Gemini: We can rebuild it, we have the technology

najab · Dec 6, 2004

>No I understand what he posted. How do you link the APAS-89 to the 'hatch'(s)?I think I may not have been clear enough. GX-3 wouldn't have the top-mounted gull-wing doors of GeminiX-1. As I picture it, there would be a round hatch on the side for ground access and a (larger?) hatch in the nose for use while docked. The docking hatch would, as it's name suggests, have to have some kind of docking mechanism associated with it.The point I was trying to make is that the smaller aeronautics firms mrmorris has tapped to build GX-3 all have experience with 'simple' round pressure doors, but they do not have experience with docking adapters for space vehicles. If your docking adapter fails while you are docked, two crews die. Therefore it is better to use an 'off-the-shelf' design which has already been tested extensively than to try and develop your own.> The Gemini docked to the Agena upper stage and that wasn't an APAS-89. APAS-001? The GX-1 docking adapter was exactly that - it allowed docking only. It wasn't androgenous which meant, for example, that two 'female' ports couldn't mate (oh, those restrictive times!) The APAS system means that each port can go either way - two GX-3's could dock in orbit for example - something which couldn't be done with GX-1.>It does work very well, but it is only used for one application, docking to ISS.As I was just saying, APAS-89 can be used for dockings between any two vehicles so equipped, it's a common standard: GX-3/Soyuz could work for example.

Swampcat · Dec 6, 2004

I have been following this thread and find it very interesting. I can imagine you must be putting a lot of effort into it and selfishly hope you will continue.<img src="/images/icons/smile.gif" /> Concerning the docking adapter...since this vehicle is putatively being designed to win the ASP and be used to dock with a Bigelow inflatable have you thought about what kind of docking mechanism is planned for these inflatables?. My guess is that it will be similar to an ISS design, but IMO it's something to look into. BTW, as far as the parachute/parafoil question is concerned...eventually commercial spaceflight is going to be much like current aviation practices where the vehicle lands and then taxis to a terminal where passengers debark through a tunnel and out of the weather. However, to have that as a requirement for the early stages is unnecessary. Getting the G-X3 down with routine safety onto an airfield-sized area seems reasonable for now and I see no reason why parachutes can't do that. Maybe a second generation G-X3 (G-X4?) could incorporate a parafoil. <div class="Discussion_UserSignature"> ------------------------------------------------------------------- "I hold it that a little rebellion now and then is a good thing, and as necessary in the political world as storms in the physical. Unsuccessful rebellions, indeed, generally establish the encroachments on the rights of the people which have produced them. An observation of this truth should render honest republican governors so mild in their punishment of rebellions as not to discourage them too much. It is a medicine necessary for the sound health of government."Thomas Jefferson </div>

mrmorris · Dec 6, 2004

"a Bigelow inflatable have you thought about what kind of docking mechanism is planned for these inflatables?. " Nothing has been put into print that I've ever been able to locate. However, what I *have* seen in print is that Bigelow's modules will require a C&C element and that this is presumed to be a Zvezda-like module to be contracted by Bigelow from the Russians. As such -- it might have either the probe and drogue of the Soyuz/Progress (as does the Zvezda on the ISS) *or* the APAS-89. Seeing as the only docking adapters currently being made are either Russian-manufactured, or Chinese knockoffs of Russian designs -- it's a pretty good bet that Bigelow will be using one of these two. The APAS-89 is by far the more adaptable of the two docking adapters, being genderless as the name suggests. However, my own guess would be that Bigelow would hedge his bets and have both available. After all -- he's interested in the widest compatibility possible. It would be easier to design G-X3 with the probe and drogue. It's my understanding that it's smaller and lighter then the APAS-89. However -- if the deisgn is made from the outset to be able to handle the APAS-89 -- then it's easy to retrofit to P&D. The reverse is not necessarily the case. "Maybe a second generation G-X3 (G-X4?) could incorporate a parafoil. " Which is the point I keep trying to get across. Referring to Atkin's law number 13: Design is based on requirements. There's no justification for designing something one bit "better" than the requirements dictate. On the (hypothetical) day that a Gemini-X3 returns safely to a desert landing using a parachute recovery -- investors and sources of income will spring up out of the woodwork. This happened with SS1 and it doesn't have one tenth the commercial possibilities that the proposed 'simplest'

mrmorris · Dec 6, 2004

"The Gemini docked to the Agena upper stage and that wasn't an APAS-89. APAS-001?" The Gemini-Agena 'docking' has almost nothing in common with what we're discussing. The two vehicles were attached to each other -- period. There was no pressurized means of leaving the Gemini and entering the Agena. It has no relevance to Gemini-X3. It had no name that I'm aware of -- nor is there a reason for one, as it was a one-off with no uses outside the one application for which it was designed. If it were to have a name -- the acronym wouldn't be APAS, as that stands for Androgynous Peripheral Attachment System, and it was not androgynous. "It does work very well, but it is only used for one application, docking to ISS. My point is if there is a reason for GeminiX then maybe a thirty year old design could be upgraded, sort of like using the electronic avionics available today instead of the mechanical systems 30 years ago. " In addition to the erroneous assumption that it's only useful for docking to the ISS that najaB pointed out -- you're also making the erroneous assumption that the design of the APAS-89 is 30 years of (it is not), and that there has been *progress* in the technology of docking systems since it was designed. Unlike avionics -- there is no industry working to improve the technology of docking systems between spacecraft. There *might* be in the future, but until there is, and some time passes, there is no ready-made 'advances' to incorporate into a new docking system. Therefore -- APAS-89 is the newest and best there is, despite being about fifteen years old (designed for Buran for docking to Mir about 1990 -- first used on Soyuz-16 in 1993) "If the only purpose of GeminiX is to dock to the ISS then your right. I wasn't intending to follow NASA's lead anyway. GeminiX doesn't have to be any more NASA comp

mrmorris · Dec 6, 2004

OK -- I really really don't understand why it takes so long for the Orbiters and the Soyuz to make it to the ISS after launch. Looking back through the various shuttle and Soyuz missions -- launch->dock times hang around 43-48 hours. Everything I'm reading about the transfer orbits indicates they should be using Hohmann Transfers /> Time-Of-Flight on a Hohman transfer is: T = pi * (a3 / m) M = 398,600.4 km3/s2 for earth, but I can't find anything that tells me what 'a' is. However -- I did find a page with pre-calculated values for a couple of Hohman transfer orbits in LEO, and it has a 300 – 500 km transfer shown at 0.7706 hrs. This jives with a tutorial I found on another site: "The most common examples of rendezvous manoeuvres are the ferry flights to the Russian MIR space station, and Shuttle missions to recover or repair on-orbit satellites. To simplify our discussion on rendezvous, we will consider two circular coplanar orbits. The target satellite is in the higher orbit. The other satellite, say the Shuttle, is at any random position in the lower orbit following a coplanar launch. Since we are always concerned with fuel efficiency, we will use a Hohmann transfer orbit to transfer from the inner orbit to the outer target orbit. When we reach the outer orbit, we will do the second of our two Hohmann transfer delta-v's to match the orbits. The only real consideration is to arrive at the outer orbit at a position and time when the target satellite is at that same position. Recall that larger orbits have larger periods. Our lower orbiting Shuttle is completing orbits faster than the target satellite. We can use that fact to get in precisely the right position to complete a Hohmann transfer to an exact rendezvous. It is a matter of phasing. To calculate the exact phasing positions, we first consider the Hohmann transfer orbit. Since the initial and final orbits are fixed, the per

najab · Dec 6, 2004

It doesn't have to take that long. At least a couple of the Gemini X1 missions made rendezvous in a couple of orbits. IIRC, the reason that they choose to do a FD3 rendezvous is as much to reduce crew workload as it is for orbital mechanics reasons, though having said that, the delta-v requirement is lower and the launch window more forgiving for multi-orbit rendezvous.

mrmorris · Dec 6, 2004

"...the delta-v requirement is lower and the launch window more forgiving for multi-orbit rendezvous." I can't see the dv requirement changing. Situation 1: The orbiter inserts to 280 km -- waits 25 orbits (~37 hours) then performs the perigee burn, H-Trans out to apogee then performs the apogee burn. Situation 2: The orbiter inserts to 280 km -- waits 1 orbit (~90 minutes) then performs the perigee burn, H-Trans out to apogee then performs the apogee burn. The Hohmann transfer dv requirements are the same. I agree that the launch window is an issue. The insertion accuracy has to be considerably better in the second situation, since there is a smaller margin of time for the angles between ISS and orbiter to match up to 'go' point for the perigee burn. If the orbiter ended up getting too far 'ahead' of the ISS angle -- it would actually have to either wait a *lot* of orbits for the ISS to come around again (orbital periods for 280 km and 350km are 90.11 and 91.54 mins respectively), or actually launch to an orbit slightly *above* the ISS and wait for it to catch up. While the orbiters have little dv to space while carrying ISS equipment, the G-X3 likely will have space capacity. Ergo it can be a bit more flexible in this sense.

najab · Dec 6, 2004

>I can't see the dv requirement changing.Hmm...I guess the dV would be the same either way - but intuatively it seems that a one-orbit rendezvous would require higher performance.

yree · Dec 7, 2004

Big Gemini Big Gemini (or "Big G") was proposed to NASA by McDonnell Douglas in August, 1969, as an advanced version of the Gemini spacecraft system. It was intended to provide large-capacity, all-purpose access to space, including missions that ultimately used Apollo or the Space Shuttle. The study was performed to generate a preliminary definition of a logistic spacecraft derived from Gemini that would be used to resupply an orbiting space station. Land-landing at a preselected site and refurbishment and reuse were design requirements. Two baseline spacecraft were defined: a nine-man minimum modification version of the Gemini B called Min-Mod Big G and a 12-man advanced concept, having the same exterior geometry but with new, state-of-the-art subsystems, called Advanced Big G. Three launch vehicles-Saturn IB, Titan IIIM, and S-IC/S- IVB-were investigated for use with the spacecraft. The Saturn IB was discarded late in the study. The spacecraft consisted of a crew module designed by extending the Gemini B exterior cone to a 419-cm-diameter heat shield and a cargo propulsion module. Recovery of the crew module would be effected by means of a gliding parachute (parawing). The parametric analyses and point design of the parawing were accomplished by Northrop- Ventura Company under a subcontract, and the contents of their final report were incorporated into the document. The landing attenuation of the spacecraft would be accomplished by a skid landing gear extended from the bottom of the crew module, allowing the crew to land in an upright position. The propulsion functions of transfer, rendezvous, attitude control, and retrograde would be performed by a single liquid propellant system, and launch escape would be provided by a large Apollo-type escape tower. In addition to the design analyses, operational support analyses and a program development plan were prepared. Specifications * Crew Size: 9 to 12 * Length: 1

najab · Dec 7, 2004

I believe mrmorris used Big-G as the starting point for GX-3.

mrmorris · Dec 7, 2004

"I believe mrmorris used Big-G as the starting point for GX-3. " Not really. I've looked at it, of course -- but it's actually further from my design goal than the original Gemini. Big-G hugely increased the exterior dimensions of the spacecraft. Essentially it superglued a crew cabin on the back of the re-entry module -- then had a big retro module behind that, then an equipment/RCS/cargo module behind that. As the post indicates -- >15,000 kg and BIG (well... unless you compare it to the shuttle <img src="/images/icons/smile.gif" /> ). I've essentially kept the exterior dimensions of the original 2-person Gemini, eliminated the heavy/volume intensive elements not required for ferry operations (primarily the water tanks and fuel cells), and consolidated everything else from the Equipment module into an expanded Return Module. The result is a spacecraft with exterior dimensions and weight approximately the same as the original Gemini (~4000-4700 kg) but with a crew of five. Since the FalconV rating is currently at 5,540 kg for a 400km orbit, and since the insertion orbit will actually be around 280-300 km so that the capsule can initiate a hohmann transfer orbit to docking -- there's plenty of lift capability to space for increasing the crew size, or adding cargo space. However -- that's a job for another day. Must focus on today's task... five people to orbit -- dock -- undock -- five people to ground level.

mrmorris · Dec 9, 2004

Since Gemini-X3 is to be powered by batteries, many of my 'What-if' scenarios revolve around an emergency that causes them to run low on juice for one reason or another. I've considered adding solar panels to the craft, but deployment and orientation issues simply add more complexity than I really want. However -- as of this morning, I HAVE THE SOLUTION!!! Yes -- I'm proposing a spaceship run by pedal power. Laugh if you will -- but the crew of Apollo 13 would have killed for one of these pedal-generators to supplement the onboard power. It indicates the average continuous power suppliable by pedalling is ~125 watts. That wouldn't be enough to power the craft at full load (although two such generators might come close -- I'll know more after I complete my power usage calculations). However -- in the event of an emergency, it would be enough to run critical systems (i.e. life support, RCS, communications) and have enough spare to recharge the batteries simultaneously. The oxygen storage on G-X3 contains *ample* oxygen for the crew for a flights of ~ten times nominal duration. From earlier calculations, the LOX/Ethanol system would require 5836.8 cm3 of LOX, and the 18" tanks I had estimated to hold 50,040 cm3. I've since refined that -- picking a specific LOX tank from ATK PSI: 80353-1 3,575 in3 (58,583 cm3) 19.03" sphere 8.5 lbs 58,583 cm3 - 5836.8 cm3 (LOX for propulsion) = 52746 cm3 (LOX for ECLSS) Wlox * 1.14 g/cc = 52746 cm3 Wlox = 46268 g (46 kg) Oxygen usage: ~.84 kg/day/person 46 kg / 5 people / .84 kg/day = 10.95 days Eleven days is a heckuva lot more time than G-X3 needs (water becoming critical before then), but the 80353-1 was the smallest cryogenic-rated tank I could find, and I wanted two for r

najab · Dec 9, 2004

>With it -- the options for extending the flight time of the G-X3 in the event of an emergency are greatly enhanced.Given G-X3's intended purpose, in an emergency we should light the retro-rockets and come home.

mrmorris · Dec 9, 2004

"...we should light the retro-rockets and come home. " Granted. However -- What if there is a failure such that the batteries are drained prematurely. The resulting charge is insufficient to run the navigation/communications/computer systems required to orient G-X3 for re-entry. This might be a component failure/short that rapidly drains the battery or a battery cell failure during flight. I'm trying to think 1960's NASAesque here. Primary option, secondary option, emergency backup to the secondary option, failover option for the emergency backup. I don't like having eleven days of oxygen and 2 days of battery power. In any event -- the pedal generator would be a very low volume and mass 'security blanket'. You are correct though -- it's would actually be a much more mission-oriented option on the CEV as opposed to the G-X3. On flights to the moon/Mars the pedal option would not only provide an additional backup power supply, but also valuable physical conditioning for the crew.

najab · Dec 9, 2004

How about a belt of solar cells around the nose of the capsule? That way you wouldn't have to worry about deployment and sun-tracking.

grooble · Dec 9, 2004

http://www.webpal.org/webpal/b_recovery/3_alternate_energy/bike/bike.htm ^ Info about Bike operated power generation. A healthy person can make 150w per 30 mins. But i don't know about zero G sports performance.

mrmorris · Dec 9, 2004

"How about a belt of solar cells around the nose of the capsule? " It's conceivable. You wouldn't get a *lot* of power. Using the entire area of the nose cone, would allow (generously, and a guess at the moment -- my notes are not with me) 12 square meters of surface area if you used the entire nosecone. Working solar power equations on that (fast and dirty), we get: The surface area of the nosecone would be ~12 m2. Of this -- only 50% would be facing the sun at any given time, and of that -- much of it at high incident angles. I'll assume that a maximum of 25% of the sunlit surface area as being effectively facing the sun at a zero incidence angle (when the G-X3's path is oriented at a right angle to the sun). Of course unless the G-X3 spins about its long axis as it orbits the Earth -- it will only have a zero incidence angle for a short period each orbit. Such spinning is unlikely -- therefore the incidence angle will change from zero to 90 and back to zero as it orbits the sunlit side of the Earth (figure 50% production during sunlit portion of orbit). Close to half of its orbit will be in the shadow of the Earth, so production is nothing 50% of the time. Triple-junction solar cells are currently getting about 35% efficiency levels, so our calculation becomes: Esa = solar flux * (surface area * lighted % nc * inc. angle for nc curve * inc. angle for G-X3 path) Esa = 1.37 kW m2 * (12. * .5 * .25 * .5) m2 * .35 Esa = ~360 Watts produced during lighted half of orbit. I don't know what the implications are of the solar cells on the skin of the G-X3 nosecone would be for launch and re-entry. I doubt they have good heating characteristics, though. The pedal-generator is cheap, light, low-tech and requires no vehicle mods. Stick it in the back of the tool cabinet and hope it never needs to be used.

mrmorris · Dec 16, 2004

In looking to build a detailed power and mass estimate for Gemini-X3, I'm searching out current technology to replace the systems that existed on the original spacecraft. Most recently, I found that all of the following equipment from the original: Inertial Measurement Unit -- 130 lbs -- 470 watts Computer -- 59 lbs -- 90 watts Auxiliary Tape Memory -- 26 lbs Structural Supports -- 9 lbs Cold Plates -- 6 lbs Attitude Control Maneuver Electronics -- 40 lbs -- 42 watts Total -- 270 lbs, 602 watts ...could be replaced with a single subcomponent: Honeywell Enhanced Space-Integrated GPS/INS (E-SIGI) -- 21 lbs -- 45 watts This is a reduction of 249 lbs and 557 watts from the original craft. Furthermore, 557 watts is about 2/5 of the power usage of the original Gemini. However -- since the components all required cooling by the ECS -- the specific power usage would have been 50% or more. Essentially that means that the power system requirements are cut in half with this one piece of equipment. The battery packs (easier to get a discrete mass for then the fuel cell system) weighed 708 lbs. The approximate real weight reduction then from this one piece of equipment would be ~603 lbs.

najab · Dec 16, 2004

Way cool! I suspect you will find similar weight/power savings in the communications system. When Gemini was built there were no ICs.

mrmorris · Dec 16, 2004

"I suspect you will find similar weight/power savings in the communications system." They look good too -- but didn't have quite as much fat to cut. I haven't completed my communications research yet. I'm not positive my solution provides everything that G-X3 needs yet. I'm doing some reading on telemetry and satellite communications. None of the COTS systems for satellite comm systems are geared towards voice (duh -- there's no manned spacecraft for them to be put on). The Kenwood transceiver I'm referencing for VHF/UHF electrical/weight specs I picked because one was delivered to the ISS a few years back so they could communicate with Ham radio operators. Ergo I know it has the range required. Unfortunately -- since it won't have any encryption/etc. -- it likely won't be useable OTS. Original Gemini: Near Earth systems -- 54.54 kg 100 watts Intercom -- 1.69 kg -- 9 watts Recovery -- 9.8 kg -- 44.5 watts Data Processing -- 3 kg -- 25 watts Instrumentation -- 41.43 kg -- 51 watts Wiring/ECS/Structure -- 134.53 kg Total: 244.99 kg -- 229.5 watts Gemini-X3: GD TDRSS 4 Transponder (S/Ku) -- 9 kg -- 45 watts Kenwood TM-D700E VHF/UHF (Voice or Data) -- 2.5 kg -- 35 watts Wiring/ECS/Structure -- 25 kg Total: 36.5 kg -- 80 watts

najab · Dec 16, 2004

Would encryption really be a requirement? I think it would be 'cool' to have hams all over listening in on our missions.

mrmorris · Dec 16, 2004

I don't know if encryption is a requirement. Much as I'm taking on in designing this -- I haven't tricked myself into thinking that I'm the one deciding on what the final requirements are. Some are obvious -- others can be assumed, and some just have to be guessed at. I'm more concerned about civilians clogging the airwaves during a mission. If the transmissions are encrypted both ways -- this possibility is reduced.

mikejz · Dec 16, 2004

From LEO you only need a couple of watts for a decent UHF link. The data they are talking about is only about 9600 baud ax.25 based however. Encryption might not be the solution, spread-spectrum would be for limiting anyone from calling up to them!.

mrmorris · Dec 16, 2004

"From LEO you only need a couple of watts for a decent UHF link. " Ayup. I listed max wattage. Reception only takes 2W. The remainder is only during transmission. "The data they are talking about is only about 9600 baud ax.25 based however. " The UHF/VHF is 99% for voice. Telemetry will go through the other transceiver. "spread-spectrum would be for limiting anyone from calling up to them!. " I'm easy -- know of a VHF/UHF transceiver that'll handle this?

mikejz · Dec 17, 2004

Here is a press relase i found on Lox based RCS Aerojet Tests LOX Ethanol Reaction Control Engine SACRAMENTO, Calif., Oct. 22, 2004 -- Aerojet, a GenCorp Inc. (NYSE: GY) company, recently conducted a successful hot fire test program for a non-toxic reaction control engine utilizing liquid oxygen (LOX) and ethanol as propellants. The reaction control engine (RCE) is being developed for NASA’s Marshall Space Flight Center and the Exploration Systems Mission Directorate as a relevant technology that is applicable to the Vision for Space Exploration. The engine testing validated the operability of a cryogenic LOX-ethanol reaction control engine and ignition system in both steady state and pulse mode operation. It included long duration steady state operations up to 240 seconds and numerous pulse mode operations typical of ascent, on-orbit and re-entry operations. The testing also included the evaluation of the effects of chamber pressure and mixture ratio variations on the engine, combustion chamber thermal response, and six different pulse duty cycles between 80 milliseconds and ten seconds. Chamber temperatures were as predicted and uniform in steady state and during pulsing, validating Aerojet’s capabilities in controlling the injection and combustion process through the use of proprietary platelet injector design. “With this testing, Aerojet again demonstrated the value of our unique platelet technology,” said Aerojet RCE Program Manager Eric Veith. “The platelet technology distributes the LOX and ethanol propellants in a more uniform manner, thus producing an even mixture ratio, which provides efficient and uniform combustion.” Pulse mode testing was conducted to demonstrate the operability requirements for minimum pulse widths for a liquid-liquid auxiliary propulsion system that used LOX Ethanol. Aerojet worked closely with its suppliers, Moog and Castor Engineering, to develop LOX & Ethanol main injector and igni

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