.
Why Boeing doesn't get into the commercial launcher biz--is something we've discussed here before, IIRC:
1) Multibodies are huge by commercial launcher standards;
2) Boeing is comfortable with DoD and NASA orders, and since the former and even to some extent the latter is jealous of foreign operations, doesn't want to risk its plum deal with the US Government by selling to other operators (even purely US corporations--after all in the era of globalization, what company is pure US anyway?)
3) The Multibodies are designed to be launched from NASA's and the Air Force's particular facilities, at Canaveral and Vandenberg, and these sites tend to be kept busy by NASA and Air Force missions. It is more in Boeing's interest to keep these two customers busy ordering their own scheduled launches and if a lull were to occur it could be awkward arranging to use these two USG sites for commercial launches. Were they to find a market for their big rockets in defiance of point 1 and risk USG displeasure (or get a clear dispensation) to address point 2, they might still need to develop a third launch site, one with facilities, if not quite as elaborate as the gargantuan VAB at Canaveral, whatever monster hangar they have at Vandenberg. Building such facilities from scratch anywhere then staffing them is going to be a big budget item, not to be considered for a marginal market share. I did once have enthusiastic hopes of there being a branch site at Kourou, but that runs straight into issues of international competition, on both sides.
So Boeing is out of the picture and I am very confused by all this talk of other companies horning in on Saturn derivatives at this late date.
------
Alan Bond, again at this late date (if I have not forgotten some earlier mention of him being involved with Horus) is a bit of an Easter Egg, but I fear possibly one who will never be found before going rotten.
OTL I'm a big fan of Skylon, but that's partly from being a romantic who roots for underdogs. Mind I'm sold on the idea that his singular approach to SSTO can work, and at a stroke come much closer to the dream of an airplane-like mode of operations (land, refuel, load in the new payload whilst doing cursory routine diagnostics on the engines and other systems that generally confirm, yep, they are all ready to go again as they are with minor tweaking and oiling, then take off and go, deliver the next cargo and possibly pick up a small down cargo, and so on for tens of sorties). Also it looks as elegant as all hell, and its concept of operations is also elegant. But REI is an underdog because it doesn't have the backing of any of the big players, only just now attracting serious investment and scrutiny from HMG and the ESA.So that can give me hope--TTL's Bond et al are no more (if no less) in the wilderness than they were OTL at this date. Perhaps. We don't know if he's where OTL Bond was in the same year, behind or just maybe perhaps a bit ahead. But as per OTL he's a long shot coming out of left field.
---
I note though that there are really not one but two schemes to achieve transfer to GEO by secondary means. Lockheed need not panic nor let Northrop's "TransOrbital" reusable vehicle steamroller them. They can repurpose their own electric busses into a similarly reusable inter-orbital shuttle, I'd think. True, at the moment they seem to be thinking of using these as stopgap kludges, to be launched with the payload on a one-shot basis. But if it can make sense to design a Centaur-based hydrogen-oxygen transstage that ferries payloads from LEO to GEO and then returns to an orbiting base (or meets a rocket sent up for the one-shot purpose of refueling it) and then picks up another load for GEO---well, the electric buses seem even better suited to that purpose! Half the mass, or nearly, of the TransOrbital missions will be chemical fuel ferried up from Earth; every time you want to send 3 tons to GEO, you still also need to launch another 3 tons of the propellants, or so. It would help if the latter were part of a pipeline of routine prop replenishment missions launched to a storage station, but one way or another the customer is still going to be paying for that propellant, albeit perhaps with some modest economies of scale helping cheapen it a bit.
The electric buses, on the other hand, while I suppose if they are solar powered will be quite slow (taking months or more to bring their loads to target--actually looking around the Web the past year or so I'd think years would not be inaccurate) will use really tiny amounts of propellant. Well, that depends on exactly what sort of ISP we are talking about--bearing in mind a spiraling low-thrust trajectory is far less efficient of delta-V than a single high-thrust impulse (I gather perhaps double the impulse, due to foregoing the benefits of the Oberth effect) if the ISP is say 2000, and the delta-V say 5000, we might need to use 1/3 the total orbited mass, and that's including the bus mass, and before considering that we want to bring the bus back and use it again. Still, even this would cut the propellant budget in half, and the mass needed to be launched for the total mission significantly lowered. Whereas the propellants used for electric propulsion are I believe more easily stored than hydrogen, and anyway having less mass are that much less of a problem.
So, in this moment of challenge, Lockheed might look into competing directly with Northrup in the orbital transfer biz.
That doesn't save their bacon in the basic launch to LEO business because what they are doing is reinforcing StarLaunch's competitiveness by offering a competitive means of rendering Thunderbolt's small payloads equivalent to the big payloads SL cannot as yet match, undercutting their own launch service all the more. But it keeps them operating on some basis or other while others might go under completely.
I suppose the slow timeframe of the electric orbital transfers might prove a major hitch, to be sure. After all by my recommendation they want their old bus back to use again--but that multiplies the effect of the time lag. Spiraling transfers are something I'd love to see a rigorous analysis of, but by my own methods of estimation, the inner orbital part recedes from or approaches LEO at an acceleration proportional to thrust to mass--coming back from GEO the bus alone will mass less than bus plus payload and so come in faster than it went out.
I figure the constraint on speed is that first of all electric thrusters of various types tend to have very low thrust-to-mass rates, and second to get them thrusting harder might not only require larger thruster masses but more importantly, very high power inputs. The general tradeoff in rocketry is, the higher the ISP, the higher the specific energy per unit of thrust obtained--thus, although a hydrogen-oxygen flame delivering ISP in the ballpark of 440 or so is already very powerful indeed, to multiply that ISP by 5 or more would then require 5 or more times the already prodigious power. Even nuclear reactors cannot be expected to deliver such powers. If we are OK with low thrust, then we need the same high energy but it is spread out over more time, meaning lower power--but to get 1/5 the thrust at 5 times the ISP we still need the same power. Solar power has many virtues but it can't get the payload where you want it to go fast; modern operations with electric propulsion operate on timescales of months or years, even decades.
Meanwhile the Northrup TransOrbital thing will be looping back to LEO in a matter of hours. If the refueling and mating to a new payload required little enough time it could put two cargoes into GEO a day, approaching three! Each time it uses as much propellant mass as the payload masses and then some to be sure.
So I guess I see why Lockheed doesn't want to jump into the orbital tug business despite the high ISP of their own tug. They'd need a power source orders of magnitude greater than solar panels to cut the transfer time down by factors of ten, if not more.
Oh well it was a thought.
It's the S-IVB stage, the prime contractor was Douglas in begin of apollo program
in 1967 McDonnell and Douglas merge together and in TTL McDonnell Douglas merge with Lockheed somewhere in 1990s.
ITTL, McDonnell-Douglas continues making them for the Saturn 1C and Multibody families, which Lockheed inherits with the merger. Boeing does the first stage, and gets a lot more of the focus, but LockMac is there too. Of course, the joint aspect of these operations is yet a fourth reason why Boeing's not hugely interested in developing commercial Multibody applications--it's not totally their rocket to sell.WIKIPEDIA said:
Indeed it was, and one we gave some though to. The issue is that time money, especially for comsats, and spending a substantial fraction of the bird's life getting to the final position. Even IOTL, Boeing's been having some trouble getting customers for the 702 electric bus variants because of the loss of months worth of revenue for just the transfer from GTO to GSO, which is only about a third of what starting from LEO would entail. If chemical fuel's cheap enough, the extra time-on-station can be more than worth it.
I'm sorry I missed this before, but I wanted to say thanks, and on behalf of Workable Goblin and myself, welcome aboard and I hope you continue enjoying it if you catch up! There's a link to the AH.com wiki in my sig where you can find a list of just the canon posts, as well as collections of various images from the timeline universe if you'd like to get caught up faster.
Winchell's a cool guy--I sent him a message on G+ too, which may have something to do with it.Update: Got an e-mail from him yesterday, and he's already updated it. What a prompt guy !
Quite. I was pretty pleased with myself when I came up with this, and I'm glad to hear people are finding it plausible. It's an interesting thought--they make upper stages and might otherwise be put out of the launch game by RLVs, but by doing TransOrbital, they become a key part of the launch game for at least one, possibly more providers, and making LEO-only RLVs requires substantially smaller payload to orbit (and a lower-energy aerobrake) than going all the way to GTO, so it helps the launch companies (like SLS for the moment, but later maybe others) too.
Quite.
Workable Goblin and I have had some very interesting talks about the way it could shake things out, and I hope what we've got in for will be fun...I stand corrected on the matter of who makes Multibodies. Two stages, two contractors, how else would Uncle Sam do it but spread the pork around?
I had the impression that in the early '70s Dark Ages in the wake of the cancellation of the Lunar Apollo program, that the contractors were backing away from NASA biz and so Douglas sold off their rights and blueprints etc cheap to Boeing so Boeing owned the whole thing.
Now I suppose the solids too are made by a third contractor.
----
Another D'Oh moment for me--looking into the whole matter of electric propulsion (trying to get a general sense of which sort of electric propulsion would be flying in the early 2000s and what its general parameters would be) led me to read up on the example of the European SMART-1 mission, which used SNECMA-built Hall thrusters with xeon as reactant, getting ISP in the ballpark of 1600. The craft massed about 1/3 of a ton, used around 60 or so kg of xeon (or more, I might be picking up just on what was needed to bring it to L-1 and switch over into the Lunar effective gravity well, then it needed more to bring it down to low Lunar orbit)--but took over a year to get to L-1, and then months more to stabilize in LLO. Pretty much as I guessed.
But what I didn't guess was that this epic slow trip started with a boost to a geosynchronous transfer orbit, perigee of 7000 km (implying that its initial launched parking orbit was already unusually high) apogee over 42,000 km. That kick must have been in the ballpark of 2000-2500 m/sec, delivered by chemical rocket, so the propellant for that must nearly have doubled the launch mass, and counting an extra rocket stage to do it, probably more than doubled.
In my naivete all my efforts to estimate the parameters of low-thrust high-ISP electric thrusters did assume starting from LEO parking orbit; I observed that if that were so, a very long portion of the flight would be spent spiraling out in nearly circular widening orbits that would have a radial component that is a slow crawl compared to the orbital speed. The way I looked at it, the tiny acceleration in an essentially tangential direction is multiplied by the radius to get increment of the angular momentum, and the actual osculating orbit that results at any instant would be very close to a circular orbit with the slowly increasing angular momentum. Another way to say this--if the craft has an outward radial component, conservation of angular momentum will slow the tangential component, so the speed outward a given tangential thrust could maintain would be the one where increasing radius slowing the craft down tangentially is compensated by the thrust. Or not quite compensated, because the successive outward circular approximate orbits have decreasing orbital speeds. Thus in the near-Earth region, the radial velocity is negligible compared to the tangential, and even if we seek to maximize rate of increasing orbital energy (by always thrusting in line with the craft's current motion and not at some angle to it) we very closely approximate tangential thrust that applies a steady, and increasing, increment of angular momentum as the craft slowly recedes from Earth.
Now gradually of course the orbital speed is slowing whereas the equilibrium radial speed a given thrust can maintain is rising (since (circular) orbital speed falls with the inverse square root of distance, and angular momentum therefore rises as the 1/2 power of distance, while a given acceleration increments angular momentum in proportion to radius. If we were trying to maximize the rate of increasing orbital energy, we'd therefore be shifting the thruster to angle outward more and more radially--even if we don't do that, the radial component rises and the tangential one falls so that we eventually reach escape velocity with the two speeds being equal and net orbital energy raised to zero--that is the osculating orbit at that point is an escape parabola. So long before then the neat little approximation above would break down. But it gave me a tool to estimate just how long the thrusters have to push to reach a given distance, until the radial speeds start becoming a big fractio of the circular orbital.
But when we start from an elongated transfer to GTO orbit like that, those approximations go out the window. And to be sure, for a solar powered craft starting from LEO they are not much good anyway, because almost half the time a LEO satellite is in shadow and the solar cells are not putting out power. If we were to launch from LEO like that, such a craft would have thrust only on the daylight side, which would include perigee, so basically the impulse could be roughly approximated by a single instantaneous burn equivalent to the time it spends in sunlight times the actual thrust rate. Even summed over 45-50 minutes like that the impulse is tiny so we'd wind up with nearly-circular orbits with perigee staying roughly the same and apogee rising slowly, roughly on the far side of Earth from the Sun. Gradually the craft would be spending more time in sunlight and so its net impulse per orbit would rise--but mostly because the osculating orbital period was rising. Also the slowest part of the orbit would be near the eclipse zone.
Launching into GTO, all of a sudden the craft is in an orbit where it spends most of its time 5 to 7 times as far away from Earth as at perigee, buying longer periods, and less eclipse time as a fraction of the orbital period. SMART-1 was on its way to the Moon and so I suppose its thrust was for most of the time it was on optimized to increase energy, and then later on to increase angular momentum so that its orbit circularized--but looking at illustrations of snapshots of various osculating orbits I can see that it started circularizing early on, that is lowering its eccentricity. With the craft spending most of its time far out on the orbit, where a given thrust in the tangential direction would deliver a bigger increment of angular momentum, whereas at lower orbital speeds (far below the circular orbital speed at those radii) the orbital energy increment was relatively low I can see how this must be so, even if one is trying to maximize the latter rate.
The slow part of a spiral out from LEO is the low part, and while boosting with electric thrust from there would indeed still save a lot of reaction mass, it also takes the ship on a leisurely barge ride through the worst intensities of the Van Allen Belts, particularly the low one that is most intense and that an Apollo-like Lunar trajectory could avoid almost completely since that belt covers a limited range of latitudes. A spiraling path from LEO has it cruising at a snails pace there in the most concentrated part of the belt for months. This prospect alarms me enough when I consider a considerably higher thrust (but still absolutely low-thrust) fusion pulse drive for manned flight from LEO to Luna. SMART-1's trajectory reminds me that other discussion of electric propulsion I've seen worries about the Van Allen Belt passage issue too--not for any hazard to human astronauts on these very slow passages, but due to damage the radiation there can do to the craft itself, notably to the solar power panels.
One possible solution is to harden the craft to endure the passage--and if we had Dr. Slough's electromagnetically driven fusion pulse rocket ready to hand today, where a 15-20 ton propulsion module could pulse a .4 kg mass of lithium at ISP of 5000 for a 20,000 Newton-second impulse once a minute (maybe once every ten seconds) so that the lunar transfer would be not a year but a few weeks for a craft massing many tens of tons all up, we could similarly harden the craft as we would in any event for solar flares, providing a "storm shelter" shielded by supplies composed mostly of low-atomic-weight elements, thus they too could withstand the Belts despite spending a couple weeks in them. But the alternative, despite the considerable cost, of simply boosting the ship with a high-thrust chemical impulse so it is well on its way, appeals--even though the subsequent spiraling orbits will still hat o many crossings of the belts, at least these will happen at high radial speeds, cutting down exposure time, and the initial boost shaves a considerable amount of incremental boosting time and places the craft where its low-thrust rocket is most efficient. It would seem SMART-1 had a similar option chosen to deal in part with the belts and also to save time, trading off much of the mass savings advantage for these benefits.
So--I was wrongly assuming that Lockheed ITTL was going to use electric thrusters to raise GEO satellites in steady spiraling near-circular orbits from LEO directly. Instead it looks like some kind of hybrid akin to the trajectory SMART-1 followed--despite the mass cost, a chemical boost of some 2500 m/sec initially to put the payload and thruster bus into transfer orbit, then a slow steady circularization of that orbit.
Now looking at SMART-1 and figures for the masses of solar panels versus their power output, it had a single SNECMA PPS-1350-G that massed 5.3 kg, delivered (as used on the mission) up to 68 mN at 1200 watts of power (SNECMA claimed up to 88 at 1500 W) and the Wikipedia page on SMART-1 cites 29 kg for the electrical propulsion system mass (excluding obviously the 80+ kg of Xeon propellant) so that 29 kg may include the solar panels. (It might not since those panels obviously had other functions too). Another Wikipedia page says 300 W/kg "are available" from space-optimized panels, and while this might not reflect the state of the art when SMART-1 was launched it suggests just 4 kg can account for SMART-1's power needs for this thruster. Say it's 36 kg all up for a system that can use 1200 W to put out 72 mN and we have 1/2 kg all up per m/N, which suggests to me a dedicated heavy bus massing 2 tons could put out 4 Newtons, using 60 of these thrusters drawing 72 kW from a quarter-ton solar panel array. In proportion then a 20 ton all up craft, comprising the two ton bus, 5 tons of xeon propellant, and 13 tons of payload, could be moved from LEO to low Lunar orbit after being boosted some 2500 m/sec using an 18 ton hydrogen-oxygen booster in 15 months on the same path as SMART-1.
That does not sound so great but I'd think payloads to GEO would typically be a lot smaller than 13 tons; say only 4 tons, then the dry mass falls to 6 from 15, the orbital delta-V the solar thruster must achieve is--well, that's a bit tricky, but say 2/3 what SMART-1 had to accomplish, so the propellant would be reduced to under 2 tons. Now we have 8 tons all up to boost to GTO with an 8 ton chemical rocket, 16 tons all up in LEO versus 38, and we still have some reserve propellant that might get the solar thruster back to LEO. How long would the trip out take? If the delta-V is 2/3 and the thrust to mass ratio initially 2.5 times greater we have maybe 4 months out. That would imply, thrusting about 60 percent of the time, propellant consumption of 1850 kg, so no we wouldn't have a lot of reserve left after all. To spiral the 2 ton bus back down to LEO--well, we wouldn't want to try to recover it from a GTO orbit so we can't just reverse the path out.
Trying to figure it as a direct boost from LEO using the thrusters all the way is tricky, especially considering the gradually reducing effect the shadow of the Earth has. I'd guess the thrust time and hence propellent consumption would more than double, to 4 tons say, and if we want to get the bus back, we'd need 1.3333 tons of extra propellant, which would raise the outbound propellant need to 5. Say 6.5 tons of propellant, 12.5 all up, it takes 8 months to boost the 6 ton payload-bus combo up to GEO--but then, the bus massing about 1/3 that it will only take a third that time to bring it back, so a bit over a year round trip, and the customer is waiting 8 months from launch for their payload to go on line. But, assuming the bus has already been launched before, just 10.5 tons need to be launched, and Lockheed gets their bus back.
Now notice, the SNECMA early-2000s Hall thruster model only gets 1600 ISP. I gather that the theoretical upper limit of Hall thrusters might be pushed to double that. For the same power input, that cuts the already weak thrust in half, but it also would cut the propellant masses needed down. Say now we only need .75 tons to bring the bus back down, cutting 7.5 outbound tons to 6.75; with doubled mass efficiency we need only 2.25 for the outbound leg, or 9 all up, down to 7 tons launched to LEO, just 3 of which is propellant, 4 being GEO payload. The whole stack is now 6/7 previous mass and .9 at GEO meaning we should shave some time off the outward boost. But wait! At constant power we've cut the thrust in half, so now it takes the better part of two years! However, not to worry, we can probably up the power--a factor of 4 will restore the original mass flow rate, and now the doubled ISP appears as double instead of half thrust. The price we pay for that is needing to quadruple 72 KW to 288, raising the array from a quarter ton to a full ton, adding 750 kg to the 2 ton bus. Well, that will undo a lot of the mass saving above but anyway we only launch the bus once.
Now we need a full ton of propellant to bring the 2.75 bus back down; we need to arrive at GEO with 7.75 tons altogether; we need 2.6 tons of xeon for the upward leg, or 10.35 tons in LEO, 7.6 were launched for this mission, 4 of payload, 3.6 of propellant.
Allowing for the different masses changing the time profile versus doubled thrust, I figure the whole cycle now takes 5 months, with the payload arriving at GEO well before 4 months have passed.
Turning instead to a chemical alternative, specifying that it too is a 2 ton dry bus that can boost at ISP 440 (typical of Centaur) and must be returned to LEO from a GEO delivery--I'm going to leave out inclination change which implies that LEO is equatorial, hence our launch was from Kourou. I didn't try to account for inclination change above, so I'm skipping it now! But these American launchers are going to want to launch to 30 degrees or more inclination and I think that adds 500 m/sec delta-V requirement at GEO circularization (and sending the bus back down to a 30 degree inclination too). Skipping that 1000 m/sec in the process merrily, if we want to return a 2 ton bus from GEO to LEO, it needs to go through 1500 m/sec deorbiting burn up high and then lose 2500 m/sec at perigee, approximately--4000 m/sec all up, so the two ton bus rocket needs a bit over 3 tons fuel just to get itself back home. Thus the load going "up" is 4 tons payload plus 5 tons bus and return fuel, or 9, this too needs to go through delta-V of 4000 at least, so again we need to add a bit over 1.5 or 13.5 tons, 22.5 all up in LEO to deliver 4 tons to GEO and return the bus--which comes back to LEO in just 8 hours plus however long it takes to fuss around in GEO to set things just right there, plus possible coasting time up to 24 hours to position the returning bus to the right part of its low equatorial orbit. 20.5 of the total had to be launched this time around; the two ton bus rocket needs to hold 14.5 tons of hydro-lox propellant. I suppose that is roughly in the right ballpark for a plausible rocket. Over 4 times the payload to GEO is fuel.
Could it have been possible for the bus to skip the 2500 m/sec braking to LEO by aero-skipping off the atmosphere to lose it instead? Say we insist on keeping 500 m/sec as maneuvering reserve, but save 2000. This means we deorbit 2.25 tons through 1500 m/sec; we need the booster to mass 3.2 now instead of 5, so up mass to GEO is 7.2, and 1.5 times that is 10.8 so the tank needs to hold 12 tons of propellant; we have all up mass of 18 tons or 4 and a half less than an all-rocket strategy demands; we only need to launch 16, of which 4 is payload; 3 times the payload to GEO is fuel.
So--pushing the envelope toward two ambitious stretches of state of the art--one where we double the ISP of a Hall thruster but maintain mass flow by quadrupling power throughput and hence double thrust; the other where we must design the two ton (dry) chemical booster to fly through an atmospheric aeroskip maneuver to LEO and therefore it needs both careful aerodynamic design and non-ablative TPS that doesn't get degraded in long stays in orbital space, we can contemplate a tradeoff of a slow transfer of almost 4 months and then a month wait to get the bus back, but using less propellant mass than payload, versus having to launch 4 times the payload mass to deliver fuel but getting the bus back later the same day.
If we deny either of these perhaps dubious advances, we can compare an even slower electric bus that takes the better part of a year to cycle but requires only 1 2/3 the payload mass in propellant, versus a very conventional Centaur type bus that requires we launch over 5 times the payload mass to orbit to deliver it to GEO and return the bus.
Note something else about the electric thrusters when frowning at their very slow transit times--we can add propellant tankage to them relatively cheaply in mass terms, and by paying an even bigger penalty in transit time, push bigger payloads to a given goal, or a given payload farther. The chemical buses, on the other hand, seem just about right-sized for this 4-ton payload I pulled out of the air intuitively. In the conventional case I think about the 14.5 tons of hydro-LOX I guessed is as much as 2 tons of structure, less engine, thrust structure and docking ring can hold; to push a bigger payload we either need more buses or to push it to a closer target. The aero-skipping version needs significantly less propellant, but designing a structure that can hold that lesser fuel load and also safely aeroskip every time probably will cut the tankage available down.
Comparing the extremes, the advanced 3200 ISP Hall thruster that still takes 4 months to deliver the goods to GEO versus the almost completely conventional Centaur with a docking ring, we need to launch just under twice a 4 ton payload to get that slow delivery accomplished, versus the old-fashioned Centaur tug that can offer same-day delivery, but requires five times the payload mass to be launched. Is the cost of launching the extra 12 tons worth saving by waiting 4 months? If not, saving less mass by waiting 8 months is clearly a bad deal, and if not, Hall thrusters as electric transfer vehicles are a non-starter for commercial business. They remain desirable for what they have been used for OTL thus far, which is deep space exploration where their slow accelerations cease to be such a drawback.
And then, electric propulsion will have to await a new power source to be useful for anything other than station-keeping and niche applications. Nuclear fission seems unlikely to be mass-effective, competing with solar panels. By the way, can we do that trick I did with the Hall thruster and just quadruple power again to get double ISP and double thrust? Well for one thing that would if possible mean taking a one ton solar array and making it 4 tons, or tripling the bus mass from the original version, so clearly we would be hitting diminishing returns. And no we can't; I was taking the high-end figure of on-line estimates of the maximum theoretical ISP for Hall thrusters and maybe exceeding it too. We can't expect it to be physically possible to double the ISP again. We might still consider doubling the power to double the bus mass all up to 5 tons and speeding up transfers a bit, but again that way lie diminishing returns.
It looks to me like we are just going to have to rely on good old hydrogen-oxygen engines and keep launching lavish amounts of propellant to service them. In fact, it looks like maybe trying to recover a Centaur-sized rocket is kind of marginal economically speaking, considering the extra propellant we need to keep launching to get it back. It might make more sense to just make the rocket cheaper and dispose of it once used.
I had the impression that in the early '70s Dark Ages in the wake of the cancellation of the Lunar Apollo program, that the contractors were backing away from NASA biz and so Douglas sold off their rights and blueprints etc cheap to Boeing so Boeing owned the whole thing.
Now I suppose the solids too are made by a third contractor.
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Another D'Oh moment for me--looking into the whole matter of electric propulsion (trying to get a general sense of which sort of electric propulsion would be flying in the early 2000s and what its general parameters would be) led me to read up on the example of the European SMART-1 mission, which used SNECMA-built Hall thrusters with xeon as reactant, getting ISP in the ballpark of 1600. The craft massed about 1/3 of a ton, used around 60 or so kg of xeon (or more, I might be picking up just on what was needed to bring it to L-1 and switch over into the Lunar effective gravity well, then it needed more to bring it down to low Lunar orbit)--but took over a year to get to L-1, and then months more to stabilize in LLO. Pretty much as I guessed.
But what I didn't guess was that this epic slow trip started with a boost to a geosynchronous transfer orbit, perigee of 7000 km (implying that its initial launched parking orbit was already unusually high) apogee over 42,000 km. That kick must have been in the ballpark of 2000-2500 m/sec, delivered by chemical rocket, so the propellant for that must nearly have doubled the launch mass, and counting an extra rocket stage to do it, probably more than doubled.
In my naivete all my efforts to estimate the parameters of low-thrust high-ISP electric thrusters did assume starting from LEO parking orbit; I observed that if that were so, a very long portion of the flight would be spent spiraling out in nearly circular widening orbits that would have a radial component that is a slow crawl compared to the orbital speed. The way I looked at it, the tiny acceleration in an essentially tangential direction is multiplied by the radius to get increment of the angular momentum, and the actual osculating orbit that results at any instant would be very close to a circular orbit with the slowly increasing angular momentum. Another way to say this--if the craft has an outward radial component, conservation of angular momentum will slow the tangential component, so the speed outward a given tangential thrust could maintain would be the one where increasing radius slowing the craft down tangentially is compensated by the thrust. Or not quite compensated, because the successive outward circular approximate orbits have decreasing orbital speeds. Thus in the near-Earth region, the radial velocity is negligible compared to the tangential, and even if we seek to maximize rate of increasing orbital energy (by always thrusting in line with the craft's current motion and not at some angle to it) we very closely approximate tangential thrust that applies a steady, and increasing, increment of angular momentum as the craft slowly recedes from Earth.
Now gradually of course the orbital speed is slowing whereas the equilibrium radial speed a given thrust can maintain is rising (since (circular) orbital speed falls with the inverse square root of distance, and angular momentum therefore rises as the 1/2 power of distance, while a given acceleration increments angular momentum in proportion to radius. If we were trying to maximize the rate of increasing orbital energy, we'd therefore be shifting the thruster to angle outward more and more radially--even if we don't do that, the radial component rises and the tangential one falls so that we eventually reach escape velocity with the two speeds being equal and net orbital energy raised to zero--that is the osculating orbit at that point is an escape parabola. So long before then the neat little approximation above would break down. But it gave me a tool to estimate just how long the thrusters have to push to reach a given distance, until the radial speeds start becoming a big fractio of the circular orbital.
But when we start from an elongated transfer to GTO orbit like that, those approximations go out the window. And to be sure, for a solar powered craft starting from LEO they are not much good anyway, because almost half the time a LEO satellite is in shadow and the solar cells are not putting out power. If we were to launch from LEO like that, such a craft would have thrust only on the daylight side, which would include perigee, so basically the impulse could be roughly approximated by a single instantaneous burn equivalent to the time it spends in sunlight times the actual thrust rate. Even summed over 45-50 minutes like that the impulse is tiny so we'd wind up with nearly-circular orbits with perigee staying roughly the same and apogee rising slowly, roughly on the far side of Earth from the Sun. Gradually the craft would be spending more time in sunlight and so its net impulse per orbit would rise--but mostly because the osculating orbital period was rising. Also the slowest part of the orbit would be near the eclipse zone.
Launching into GTO, all of a sudden the craft is in an orbit where it spends most of its time 5 to 7 times as far away from Earth as at perigee, buying longer periods, and less eclipse time as a fraction of the orbital period. SMART-1 was on its way to the Moon and so I suppose its thrust was for most of the time it was on optimized to increase energy, and then later on to increase angular momentum so that its orbit circularized--but looking at illustrations of snapshots of various osculating orbits I can see that it started circularizing early on, that is lowering its eccentricity. With the craft spending most of its time far out on the orbit, where a given thrust in the tangential direction would deliver a bigger increment of angular momentum, whereas at lower orbital speeds (far below the circular orbital speed at those radii) the orbital energy increment was relatively low I can see how this must be so, even if one is trying to maximize the latter rate.
The slow part of a spiral out from LEO is the low part, and while boosting with electric thrust from there would indeed still save a lot of reaction mass, it also takes the ship on a leisurely barge ride through the worst intensities of the Van Allen Belts, particularly the low one that is most intense and that an Apollo-like Lunar trajectory could avoid almost completely since that belt covers a limited range of latitudes. A spiraling path from LEO has it cruising at a snails pace there in the most concentrated part of the belt for months. This prospect alarms me enough when I consider a considerably higher thrust (but still absolutely low-thrust) fusion pulse drive for manned flight from LEO to Luna. SMART-1's trajectory reminds me that other discussion of electric propulsion I've seen worries about the Van Allen Belt passage issue too--not for any hazard to human astronauts on these very slow passages, but due to damage the radiation there can do to the craft itself, notably to the solar power panels.
One possible solution is to harden the craft to endure the passage--and if we had Dr. Slough's electromagnetically driven fusion pulse rocket ready to hand today, where a 15-20 ton propulsion module could pulse a .4 kg mass of lithium at ISP of 5000 for a 20,000 Newton-second impulse once a minute (maybe once every ten seconds) so that the lunar transfer would be not a year but a few weeks for a craft massing many tens of tons all up, we could similarly harden the craft as we would in any event for solar flares, providing a "storm shelter" shielded by supplies composed mostly of low-atomic-weight elements, thus they too could withstand the Belts despite spending a couple weeks in them. But the alternative, despite the considerable cost, of simply boosting the ship with a high-thrust chemical impulse so it is well on its way, appeals--even though the subsequent spiraling orbits will still hat o many crossings of the belts, at least these will happen at high radial speeds, cutting down exposure time, and the initial boost shaves a considerable amount of incremental boosting time and places the craft where its low-thrust rocket is most efficient. It would seem SMART-1 had a similar option chosen to deal in part with the belts and also to save time, trading off much of the mass savings advantage for these benefits.
So--I was wrongly assuming that Lockheed ITTL was going to use electric thrusters to raise GEO satellites in steady spiraling near-circular orbits from LEO directly. Instead it looks like some kind of hybrid akin to the trajectory SMART-1 followed--despite the mass cost, a chemical boost of some 2500 m/sec initially to put the payload and thruster bus into transfer orbit, then a slow steady circularization of that orbit.
Now looking at SMART-1 and figures for the masses of solar panels versus their power output, it had a single SNECMA PPS-1350-G that massed 5.3 kg, delivered (as used on the mission) up to 68 mN at 1200 watts of power (SNECMA claimed up to 88 at 1500 W) and the Wikipedia page on SMART-1 cites 29 kg for the electrical propulsion system mass (excluding obviously the 80+ kg of Xeon propellant) so that 29 kg may include the solar panels. (It might not since those panels obviously had other functions too). Another Wikipedia page says 300 W/kg "are available" from space-optimized panels, and while this might not reflect the state of the art when SMART-1 was launched it suggests just 4 kg can account for SMART-1's power needs for this thruster. Say it's 36 kg all up for a system that can use 1200 W to put out 72 mN and we have 1/2 kg all up per m/N, which suggests to me a dedicated heavy bus massing 2 tons could put out 4 Newtons, using 60 of these thrusters drawing 72 kW from a quarter-ton solar panel array. In proportion then a 20 ton all up craft, comprising the two ton bus, 5 tons of xeon propellant, and 13 tons of payload, could be moved from LEO to low Lunar orbit after being boosted some 2500 m/sec using an 18 ton hydrogen-oxygen booster in 15 months on the same path as SMART-1.
That does not sound so great but I'd think payloads to GEO would typically be a lot smaller than 13 tons; say only 4 tons, then the dry mass falls to 6 from 15, the orbital delta-V the solar thruster must achieve is--well, that's a bit tricky, but say 2/3 what SMART-1 had to accomplish, so the propellant would be reduced to under 2 tons. Now we have 8 tons all up to boost to GTO with an 8 ton chemical rocket, 16 tons all up in LEO versus 38, and we still have some reserve propellant that might get the solar thruster back to LEO. How long would the trip out take? If the delta-V is 2/3 and the thrust to mass ratio initially 2.5 times greater we have maybe 4 months out. That would imply, thrusting about 60 percent of the time, propellant consumption of 1850 kg, so no we wouldn't have a lot of reserve left after all. To spiral the 2 ton bus back down to LEO--well, we wouldn't want to try to recover it from a GTO orbit so we can't just reverse the path out.
Trying to figure it as a direct boost from LEO using the thrusters all the way is tricky, especially considering the gradually reducing effect the shadow of the Earth has. I'd guess the thrust time and hence propellent consumption would more than double, to 4 tons say, and if we want to get the bus back, we'd need 1.3333 tons of extra propellant, which would raise the outbound propellant need to 5. Say 6.5 tons of propellant, 12.5 all up, it takes 8 months to boost the 6 ton payload-bus combo up to GEO--but then, the bus massing about 1/3 that it will only take a third that time to bring it back, so a bit over a year round trip, and the customer is waiting 8 months from launch for their payload to go on line. But, assuming the bus has already been launched before, just 10.5 tons need to be launched, and Lockheed gets their bus back.
Now notice, the SNECMA early-2000s Hall thruster model only gets 1600 ISP. I gather that the theoretical upper limit of Hall thrusters might be pushed to double that. For the same power input, that cuts the already weak thrust in half, but it also would cut the propellant masses needed down. Say now we only need .75 tons to bring the bus back down, cutting 7.5 outbound tons to 6.75; with doubled mass efficiency we need only 2.25 for the outbound leg, or 9 all up, down to 7 tons launched to LEO, just 3 of which is propellant, 4 being GEO payload. The whole stack is now 6/7 previous mass and .9 at GEO meaning we should shave some time off the outward boost. But wait! At constant power we've cut the thrust in half, so now it takes the better part of two years!
However, not to worry, we can probably up the power--a factor of 4 will restore the original mass flow rate, and now the doubled ISP appears as double instead of half thrust. The price we pay for that is needing to quadruple 72 KW to 288, raising the array from a quarter ton to a full ton, adding 750 kg to the 2 ton bus. Well, that will undo a lot of the mass saving above but anyway we only launch the bus once. Now we need a full ton of propellant to bring the 2.75 bus back down; we need to arrive at GEO with 7.75 tons altogether; we need 2.6 tons of xeon for the upward leg, or 10.35 tons in LEO, 7.6 were launched for this mission, 4 of payload, 3.6 of propellant.
Allowing for the different masses changing the time profile versus doubled thrust, I figure the whole cycle now takes 5 months, with the payload arriving at GEO well before 4 months have passed.
Turning instead to a chemical alternative, specifying that it too is a 2 ton dry bus that can boost at ISP 440 (typical of Centaur) and must be returned to LEO from a GEO delivery--I'm going to leave out inclination change which implies that LEO is equatorial, hence our launch was from Kourou. I didn't try to account for inclination change above, so I'm skipping it now! But these American launchers are going to want to launch to 30 degrees or more inclination and I think that adds 500 m/sec delta-V requirement at GEO circularization (and sending the bus back down to a 30 degree inclination too). Skipping that 1000 m/sec in the process merrily, if we want to return a 2 ton bus from GEO to LEO, it needs to go through 1500 m/sec deorbiting burn up high and then lose 2500 m/sec at perigee, approximately--4000 m/sec all up, so the two ton bus rocket needs a bit over 3 tons fuel just to get itself back home. Thus the load going "up" is 4 tons payload plus 5 tons bus and return fuel, or 9, this too needs to go through delta-V of 4000 at least, so again we need to add a bit over 1.5 or 13.5 tons, 22.5 all up in LEO to deliver 4 tons to GEO and return the bus--which comes back to LEO in just 8 hours plus however long it takes to fuss around in GEO to set things just right there, plus possible coasting time up to 24 hours to position the returning bus to the right part of its low equatorial orbit. 20.5 of the total had to be launched this time around; the two ton bus rocket needs to hold 14.5 tons of hydro-lox propellant. I suppose that is roughly in the right ballpark for a plausible rocket. Over 4 times the payload to GEO is fuel.
Could it have been possible for the bus to skip the 2500 m/sec braking to LEO by aero-skipping off the atmosphere to lose it instead? Say we insist on keeping 500 m/sec as maneuvering reserve, but save 2000. This means we deorbit 2.25 tons through 1500 m/sec; we need the booster to mass 3.2 now instead of 5, so up mass to GEO is 7.2, and 1.5 times that is 10.8 so the tank needs to hold 12 tons of propellant; we have all up mass of 18 tons or 4 and a half less than an all-rocket strategy demands; we only need to launch 16, of which 4 is payload; 3 times the payload to GEO is fuel.
So--pushing the envelope toward two ambitious stretches of state of the art--one where we double the ISP of a Hall thruster but maintain mass flow by quadrupling power throughput and hence double thrust; the other where we must design the two ton (dry) chemical booster to fly through an atmospheric aeroskip maneuver to LEO and therefore it needs both careful aerodynamic design and non-ablative TPS that doesn't get degraded in long stays in orbital space, we can contemplate a tradeoff of a slow transfer of almost 4 months and then a month wait to get the bus back, but using less propellant mass than payload, versus having to launch 4 times the payload mass to deliver fuel but getting the bus back later the same day.
If we deny either of these perhaps dubious advances, we can compare an even slower electric bus that takes the better part of a year to cycle but requires only 1 2/3 the payload mass in propellant, versus a very conventional Centaur type bus that requires we launch over 5 times the payload mass to orbit to deliver it to GEO and return the bus.
Note something else about the electric thrusters when frowning at their very slow transit times--we can add propellant tankage to them relatively cheaply in mass terms, and by paying an even bigger penalty in transit time, push bigger payloads to a given goal, or a given payload farther. The chemical buses, on the other hand, seem just about right-sized for this 4-ton payload I pulled out of the air intuitively. In the conventional case I think about the 14.5 tons of hydro-LOX I guessed is as much as 2 tons of structure, less engine, thrust structure and docking ring can hold; to push a bigger payload we either need more buses or to push it to a closer target. The aero-skipping version needs significantly less propellant, but designing a structure that can hold that lesser fuel load and also safely aeroskip every time probably will cut the tankage available down.
Comparing the extremes, the advanced 3200 ISP Hall thruster that still takes 4 months to deliver the goods to GEO versus the almost completely conventional Centaur with a docking ring, we need to launch just under twice a 4 ton payload to get that slow delivery accomplished, versus the old-fashioned Centaur tug that can offer same-day delivery, but requires five times the payload mass to be launched. Is the cost of launching the extra 12 tons worth saving by waiting 4 months? If not, saving less mass by waiting 8 months is clearly a bad deal, and if not, Hall thrusters as electric transfer vehicles are a non-starter for commercial business. They remain desirable for what they have been used for OTL thus far, which is deep space exploration where their slow accelerations cease to be such a drawback.
And then, electric propulsion will have to await a new power source to be useful for anything other than station-keeping and niche applications. Nuclear fission seems unlikely to be mass-effective, competing with solar panels. By the way, can we do that trick I did with the Hall thruster and just quadruple power again to get double ISP and double thrust? Well for one thing that would if possible mean taking a one ton solar array and making it 4 tons, or tripling the bus mass from the original version, so clearly we would be hitting diminishing returns. And no we can't; I was taking the high-end figure of on-line estimates of the maximum theoretical ISP for Hall thrusters and maybe exceeding it too. We can't expect it to be physically possible to double the ISP again. We might still consider doubling the power to double the bus mass all up to 5 tons and speeding up transfers a bit, but again that way lie diminishing returns.
It looks to me like we are just going to have to rely on good old hydrogen-oxygen engines and keep launching lavish amounts of propellant to service them. In fact, it looks like maybe trying to recover a Centaur-sized rocket is kind of marginal economically speaking, considering the extra propellant we need to keep launching to get it back. It might make more sense to just make the rocket cheaper and dispose of it once used.
Hi everyone. Sorry for the delay in this week's illustration, we had a few late adjustments to make. Here is a presentation of TransOrbital's concept of operations.
To summarize briefly, if I might:
(1) Electric transfer tugs for LEO-to-GTO suffer too many drawbacks to make a reasonable alternative in the eyes of a comsat buyer. Could work, but there's a lot of time value in the satellite's life lost in nearly a year or spiraling with the best current technology.
(2) It'd be nice if you could use aerobrake/capture for some of the tug's return to LEO, because it'd dramatically cut down on the reuse penalty of the prop you have to carry to GTO with the payload in order to reuse the tug, but no one's done that in practice and it'd mean TransOrbital would cost Northrop a lot more in development than here, where it's probably under a billion in terms of actual R&D costs (studies and such to sell it to comsat providers and users might be more, but that's a cost they're sharing with StarLaunch).
(3) At the very end, Shevek raises the question of whether the prop launch costs to recover the tug are higher than simply launching a new tug every time and expending it as you would a traditional second stage.
My math shows about 4.2 metric tons of prop are used by the tug for GTO-to-LEO return and carrying that prop LEO-to-GTO. At $2,500/kg, that has a price to Northrop of about $10.6m. My best-guess estimate for a Centaur's cost ITTL is only a million or so more, so it's a valid question. However, I think they'd prefer to reuse the tug for a couple reasons.
First, even a savings of a million or so adds up over several flight. Second, a reusable tug means that once the system is launched, the support costs are lower--essentially, once it's operational, all you have to launch is prop. Third, at the moment, every Delta and even Thunderbolt expends a Centaur in LEO, and the price of prop is $2,500/kg. However, if and when Thunderbolt or a competitor develops a reusable second stage, then the launch cost for these expendable Centaurs must also be added to their build cost, while the price of prop (and thus the "reusability tax") will drop, tilting even further. Northrop's willing to gamble that'll happen eventually, but not enough to immediately sink a billion or two into an aerocapture tug (which would cost only about $1m in reusability tax).