Rocket engine with electric pumps

[Enthalpy]

Hello dear friends!

 

As an alternative to heavy pressurized tanks and to complicated turbopumps, an electric pump can feed the propellants in the chamber(s). Conceivable at small chemical thrusters, where a high pressure improves the efficiency and injects enthalpy from the Solar panels.

-----

Scale up: electronics can control 1MW motors from the main's voltage. A 70% efficient centrifugal pump brings then 72kg/s of oxygen and farnesane C₁₅H₃₂ to 96b; expansion from 80b to 0.02b in a 2.6m nozzle to push 260kN with Isp=375s, enough for the main engine of a 30t upper stage.

Rotating at 780Hz in vacuum, the motor is small. Its rotor can be a permanent magnet of Magnetoflex 93, D=250mm L=200mm. The D=350mm 3-phase 8-pole stator loses <1kW in its braided "Litz" wires and 70W in the Nanoperm magnetic circuit; it's cooled by the fuel. The motor weighs 140kg. It accelerates in 30s.

The safer Li-MnO2 primary battery brings ~650kJ/kg in the too quick discharge. The stage bringing 5900m/s would burn 24t; at full thrust, this would require a 470kg battery (20kg per ton of propellants), but throttling is easy, useful, and lightens the battery.

The battery is easy to integrate and welcome for the gimbal actuators, the igniters...

-----

Pressurized steel tanks for the same 24t but with only 36b in the chamber, throttling to 20b, would weigh 1930kg and provide 13s less Isp. Graphite tanks could weigh about 1100kg.

A gas generator cycle could waste 42kg per ton of propellants, later ejected at 1500m/s thus counting as 35kg/t - but this equivalent overhead is ejected all the way and a battery supposedly not.

Staged combustion is more efficient.

-----

Scale up further: power components control 6MW railway engines. This would provide 1MN thrust with 120b in the chamber, enough for 115t at a second stage or 70t per 6MW slice at a first stage.

-----

Smaller roll and vernier engines can also have electric pumps, for instance at a solid engine stage.

For a lander or descent-ascent module, I like the ease of starting and restarting electric pumps.

-----

Fuel cells would have been fantastic for a hydrogen-oxygen engine with electric pumps, but are still too heavy.

Marc Schaefer, aka Enthalpy

 

=====================================================================================

At an RL-10B equivalent with electric pumps, injectors shall drop 17% pressure and pumps be 65% efficient, then the shafts need 145kW and 399kW, and the motors cumulate 80kg. The Li-MnO₂primary battery weighs 36kg per ton of propellants, a bit less by throttling. This is mass where not desired, and an expansion cycle performs better, but the electric pump is simpler.

 

The battery mass is a good surprise, this is how I compute it:

 

1kg of oxygen-hydrogen m5.88:1 occupies 2.8dm³ as liquids boiling at 1 atm. Because no turbine is needed and I forgot the drop in the cooling jacket, the pumps must bring the propellants to 52.3b at the injectors. The power electronics shall be 97% efficient, the motor 99.9%, the pumps 65% because of hydrogen. 23kJ of electricity is used per kg of propellants.

Li-MnO₂ batteries use to store 240Wh/kg = 864kJ/kg or 33Ah*2,7V/355g = 904kJ/kg at room temperature
http://media.duracell.com/media/en-US/pdf/gtcl/Technical_Bulletins/Lithium%20Technical%20Bulletin.pdf
http://www.saftbatteries.com/doc/Documents/primary/Cube656/FRI_M62.2cf814ca-1181-4dc6-85f1-7c679f35054c.pdf

but are meant for multi-hour discharge. Winding the electrodes and electrolyte thinner must permit a faster discharge (and self-discharge); I take only 650kJ/kg because of the redesign. Hence the 36kg per ton of propellants.

This extra mass is similar to a gas generator cycle because both energy sources are chemical, and both happen to be equally inefficient. 36kg/t is as much as the Shuttle's external tank, alas, so a lower pressure is
probably better.

Throttling permits to reduce the pump pressure at the end, saving energy.

Li-SOCl₂ and others are lighter, fit fast discharge, cold... but they catch fire, explode or emit toxic gas if crushed or pierced. I don't want half a ton of them over 20t of explosive gas.

-----

The power electronics needs a special design cooled by the fuel. An eight-pole motor at 780 Hz (46,800 rpm) needs three-phase at 3120 Hz so the electronics makes probably a 1/3 - 2/3 waveform instead of a sine by pwm.

The motor is exotic. Its permanent magnet rotor is uncooled; Magnetoflex 93 with HV=950 rotates at 613m/s, key to the tiny design. A stepper motor would accept more banal materials but demands vacuum, and might combine with the impeller - maybe perhaps.

 

=====================================================================================

 

Many hydrogen engines for upper stages offer no roll control. Vernier thrusters control the roll, the launcher's attitude at spacecraft separation, the precise injection speed to orbit.

Instead of toxic and less efficient hydrazine, these thrusters could use the hydrogen and oxygen, but they must operate after the main engine shuts off, and a restartable extra turbopump is too complicated for them. Propellants at ~1.5b from the tanks may cavitate and need bigger thrusters. Electric pumps would improve this.

This holds for other propellants as well, but pressure vessels are especially bad for hydrogen, hence the pumps.

As roll actuators, they could push only when needed, saving further electricity hence battery mass.

A positive displacement pump can be considered here.

A thruster taking 5b hydrogen-oxygen exists already at DLR. For this purpose?

Marc Schaefer, aka Enthalpy

 

=====================================================================================

 

Many communication satellites are put by the launcher on Geosynchronous Tranfer Orbit and reach Geo-Synchronous Orbit with a 1600m/s kick by their MMH+N2O4 apogee motor. While better than solids, these motors use inefficient propellants pressure-fed by heavy tanks to burn at a low pressure. Electric pumps would improve.

The satellite shall weigh 4000kg including the engine - wherever this one is. The nozzle is 0.6m wide. 1600m/s are brought in 2000s, needing about 4kN.

 

 

Hydrogen-oxygen in the satellite improves most. The satellite's existing battery loaded by Sunlight brings all electricity (75MJ) for 100b in the chamber, hence the Isp and inert mass. Though, the hydrogen takes a 1.7m sphere. The sphere can be multilayer-insulated and held to a truss by polymer straps.

Hydrogen-oxygen in a launcher's special stage takes an extra battery of 16kg/t to achieve only 20b. No big improvement over a Vinci stage for instance, but some launchers (Zenit, Falcon...) could then offer GSO delivery to satellites without an apogee motor or to replace MHM.

Syntin-oxygen in the satellite can burn at 250b, or even more if the apogee kick is spread over several orbits. A strained amine is as efficient as Syntin and possibly cheaper.

Syntin-oxygen in a launcher's special stage burns at good 80b. All-kerosene launchers may prefer it. The gain over MMH is still impressive.

Space probes have similar needs, for instance to get captured by Saturn. A cryocooler can keep the propellants liquid.

Un pensamiento para Don Hugo.
Marc Schaefer, aka Enthalpy

 

=====================================================================================

An oxygen screw pump fits a 4kN apogee motor burning Syntin.
http://en.wikipedia.org/wiki/Rotary_screw_compressor

Its two rotors can have (before optimization) D=20mm and a core of d=16mm, a single thread of 10mm pitch and 50% solid, for instance of nearly-sine profile. Drive at 418Hz and 3.3N*m (8.6kW) achieves 96bar and 669cm3/s if efficiency is 75%. With 10 screw turns in 100mm length, oxygen inertia lets it leak at 41m/s; 20µm radial clearance limit the leak to 104cm3/s - less if the pump is smaller. The stator and rotors need roughly matched temperatures and expansion coefficients. Turning or milling tools are made to the desired tiny profile. The pump weighs about 3kg.

The electric motor can have a D=40mm L=60mm rotor (only 53m/s) with four poles of Nd-magnets held by a steel sleeve. The three-phase stator then looses about 60W. The motor weighs 2.5kg.

Example of an electrically pumped apogee stage:

 

 

Marc Schaefer, aka Enthalpy

 

=====================================================================================

A main hydrogen-oxygen stage pushing 1MN with 40bar in the chamber(s) is feasible with battery power.

The impeller for the hydrogen centrifugal pump for 33kg/s or 467dm³/s at 48bar has for instance D=214mm (not optimized), its channels are 19mm high, and it rotates at 503Hz. If 60% efficient, it needs 3.9MW and 1.2kN*m at the shaft. I won't detail the booster pump.

The electric motor can have 10mm thick Nd-Fe-B magnets moving at 200m/s at the D=126mm L=400mm rotor, held by a 3mm thick sleeve. The 3-phase 4-pole (or more) stator loses about 20kW in its coils with few turns of rectangular wire, and the motor weighs 200kg.

The traditional sleeve is cold-rolled austenitic steel. It could instead be helix-rolled and welded Maraging sheet to offer permeability. But for precision, and to reduce eddy currents in the sleeve, I'd prefer the same 3mm wound of unidimensional graphite prepreg. Thin elastic material can fit between the steel core and the magnets.

A stepper motor, with a D=400mm L=50mm rotor and overlapped phases and coils, may work and weigh 70kg, but I won't invest more time to check this attempt.

The oxygen pump can have symmetric inlets, then D=174mm lets it rotate at 165Hz, so a ring motor can have D~380mm with many poles and be flat - I won't detail it.

The magnets cost 1000€ per motor. 80t of propellants need 2.7t of Li-MnO₂ batteries, which sell for ~50€/kg in small amount.

Marc Schaefer, aka Enthalpy

 

=====================================================================================

 

A booster burning 225:100 of O₂:Pmdeta expanded from 110b to 0.35b in a D=2.4m nozzle to produce 2MN and 337s @vac needs 419kg/s and 367dm³/s of oxygen. The centrifugal pump with symmetric inlets can rotate at 217Hz with a D=204mm impeller; being 72% efficient, it receives 6.75MW and 4.95kN*m from the shaft.

 

 

(Click to magnify) A corresponding electric motor has 10mm thick Nd-Fe-B magnets moving at 200m/s on D=290mm. 3mm thick wound unidimensional graphite composite holds the magnets. 18 poles allow 16mm thin iron at the stator and the hollow rotor, and a single turn (36 passes, 36 grooves, 3 phases) of D=8mm twisted "Litz" wire gets 1.7kVpk induced, while 2.0kApk take 12 big IGBT.

Copper looses some 50kW and the motor weighs 160kg for almost 7MW: it's lighter than a gas turbine. On an aeroplane, it can drive a fan directly, or a propeller through a gear. We still lack light fuel cells.

Marc Schaefer, aka Enthalpy

Edited March 12, 2013 by Enthalpy

[CaptainPanic]

A few comments:

 

Why do you propose solar panels at the top of your post? You need lots of power for a relatively short period of time, so I would go with a battery or even a (super?)capacitor.

 

I am confused how you arrive at all the values for the pumps / compressors. How do you calculate the weight of the pump? How do you calculate / find the pressure that the pump can generate? Care to either explain, or give a link (reference)?

 

Without any references, it is difficult to check, but you seem to sound pretty optimistic.

[Enthalpy]

Thanks for your interest! And please forgive the late answer, I had no electricity for two days.

 

You're absolutely right that the pumps need much power over a short time, so they must be fed by some chemical means. The secondary battery already in the satellite and full happens to suffice, which is excellent news, as this battery can be later re-charged by the Solar panels, so electricity for the pumps needs no extra mass.

 

That's why I compute with a higher chamber pressure if the apogee rocket engine is part of the satellite rather than the launcher: enough energy is available at no extra mass, and pressure improves performance.

 

Supercapacitors are still heavier than chemical batteries. Secondary (=rechargeable) batteries can fit quick discharge, but use to weigh more than primary ones do - I didn't check that with recent technologies. And I agree that the Li-MnO₂ I've seen can't be discharged within few minutes; I just suppose that thinner electrodes and electrolyte enable that - take a different technology if not.

 

-----

 

I extrapolate centrifugal pumps from RD-170's liquid oxygen ones.

http://www.lpre.de/energomash/RD-170/index.htm (section Турбонасосный агрегат)

 

The outlet pressure varies as density*square(linear speed). I keep the diameter/(blade height) ratio and make this area*speed proportional to the volume throughput, where I keep the fluid's radial speed proportional to the impeller's tangential speed.

 

In some cases, the pump is symmetric, to double the relative blade height, reduce the impeller's diameter and increase the angular velocity, in favour of the electric motor.

 

This design could be improved, especially the fluid's radial speed, but at least the direct extrapolation looks safe to my eyes. Or isn't it? Maybe I should not extrapolate to hydrogen.

 

As the electric motor can run at 200m/s and is lighter then, fast impellers at oxygen and dense fuel would improve, which suggests some kind of axial pump instead, like at a booster pump, but I suppose cavitation is very difficult to avoid at the inlet then.

 

-----

 

I give a mass for the screw pump.

http://en.wikipedia.org/wiki/Rotary_screw_compressor

 

The one I consider runs really fast at 418Hz as compared with 60Hz or 25Hz usually, and that makes it lighter. I consider it possible because it has no seals at the rotors, only at the low pressure side of the shafts. The throughput results from volume*frequency, and the mass from the size. 20µm radial clearance is easily achieved by machining (details later); because oxygen is very thin, leak speed is limited by inertia, or (10 steps) 9.6bar = 0.5*d*V² with d=1141kg/m³. The leak isn't brilliant, but a thread profile more hollow would reduce the leaking area, and more turns would reduce the pressure per step.

 

Centering the rotors needs adjustments, like eccentric parts to hold the shafts. Thermal expansion isn't critical: less than 24µm over the radius for steel and 200K, before stator-rotor matching. Put the pump within its oxygen inlet line?

Edited March 15, 2013 by Enthalpy

[Enthalpy]

Researchers in Buenos Aires and Roma have already studied electric feed for propellants:
http://www.dima.uniroma1.it/STAFF2/jpp12r3.pdf
They evaluate masses identical to mines, within computing noise.

- They take rechargeable lithium technology and properly observe the peak power of existing cells.
- I carefully match the rotation speeds of the pumps and the specially designed motors.
- The dear reader makes the smart and astute synthesis.

My strong feeling is that electric propellant feeding will exist on apogee stages or satellites, because it's easily developed, and among the possible solutions (as a turbine would be too small), it brings a big performance improvement. At lower stages, it's rather a good, simple and cheap solution, which brings a launcher early to the market, before a more efficient turbine cycle possibly replaces it.

Marc Schaefer, aka Enthalpy

[Enthalpy]

For a small propellant throughput, say at an apogee stage, a screw pump is more practical than a centrifugal one.

An accurate screw profile limits leaks. For a small screw like D=20mm, a (pair of?) accurate tailor-made tool can cut the profile, just like a fastening screw is cut:

 

 

sketched here for turning; milling is possible. Notice the tool cuts the outer diameter as well, so the whole profile is precise and easy to measure. The stator and rotors will likely receive some protective layer against oxygen and wear; if thin Ptfe-impregnated nickel is acceptable, its thickness can finely adjust the parts' diameter. Or etch the parts (electro-) chemically for the last few µm.

Electric Discharge Machining is one alternative here to chip-cutting.

The screws' outer edge can comprise several crests instead of one, separated by grooves, to reduce the leak. The whole profile must be designed accordingly. Pump specialists may know more tricks, perhaps even seals.

A thick stator deforms very little at 300b.

Marc Schaefer, aka Enthalpy

Edited March 16, 2013 by Enthalpy

[Enthalpy]

To illustrate electric pumps, this launcher puts 3.0 to 8.3t in geosynchronous orbit directly. It can also send 9.0t or 2.8t towards Mars and Saturn directly, or without the upper stage, put 8.2 to 25t in low-Earth-orbit.

The upper stage burns 12.1t of hydrogen-oxygen at 30bar, expands in D=2.15m to achieve 75kN and Isp=4565m/s. Balloon tanks within a taller fairing would gain almost 300kg payload.

The lower central stage burns 54t of hydrogen-oxygen at 40bar, expands in D=3.0m to achieve 650kN and Isp=4352m/s. 4 Vinci or 6 RL-10 chambers, with short nozzles, can replace the single one, and share a set of actuators and electric pumps, later maybe turbopumps.

Each booster burns 52t of Pmdeta-oxygen at 110bar, expands to 0.35bar in D=2.0m to achieve Isp=3304m/s in vacuum and, at sea level, 1.2MN and Isp=2594m/s. 2, 3, 4, 6 boosters, or 6 followed by 2, match varied payloads. The launcher's strongest pump takes 5.0MW shaft power.

The boosters connect at the reinforced upper section of the lower central stage. Boosters share a fastening with each neighbour at the core. The core of extruded aluminium profile spreads the push and pull forces by shear; a small internal truss prevents bending.

All skins consist of extruded profiles as I describe elsewhere. Batteries are still Li-MnO₂ despite the uncertain peak power, but that's a matter of 30kg/t weighing 1/3 more. All empty stages weigh less than 102kg per ton of propellants. Here all tanks start full even with two boosters, but bigger tanks would widen the adaptation range.

Marc Schaefer, aka Enthalpy

[Enthalpy]

To my surprise, rechargeable batteries are as light as primary ones, and would ease trials of the engines. Li-ion can fit the long burn of an apogee stage, perhaps a perigee stage, but seemingly not a booster, and already caught fire on aeroplanes, cars, laptops.

Li-polymer exist for 4min discharge (15*C) - fits a booster. Polyquest-Enerland developed one series, but does A123 Systems still sell them?
http://enerland.en.ec21.com/ "RC series"
Toshiba developed an other series for <5min operation including at cold
http://www.toshiba.co.jp/about/press/2007_12/1101/SCiB.pdf
Altair Nano makes something, but what?
http://www.altairnano.com/

Fortunately, hobbyists give more data about batteries and fast discharge
http://www.aircraft-world.com/prod_datasheets/polyquestxp.htm
http://www.conrad.de/ce/de/product/238913/Top-Fuel-LiPo-Akku-222-V-2500-mAh-40-C-
Li-polymer meant for permanent 40*C, or discharge in 90s!
http://www.conrad.de/ce/de/product/269802/Graupner-LiFePO-4-Akkupack-165-V-4000-mAh-35-C-Stecksystem-G35-EH
Li-FePO₄ meant for permanent 25*C, or discharge in 144s!

From data at Conrad, Li-polymer would bring 455kJ/kg at a reasonable pace; if snap discharge drops the mean cell voltage to 3.45V, the battery would still bring 424kJ/kg. Data for Polyquest batteries at aircraft-world.com even tells 509kJ/kg slow, 475kJ/kg fast.

The safe Li-FePO₄ (through Conrad) still brings 371kJ/kg at moderate pace.

 

As a complete launcher would use >10 tons of batteries, these can be tailor-made, say from prototypes a manufacturer didn't market. Here we can accept fewer cycles and a faster self-discharge, which uses to enable a faster discharge as well.

And for sure, the pumps, electric motors and inverters would be new designs, saving much mass over the cited research paper's figures.

-----

Li-S and Na-S work only hot - I dislike that... I'd even prefer some Li-O₂. Inject the product in the main chamber.

-----

How difficult would be a light fuel cell? It needs to operate for few minutes, can work at 600°C or 1400°C, get hydrogen at 100bar... At an apogee stage, providing only 50kW in 100kg, it still outperforms a lithium battery, and car prototypes have already such fuel cells.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Launching from Baikonur, Zenit achieves an orbit inclined 51.4°; a third stage brings to 1500m/s below geosynchronous orbit the satellite, leaving the same effort as from an equatorial transfer orbit.

I suggest here a smaller third stage, as it starts from low-Earth-orbit already and attains the geosynchronous orbit directly. Its electric pump fits the many ignitions (~5 burns at perigee, 2 at apogee) and small thrust (10kN) that, with 60 bar in the chambers and four D=1.0m nozzles, extract Isp=3985m/s=406s from RG-1 and oxygen .

 

Zenit-2 places 12940kg on the inclined low orbit. Correcting the inclination then lets the apogee cost 2500m/s instead of 1600m/s, while the perigee still takes 2500m/s, so the 8775kg propellants leave 3690kg.

 

 

A truss of welded aluminium tube, 330mm wide and 2mm thick, holds the payload and the engines to the previous stage and contains the fuel. Polymer belts hold the oxygen balloon tank, made of brazed 150µm thin steel covered with foam and multi-layer insulation, to the truss; electric motors compensate the thermal expansion. The tanks with belts and insulation weigh 236kg, or 27kg/t of propellants.

The 10kN four-chamber engine, with one set of pumps and actuators, plus verniers and their bladder accumulators, shall weigh 200kg. Existing Lithium-polymer accumulators weigh 142kg, the inverter 8kg, summing 350kg for the pumped engine.

Sensors, radiocomms, control and their energy are to fit in 100kg. Three halves of separation actuators weigh 75kg. Sum 129kg unaccounted items, and the stage puts 2800kg in GSO . This is more than the DM-SLB followed by a pressure-fed apogee engine at the satellite, and it saves one stage.

This design adapts to Falcon and others, and can be considered for hydrogen.

Marc Schaefer, aka Enthalpy

[Enthalpy]

This design (click to enlarge) with electric pumps everywhere puts for instance a Soyuz spaceship in orbit, but is smaller and lighter (201t) than the venerable Soyuz launcher (315t).

 

 

The 4m body accommodates efficient 2.6m nozzles that deploy like Vinci does; 140 bar, 140 bar and 100 bar make the rest of the Isp. Rolls near the Vernier thrusters guide the previous stage during separation, as on Zenit's Dm-slb, to prevent collisions.

The whole cylindrical body, plus the payload adapters, the third stage's oxygen tank and possibly the fairing, consist of assembled aluminium extrusions as I suggest in

http://www.scienceforums.net/topic/60359-extruded-rocket-structure/

here of AA6005 with t₁ = t₂ = 1mm, a = 45°, B =22mm, oriented axially, to resist 8MN compression and 8MN*m bending moment, twice the expected stress. Mastering thinner extrusion of stronger weldable alloys (AA7022) would improve, as would magnesium alloys.

Cylindrical tanks are wrapped tightly in welded bands of AA6082, 0.3mm at the top, up to 3.0mm at the bottom of the first stage's oxygen. Tank heads are of AA7022, as is the second kerosene tank and its holding cones, as well as all trusses - all welded together. At the third stage, the structural fuel tank is a prismatic toroid of the aluminium extrusion, while the balloon oxygen tank is of 250µm brazed steel, plus foam, hold to the toroid by polymer belts.

Per ton of combined propellants, the structural sets of tanks weigh 34.1kg, 28.9kg and 28.1kg. Li-polymer batteries weigh 49.2kg, 49.2kg and 35.1kg per ton (throttling unaccounted), but the dry stages keep at 121kg, 109kg and 108kg per ton.

The first oxygen pump takes 15MW at 13800rpm; its motor can be 360mm long, 530mm wide and hollow, to weigh 380kg.

A two-stage design would weigh 330t, with poor capability for slightly higher orbits.

An optional fourth stage, scaled from my description for Zenit, or from the present third stage, brings payloads to geosynchronous orbit from Baikonur or the Equator, or towards Mars... A 2.6m nozzle would improve further, as would strained fuels and the smaller thrust.

Marc Schaefer, aka Enthalpy

[Enthalpy]

An efficient upper stage for the KSLV is sooner developed if electrically pumped. To achieve 202kN and isp=369s, it burns RG-1 and oxygen at 140bar and expands in a D=1.45m nozzle; rolls at the vernier thrusters guide the interstage during separation. The stage begins with 22.6t and ends with 5.5t.

 

 

Axially oriented AA6005 extrusion, with t₁=t₂=1mm, a=45°, B=22mm, as described there
http://www.scienceforums.net/topic/60359-extruded-rocket-structure/
makes the outer cylinder, interstage, payload adapter and maybe the fairing. The rest is AA7022 sheet and tube; all is welded together.

The RG-1 tank is a D=2.4m d=1.7m t=1mm ellipsoid hold by two 45° t=0.5mm cones. The oxygen tank is hold by two t=0.6mm cones; it's a D=2.8m t=1mm sphere covered with 10mm foam and multilayer insulation. The sphere's diameter and height shrink at cold, and the cones less so; their best angle is near 45° to reach the cylinder without stress.

The tanks and cylinder weigh 490kg, the payload adapter 65kg, the engine with pumps and truss 375kg, the Li-polymer batteries 850kg, sensors and control 200kg, unaccounted items 100kg. The earlier separated interstage and fairing weigh 330kg and 570kg. The empty stage weighs around 2074kg, leaving 3450kg payload at 4400+5100m/s performance.

74% efficient centrifugal pumps need 836kW and 376kW shaft power. The oxygen impeller can be 85mm wide and 8mm high running at 586Hz, its electric motor 152mm wide and its poles 100mm long.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Fuels cells limit the power more than the energy. In the following perigee-apogee stage that accelerates only 1m/s², a 100kW Honda cell for marketed cars is to weigh <100kg
http://en.wikipedia.org/wiki/FCX_Clarity
http://www.ansoft.com/workshops/altpoweree/Andy_Bosco_GM.pdf
and it replaces 598kg of Li-polymer batteries with 21kg of H₂ and O₂ ejected continuously.

The stage starts with 15t on a 200km 30° orbit, as for instance the H-IIA achieves from Tanegashima, and provides the following 4330m/s. 15kN thrust need many perigee and apogee kicks but let 100kW achieve 63bar in the chamber (mass 750:100, H₂ pump 55%, O₂ pump 70%, injectors 1.2), and four 1.0m nozzles combine to isp=490s including the 236pppm used in the cell.

The ellipsoidal H₂ tank weighs 64kg with 120µm brazed steel, foam, multilayer insulation and the holding polymer belts. The prismatic toroidal O₂ tank weighs 198kg of already described AA6005 extrusion, the payload adapter 66kg of same material. The four-chamber engine shall weigh 200kg, the sensors and control 200kg, unaccounted items 100kg, for a 828kg dry stage.

The longer interstage of AA6005 extrusion is 349kg heavier, the fuel takes 8706kg, leaving 5.1t payload in GSO.

One added fuel cell for more pressure wouldn't pay for its mass. A single deployable nozzle can replace the four but needs verniers. The H-IIA would better put a lighter stage on a higher orbit. The hydrogen throughput is little for a centrifugal pump but much for a screw one; maybe a single axial pump, like at booster pumps but faster.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Antares has flown for the first time, congratulations!

More capacity would result from a liquid second (and optionally third) stage, with electric pumps easing their development, possibly as a step towards turbopumps.

 

 

The first stage is kept; side boosters as at Ariane 4 would permit to fill it despite the heavier second stage.

The second stage has a common tank head of 3mm plain AA6005 sheet. Four chambers save length hence mass. The rest resembles previous designs. 140b and 1.2m nozzles achieve 400kN and isp=376s . 33.5t propellants and 3.3t dry stage put 8.6t in Leo (=9500m/s). A future turbopump would save 1650kg batteries and add little mass.

The optional third stage provides 4520m/s to joint Gso from a Wallops Island Leo. Its structure is the conical payload adapter of AA6005 extrusion (t₁=t₂=1mm, a=45°, B=22mm) and a truss of AA6082 tubes, D=70mm t=1mm, to the engine. RP-1 fits in a 0.5mm thin half-cylindrical alu torus welded to the cone. O₂ has an ellipsoidal tank of 150µm brazed steel, plus foam and multilayer insulation, hold by polymer straps. 50b and four 0.8m nozzles achieve 8.5kN (1m/s², successive kicks) and isp=401s . A turbopump would hardly challenge the 81kg batteries. 6.0t propellants and 826kg dry stage leave 1.9t in Gso.

Marc Schaefer, aka Enthalpy

[krash661]

hmmmm.

your name is familiar to me,

I can not place it.

[Enthalpy]

Vega puts 2300kg in Low-Earth-Orbit with three solid and one toxic stages. Accumulators and electric pumps do it with two stages.

 

 

The first stage burns 61.5t of Pmdeta and O₂ at 100bar and expands to 0.30bar in four D=1m nozzles to achieve 1.23MN & 271s @sl, 1.53MN & 338s @vac. It weighs 6.1t empty including 2.2t accumulators, and brings 4237m/s.

The second burns 11.9t at 60bar and expands to 4.6kPa in four D=0.6m nozzles to achieve 358s and excessive 215kN. It weighs 1.43t empty including 0.25t accumulators, and brings 4237m/s.

Extruded AA6005 panels make the cylindrical skin, with reinforcing bands welded at the tanks' lower end. The heads are of AA6082; intermediate ones are thicker to resist buckling at 0.5bar difference, and can bear a balsa insulation instead of foam there.

-----

An optional third stage starts with 2282kg at Leo to reach the geosynchronous orbit (4328m/s from 30° inclination) or a Mars transfer (3800m/s). It burns 1493kg at 40bar and expands to 62Pa in four D=0.6m nozzles to achieve 405s and 3500N, so the fourth kick from geostationary transfer orbit to Mars transfer lasts 390s.

Both tanks are of brazed 100µm steel hold by polymer bands, with foam and multilayer insulation at oxygen. They cumulate 16kg. The truss of D=60mm e=1mm AA7020 tubes that holds them and the payload weighs 42kg. The chambers, nozzles and pumps add 70kg, the accumulators 21kg. Sensors, datacomms, control are heavy, so the empty stage is evaluated at 264kg.

-----

Electric pumps ease an affordable Mars access , so this upper stage is also a migration path towards more liquid stages at Vega, while lateral Zefiro-23 can adjust the performance:

 

 

Marc Schaefer, aka Enthalpy

[EdEarl]

You post is too much for me, being a bit dyslexic. I read the first bit about electric pumps. I am not a rocket scientist, but I expect that a turbine would be lighter weight and more capable of withstanding high-g and extreme temperatures than an electric motor. Thus, I suspect pumps that inject liquid fuel and oxidizer into a rocket engine are not electric. Maybe someone here is a rocket scientist and knows

[Enthalpy]

If the propellants pumps were already electric, I wouldn't make a long description nor append my name.

Extreme temperatures are not necessary at the electric motor.

A quick electric motor is always lighter than a turbine; the accumulator is not, and I provide estimates showing that the extra mass is manageable.

[EdEarl]

I had in mind something more like an automobile turbocharger than an air breathing microturbine. A small amount of pressurized fuel and oxidizer could start it, and pressurized gas from the main engines could run it after being started. But, you know more than I.

[arc]

This is definitely the future in rocket motor design. I was struck by the thought that in years past technology from the aerospace industry took many years to trickle down to lower level applications in light aircraft and automotive systems. You may be now applying technology that is currently in the most advanced electric cars. I can't even imagine what the technological possibilities of your concept would be if advanced in the traditional research and development of a space program. That would surly benefit the electric auto manufacturers need for advances in energy storage and delivery. arc

[Mr Monkeybat]

Fly wheels would seem to be an appropriate form of energy storage for this application.

[Enthalpy]

The safe Li-FePO₄ brings 371kJ/kg. A flywheel of good steel (400m/s toroid) stores 80kJ/kg, of graphite-epoxy (800m/s) 320kJ/kg - in ideal cases, and exploiting the energy down to zero speed.

[Enthalpy]

The outline of the future Ariane 6 has been chosen recently, with two solid stages and one hydrogen: explosive, highly flammable, polluting, and if hydrazine provides the roll control, carcinogenic and toxic.

Criticism is easy, art is difficult, so here's an alternative with oxygen and Pmdeta. Click the sketch to magnify

• The last stage has electric pumps as described here previously.
• The two first "sailback" stages are pressure-fed and reused as I describe there
  http://saposjoint.net/Forum/viewtopic.php?f=66&t=2554

 

 

Liquids lighten the launcher, even without hydrogen, and with pumps only at the small upper engine. To put 4.1t on geosynchronous orbit (or 6.8t on transfer) from the Equator:

• The last stage begins 2000m/s before Low-Earth-Orbit with 26.4t composite, burns 20.8t (or 18.1t) at 60bar expanded to 335Pa for isp=394s, ending 3953m/s (or 2457m/s) above Leo. 1560kg dry mass include batteries for the pumps, balloon tanks, an aluminium truss, sensors, control, transmissions.
• The second stage provides 4100m/s by burning 88.5t at 36bar (throttled to 18bar at the end) expanded to 0.081bar for isp=341s. The composite begins with 125.3t; stage dry mass is 9.9t with steel tanks, helium, engines, paraglider and the rest.
• The first stage provides 3400m/s by burning 398t at 36bar (throttled to 18bar) expanded to 0.47bar; isp=303s in vacuum and 254s at sea level. The composite begins with 569t; stage dry mass is 44.6t.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Esc-A, Ariane V's upper stage, has too much dry mass; now Esc-B is rumoured at 5650kg - shattering 200kg per ton of propellants. This hampers missions to geosynchronous orbit (Gso), one goal of Esc-B, and precludes missions beyond Mars. Worse: Ariane 6's upper stage inherits from the Esc-B, but dry mass hampers more a smaller launcher, and claims for Gso performance have disappeared.

Maybe the extruded skin brings something, or maybe not:
http://www.scienceforums.net/topic/60359-extruded-rocket-structure/
Also, if Esc-B controls roll by hydrazine and tetroxide pressure-fed in Verniers, it would be time to switch to hydrogen and oxygen pumped electrically. DLR has 500N engines on the shelf.

I propose here to add over the Esc a small hydrogen-oxygen stage, easier to develop thanks to electric pumps, which enables deep-space missions and improves the Gso performance. At D=5.4m it fits the Esc-B, Esc-A, Ariane 6, and can also propel probes, say for capture at remote planets.

 

 

It burns up to 6747kg of 720:100 hydrogen and oxygen at 25 bar, expanded to 125Pa in six 0.9m nozzles to achieve isp=4658m/s=475s. 29kN thrust permits 3300m/s in one perigee burn, to send 4.1t to Jupiter from a near-escape ellipse, reached by a first 548m/s burn from the 12t at geosynchronous transfer (Gto) provided by the Esc-B. The Esc-B alone would be marginal here, like 1.9t to Jupiter.

-----

The hydrogen and oxygen tanks have 50µm of brazed steel, 20mm and 30mm foam to evaporate 15kg and 19kg in 10mn before launch, 10 and 20 plies of multilayer insulation to evaporate 1kg and 1kg per 10h in vacuum, and polymer belts to hold at the truss; they weigh 58kg and 77kg. The truss of welded AA6005 tube weighs 200kg. Add a loose polymer fabric or net against falling objects, plus separation belts: the structure and tanks weigh 390kg, and they need no internal pressure to hold the payload.

The nozzles shall weigh 115kg, the chambers 40kg, actuators and pipes 15kg, for a 170kg passive engine; the hydrogen and oxygen screw pumps 30kg and 15kg, their motors as much, both inverters 10kg, summing to 270kg for the pumped engine. The Li-poly battery weighs 158kg.

Transmissions, sensors, steering, controls count as 200kg, unlisted items 100kg. The electric-pumped dry stage weighs 1118kg. Centrifugal pumps would enlighten, a good car fuel cell also, and turbopumps further.

-----

For lower energy missions, I consider the Esc-B is filled less to keep its start mass.

To 4500m/s above Earth's gravity, for instance for my Solar thermal rocket engine, the stage would deliver 8.0t: worthy of the heavy launcher.

To geosynchronous orbit, 8.4t instead of optimistic 7t for the Esc-B alone.

Marc Schaefer, aka Enthalpy

[Moonguy]

Your work is impressively thorough.

Correct me if I'm wrong, but the propellants you suggest are all sustainable from off-Earth resources? Including the 'Pmdeta'?

It would seem this system is mechanically easier to function with in remote locations (i.e. Mars, Mercury. . .,) and it lends itself to reusable stage designs. . .

[Enthalpy]

Elemental hydrogen exists at giant planets, from which escaping is too difficult. Oxygen if often bound with silicon, metals, carbon... where hydrogen is scarce, like our Moon or asteroids.

 

Oxygen and hydrogen are at the same place in the form of snow: ordinary comets, comets in the asteroid belt. Separating them takes much energy, but is simple enough that a robotic mission conceivably achieves it.

 

There I describe how to bring a main belt comet to martian orbit (or Earth, etc.) before separating hydrogen and oxygen:

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry757663

this is the first time I see clearly how in-situ propellant can work.

 

Ice exists at other places, but then very dilute (polar Lunar craters) or deep in the soil (Mars), which looks much more difficult with limited unmanned means.

 

-----

 

Pmdeta is an amine very common in human industry:

it's more widely available than RP-1, less flammable, marginally more efficient, fluid and dense, dirt-cheap - so I take it everywhere as a banal dense fuel.

 

While its production on Earth is very easy (with ethylene, ammonia, acetone and the like) it's less suited to small robots, for which even the simplest hydrocarbons are already a challenge.

 

Last time I considered fuel production on Mars from brought hydrogen and the atmosphere found there, my conclusion was to produce only oxygen locally, and burn hydrogen as-brought.

 

-----

 

My Solar thermal rocket engine is an enabling technology for many missions and this one ejects hydrogen or, in special cases, water without prior dissociation.

 

-----

 

Electric pumps are presently my preferred way everytime the (reasonable) battery mass is acceptable. Especially to bring back astronauts or samples from Mars, Moon and more, because I trust such an engine to start, and it brings performance unattainable to pressure feed.

[Enthalpy]

Falcon 9 v1.1 has flown, congratulations!

An additional stage would enable missions of higher energy, where O₂+Rp-1 and electric pumps bring performance. Similar engines can be roll and injection Verniers at the second stage.

A quasi-Hohman travel to Jupiter from a Geosynchronous Transfer Orbit (Gto) takes 3850m/s in a single kick, which imposes 13943kg propellant capacity and 32kN thrust for the Heavy launcher. Relaxing this case would permit a fainter, more efficient engine. Performance estimates adjust the tanks filling (this is suboptimum) to start with 13150kg at 200km 28.5° Low-Earth-Orbit (Leo) or, for the Heavy launcher, with 21280kg at 28° Gto.

 

Four 1m nozzles expanding from 60 bar to 289Pa achieve isp=392s=3846m/s. 70% efficient pumps take 58kW and 25kW at the shafts to achieve 68.2 bar. Lithium-polymer bring 475kJ/kg and weigh 20.8kg/t of propellants, up to 290kg at full tank capacity.

The truss shall break at 6MN*m. Made of AA7022 tube, it weighs some 350kg. The skin is dropped early, as in
http://www.scienceforums.net/topic/60359-extruded-rocket-structure/page-3#entry764231
15mm foam isolate the oxygen for 30min in air, and 8kg multilayer insulation for 80 days in vacuum: can be more. 200µm steel contain the oxygen, 50µm the Rp-1, polymer belts hold the tanks to the truss. The tanks weigh 76kg.

The pumped engine shall weigh 150kg, the sensors+transmissions+controls 200kg, four half separation belts 80kg, unaccounted 100kg. Empty mass is 956kg plus up to 290kg battery, or 89kg/t of propellants.

Marc Schaefer, aka Enthalpy