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GSLV MkII design, development, operations
by
Proponent
on 23 Jan, 2009 04:32
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According to astronautix.com, the original GSLV featured four UDMH/N2O4 liquid boosters strapped onto a solid-propellant first stage:
http://www.astronautix.com/lvs/gslv.htm.
Why the heck would you want liquid strap-ons, with an Isp of 281 s and a thrust-to-weight ratio of 1.6, on a solid core stage with an Isp of 266 s and a ratio of 3.2?
Another weird thing is that the core stage burns out after 93 s, whereas the LRBs burn for 159 s. So you have to drag the 28 metric tons of the burned-out core stage along for about a minute.
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#1
by
johnxx9
on 23 Jan, 2009 04:46
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They need LSBs because they can switched off and on and provides more specific impulse as you said.
The perfect example is Ariane-4. The basic Ariane 40 model with 2 solid boosters could launch around 2,500 kilograms into Geostationary transfer orbit . The 44L configuration could launch 4,790 kg to the same orbit with four liquid boosters added.
Yes, the problem with GSLV is that the solis stage burns out more quickly than the LSBs and the weight is dragged until the LSBs burn out. Increasing the solid propellant and by using solid fuels with more specific impulse could solve the problem.
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#2
by
edkyle99
on 23 Jan, 2009 04:59
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According to astronautix.com, the original GSLV featured four UDMH/N2O4 liquid boosters strapped onto a solid-propellant first stage: http://www.astronautix.com/lvs/gslv.htm.
Why the heck would you want liquid strap-ons, with an Isp of 281 s and a thrust-to-weight ratio of 1.6, on a solid core stage with an Isp of 266 s and a ratio of 3.2?
Another weird thing is that the core stage burns out after 93 s, whereas the LRBs burn for 159 s. So you have to drag the 28 metric tons of the burned-out core stage along for about a minute.
ISRO developed PSLV first, which used the S125 core solid motor augmented by up to six S9 strap on solid motors, topped by an L40 liquid second stage a solid third stage, and a liquid fourth stage. The agency's approach to create GSLV was to upgrade PSLV by replacing the solid strap-on boosters with L40 liquid strap-on boosters. The L40 second stage remained, but the upper stages were replaced by a single liquid hydrogen stage. It was an expedient, not a clean-sheet-ideal, approach. The L40 strap-on boosters provide roll control, a bit more thrust, and quite a bit more total impulse, than the PSLV boosters.
GSLV Mark III will be India's "clean sheet" approach.
- Ed Kyle
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#3
by
William Graham
on 23 Jan, 2009 06:45
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"Liquid SRBs" is a contradiction in terms
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#4
by
Proponent
on 23 Jan, 2009 18:48
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"Liquid SRBs" is a contradiction in terms
Not if SRB means
strap-on
rocket
booster!
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#5
by
William Barton
on 23 Jan, 2009 19:23
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What's TO like with solid core, liquid strap-ons?
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#6
by
isro-watch
on 24 Jan, 2009 05:35
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yes, this is the basic problem with a GSLV... the dragging up of core solid stage curtails its launch capability.... ISRO has talked of it....But says it is a design constraint....
In the years to come....ISRO may leave GSLV (and start using only GSLV MKIII version) just like it left ASLV and SLV after PSLV became operational...
There is also talk of increasing the strapons of MKIII from 2 to 4 in future to launch heavier payloads...
even then...both MKIII and GSLV... are fairly complex vehicles unlike ones like ARAINE...
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#7
by
isro-watch
on 24 Jan, 2009 05:37
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ISRO has infact indicated that all problems associated with GSLV will be solved by MKIII version....
let us hope so... a developmental flight is scheduled in 2010....
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#8
by
johnxx9
on 24 Jan, 2009 07:41
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yes, this is the basic problem with a GSLV... the dragging up of core solid stage curtails its launch capability.... ISRO has talked of it....But says it is a design constraint....
In the years to come....ISRO may leave GSLV (and start using only GSLV MKIII version) just like it left ASLV and SLV after PSLV became operational...
There is also talk of increasing the strapons of MKIII from 2 to 4 in future to launch heavier payloads...
even then...both MKIII and GSLV... are fairly complex vehicles unlike ones like ARAINE...
GSLV-Mk III is not as complex as Ariane-5 or any other vehicle.
If observed closely you will see that ISRO uses the largest amount of earth-storable fuels. That has been the main cause for their cheap launch costs and good launch history.
The fact that GSLVMk III will be more close to Titan-III than Ariane-5 will make it more less complex than Ariane-5 which has 2 Cryogenic stages.
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#9
by
publiusr
on 20 Mar, 2009 17:22
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I wonder if structural concerns were behind having the liquids atatch to a more sturdy solid core stage. With the Mk III they evidently feel this is no longer an issue.
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#10
by
hop
on 20 Mar, 2009 20:50
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GSLV-Mk III is not as complex as Ariane-5 or any other vehicle.
Do you have any documentation to support this claim ?
If observed closely you will see that ISRO uses the largest amount of earth-storable fuels. That has been the main cause for their cheap launch costs and good launch history.
Uhm, every one of those statements appears to be incorrect.
- Russia and China use more. Russia launched 10 Protons (all hypergolic with Briz M, or everything but the upper stage with Blok DM) in 2008, plus numerous smaller hypergolic vehicles. All Chinese launchers uses hypergols in the lower stages as well AFAIK.
- Labor costs have more influence than vehicle design, you can see this by looking at Russia.
- What "good launch history" ? ISROs record is not particularly good. See
http://www.geocities.com/launchreport/reliability2009.txtGSLV is 2/5 and PSLV is 12/14
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#11
by
johnxx9
on 25 Mar, 2009 09:44
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GSLV-Mk III is not as complex as Ariane-5 or any other vehicle.
Do you have any documentation to support this claim ?
If observed closely you will see that ISRO uses the largest amount of earth-storable fuels. That has been the main cause for their cheap launch costs and good launch history.
Uhm, every one of those statements appears to be incorrect.
- Russia and China use more. Russia launched 10 Protons (all hypergolic with Briz M, or everything but the upper stage with Blok DM) in 2008, plus numerous smaller hypergolic vehicles. All Chinese launchers uses hypergols in the lower stages as well AFAIK.
- Labor costs have more influence than vehicle design, you can see this by looking at Russia.
- What "good launch history" ? ISROs record is not particularly good. See http://www.geocities.com/launchreport/reliability2009.txt
GSLV is 2/5 and PSLV is 12/14
Ariane-5 uses a cryogenic upper stage and also a cryogenic core stage whereas Mk-III basically has a UDMH/N2O4 core stage. It's common sense that earth storable engines or use of such fuels is comparitively less complex than Cryogenic engines and storage of cryogenic fuels. But the overall design aspects of both the vehicles are same.
Proton uses a Semi-Cryo upper stage and so does many Chinese vehicles. Their CZ-5 is going to be a complete Semi-Cryo vehicle expect for the upper stage which is Cryogenic. I believe even Angara would use Semi-Cryo and Cryo to the max extent.
Anyway Earth storage fuels doesn't only refer to Hydrazine. It also includes solid fuel. We don't see solid boosters on Russian rockets like Proton, Soyuz etc. And also the Chinese, they don't use solid rockets on bigger launchers.
Sorry, but GSLV's launch record is 4 successes and 1 failure. PSLV's record is 12 successes, 1 failure and 1 partial failure. And also we have to take into account that Indian launchers don't fly as often as
US or Russian launchers.
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#12
by
Jim
on 25 Mar, 2009 10:09
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Anyway Earth storage fuels doesn't only refer to Hydrazine. It also includes solid fuel.
Incorrect, Earth storage fuels DOES only refer to Hydrazine like propellants. Solids are in a separate category
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#13
by
Jim
on 25 Mar, 2009 10:14
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Proton uses a Semi-Cryo upper stage
He just stated not the Brez version which is the one more in use at this time. But anyways, the amount of propellant in the first 3 stages is huge compared to the cryogenic Block DM.
You said "ISRO uses the largest amount of earth-storable fuels" The size of the Proton validates Hop's point that ISRO doesn't use the largest amount
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#14
by
hesidu
on 25 Mar, 2009 13:51
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yes, this is the basic problem with a GSLV... the dragging up of core solid stage curtails its launch capability.... ISRO has talked of it....But says it is a design constraint....
In the years to come....ISRO may leave GSLV (and start using only GSLV MKIII version) just like it left ASLV and SLV after PSLV became operational...
There is also talk of increasing the strapons of MKIII from 2 to 4 in future to launch heavier payloads...
even then...both MKIII and GSLV... are fairly complex vehicles unlike ones like ARAINE...
GSLV-Mk III is not as complex as Ariane-5 or any other vehicle.
If observed closely you will see that ISRO uses the largest amount of earth-storable fuels. That has been the main cause for their cheap launch costs and good launch history.
The fact that GSLVMk III will be more close to Titan-III than Ariane-5 will make it more less complex than Ariane-5 which has 2 Cryogenic stages.
As long as i know, UDMH/N2O4 which used on PSLV and GSLV is more expensive than RP-1/O2, if not more expensive than H2/O2.
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#15
by
hesidu
on 25 Mar, 2009 13:52
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GSLV-Mk III is not as complex as Ariane-5 or any other vehicle.
Do you have any documentation to support this claim ?
If observed closely you will see that ISRO uses the largest amount of earth-storable fuels. That has been the main cause for their cheap launch costs and good launch history.
Uhm, every one of those statements appears to be incorrect.
- Russia and China use more. Russia launched 10 Protons (all hypergolic with Briz M, or everything but the upper stage with Blok DM) in 2008, plus numerous smaller hypergolic vehicles. All Chinese launchers uses hypergols in the lower stages as well AFAIK.
- Labor costs have more influence than vehicle design, you can see this by looking at Russia.
- What "good launch history" ? ISROs record is not particularly good. See http://www.geocities.com/launchreport/reliability2009.txt
GSLV is 2/5 and PSLV is 12/14
Ariane-5 uses a cryogenic upper stage and also a cryogenic core stage whereas Mk-III basically has a UDMH/N2O4 core stage. It's common sense that earth storable engines or use of such fuels is comparitively less complex than Cryogenic engines and storage of cryogenic fuels. But the overall design aspects of both the vehicles are same.
Proton uses a Semi-Cryo upper stage and so does many Chinese vehicles. Their CZ-5 is going to be a complete Semi-Cryo vehicle expect for the upper stage which is Cryogenic. I believe even Angara would use Semi-Cryo and Cryo to the max extent.
Anyway Earth storage fuels doesn't only refer to Hydrazine. It also includes solid fuel. We don't see solid boosters on Russian rockets like Proton, Soyuz etc. And also the Chinese, they don't use solid rockets on bigger launchers.
The third stage of CZ-2E use solid rocket.
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#16
by
edkyle99
on 25 Mar, 2009 16:09
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Sorry, but GSLV's launch record is 4 successes and 1 failure.
There have been three GSLV launch vehicle failures in five attempts.
On April 18, 2001, the GSLV-D1 third stage shut down 12 seconds early, leaving GSat-1 in a 181 x 32051km x 19.2deg transfer orbit versus the planned 180 x 35975km x 19.2deg orbit. GSat-1 had insufficient fuel to maneuver to its planned geostationary orbit.
On July 10, 2006, GSLV-F2 failed to reach orbit after one of its strap-on booster Vikas engines failed two seconds into flight.
On September 2, 2007, GSLV-F4 deposited Insat-4CR into a 168 x 31,786 km x 15.8 deg orbit, well short of its planned 170 x 35,975 km x 21.7 deg orbit. Insat-4CR was able to reach its planned orbit, but at the cost of life-shortening on-board propellant.
PSLV's record is 12 successes, 1 failure and 1 partial failure. And also we have to take into account that Indian launchers don't fly as often as
US or Russian launchers.
The two PSLV failures (in 14 launches) include PSLV-D1, which failed to reach orbit, and PSLV-C1, which left IRS-1D in a 306 x 822km x 98.5deg orbit (versus a planned 817 km circular orbit) after the fourth stage suffered a helium pressurant leak. IRS-1D could not make up the 130 meter per second delta-v shortfall. Instead, it used more than 70% of its own propellant to reach a 742 x 822 km "functional" orbit.
In all of the above cases, the launch vehicles suffered some type of hardware or software failure and failed to complete their planned mission assignments, regardless of the final payload disposition. It is disingenuous, in my opinion, to call these events "successes" or "partial failures". (Partial failure? Is that like only having part of my leg cut off?) A launch vehicle failure is a launch vehicle failure, whether the rocket has flown a thousand times or once.
- Ed Kyle
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#17
by
hop
on 26 Mar, 2009 01:15
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Ariane-5 uses a cryogenic upper stage and also a cryogenic core stage whereas Mk-III basically has a UDMH/N2O4 core stage. It's common sense that earth storable engines or use of such fuels is comparitively less complex than Cryogenic engines and storage of cryogenic fuels. But the overall design aspects of both the vehicles are same.
Your claim was
GSLV-Mk III is not as complex as Ariane-5 or any other vehicle.
This is a bold claim. For example Atlas V 401 uses only 2 stages with one engine each to get to GTO.
Jim and Ed have addressed the other issues.
I'm not dissing ISRO. They are doing a good job with the resources they have, and historically new organizations have more
failures err partial successes.
@ed
I think the "partial failure" terminology makes sense if you are talking about the status of the mission. If your payload gets into orbit, but has reduced service life, then you got part of your mission. If you are talking about the LVs ability to meet it's advertised parameters (as we are here), then it's a failure.
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#18
by
sanman
on 28 Jul, 2011 13:27
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#19
by
Salo
on 29 Jul, 2011 03:49
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Mk III is a mistake in headline.
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#20
by
Danderman
on 28 Dec, 2013 19:01
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Rather than fill up the update thread with discussion, this is the place to state your opinion on whether this work or not: extra points for predicting the exact failure mode (if any) or time.
I cannot see this making it through 3rd stage operation, but I am rarely right about launches.
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#21
by
AJA
on 28 Dec, 2013 19:31
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Wow. OK.. care to substantiate your argument?
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#22
by
kanaka
on 29 Dec, 2013 06:03
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Rather than fill up the update thread with discussion, this is the place to state your opinion on whether this work or not: extra points for predicting the exact failure mode (if any) or time.
I cannot see this making it through 3rd stage operation, but I am rarely right about launches.
Could you please justify your conclusion? Basis of analysis is appreciated.
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#23
by
Danderman
on 30 Dec, 2013 00:21
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First off, this particular launcher in general has a spotty history, plus the LH2 engine has a poor flight record.
I am concerned that ISRO is continually attempting to upgrade the launcher before demonstrating that any one variant has been proven.
There are two variants of the Mark I, plus a Mark 2 plus a Mark 3 coming soon.
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#24
by
antriksh
on 30 Dec, 2013 02:16
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I am concerned that ISRO is continually attempting to upgrade the launcher before demonstrating that any one variant has been proven.
There are two variants of the Mark I, plus a Mark 2 plus a Mark 3 coming soon.
Your concern is unwarranted because ISRO is not doing anything of that sorts. GSLV (Mk1, MK2) and the upcoming GSLV Mk3 or LVM3 are two completely different launch vehicles. Most people have this misconception that LVM3 is an upgraded version of GSLV. LVM3 is a next generation launch vehicle of ISRO which comes under its expansion phase (SLV, ASLV- experimental phase, PSLV, GSLV - operational phase). All stages of LVM3 are of completely new design and share nothing in common with GSLV.
As far as GSLV is concerned, the only challenge left is mastering the cryo engine tech as there is no problem with the aerodynamic structure of the LV and the performance of both the first and second stages have been nominal.
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#25
by
Danderman
on 30 Dec, 2013 03:37
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I am suggesting that the flight history of GSLV has been so short that there may be failure modes that have not been experienced yet.
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#26
by
cave_dweller
on 30 Dec, 2013 22:33
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I am suggesting that the flight history of GSLV has been so short that there may be failure modes that have not been experienced yet.
However, is that enough information for a foregone conclusion?
Technological development is always incremental. Almost all the components of GSLV with the exception of Indigenous Cryogenic Upper Stage (CUS) have been proven.
The Indian CUS was tested once in D3 flight (failure).
http://www.isro.org/gslv-d3/gslv-d3.aspxGiven the background of ISRO, I'd say the odds of success are in favor this time.
So it remains to be seen the outcome of this flight.
Best to you ISRO. Go get it done!!
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#27
by
vyoma
on 31 Dec, 2013 03:11
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GSLV-D5 has many fixes and improvements over GSLV-D3, which was the first flight of indigenous CUS. Quoting from GSLV-D5 brochure
http://www.isro.org/gslv-d5/pdf/brochure.pdf:
These include:
• Redesign of Lower Shroud which protects the cryogenic engine during atmospheric flight of GSLV-D5
• Redesign of the wire tunnel of the cryo stage to withstand larger forces during flight
• Revised Aerodynamic characterisation of the entire launch vehicle
• Inclusion of Video Imaging System to monitor lower shroud movement during various phases of flight
• Improvements in the Cryogenic upper Stage:
→ Modified design of the Fuel Booster Turbo Pump (FBTP), taking care of the expansion and contraction of the bearings and casing at cryogenic temperatures
→ Modification of Ignition Sequence to ensure the smooth, successful and sustained ignition for Main Engine (ME), Steering Engine (SE) and Gas Generator (GG)
In addition, indigenisation of many critical systems including Liquid Hydrogen Propellant Acquistion System (to prevent the possibility of contamination), Polyimide pipelines and Liquid Oxygen & Liquid Hydrogen Level Sensors has been successfully accomplished.
In order to validate the design improvements, the following extensive qualification test have been carried out on the engine at the Main Engine Test (MET) facility and the High Altitude Test (HAT) facility:
• Two acceptance tests for flight unit of FBTP
• High altitude tests to confirm the ignition sequence in flight under vacuum
• Cryogenic Main Engine (200 sec) and Steering Engine (100 sec) acceptance tests
I guess, all eyes will be on turbo pumps and CUS lower shroud.
All the best, ISRO. Let's do it
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#28
by
johnxx9
on 05 Jan, 2014 10:41
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Rather than fill up the update thread with discussion, this is the place to state your opinion on whether this work or not: extra points for predicting the exact failure mode (if any) or time.
I cannot see this making it through 3rd stage operation, but I am rarely right about launches.
You have an answer now!
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#29
by
isro-watch
on 05 Jan, 2014 10:58
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Yes, the answer !
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#30
by
Mazo
on 05 Jan, 2014 11:20
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After a technology denial regime instituted by the US Government that prevented Russia from transferring the cryogenic staged combustion cycle engine technology to India in 1992 citing MTCR and other meaningless treaties (while NASA was buying Soviet RD-180 technology because it was "better"), today the ISRO has built, tested and demonstrated its own cryogenic engine with a text book launch using an ALL INDIAN vehicle.
25 years later- mission accomplished! Success or Failure - you decide.
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#31
by
xm11
on 05 Jan, 2014 11:25
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GSLV-D5 successfully launches GSAT-14 from SDSC SHAR, Sriharikota
on Jan 05, 2014
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#32
by
xm11
on 05 Jan, 2014 11:28
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#33
by
Danderman
on 05 Jan, 2014 14:31
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I cannot see this making it through 3rd stage operation, but I am rarely right about launches.
I am pleasantly surprised to be wrong on this one.
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#34
by
sanman
on 05 Jan, 2014 16:33
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I was reading that ISRO is working on uprating the cryogenic engine from its current 73kN to 90kN thrust:
http://isp.justthe80.com/launchers/isro-s-cryogenic-upper-stage-cus#TOC-More-Powerful-Cryogenic-EngineThat still doesn't correct the design flaw of carrying the dead weight of the 1st-stage core post-burnout. But I guess ISRO doesn't mind because of the cheapness of developing GSLV-Mk1/2.
Is there any opportunity to correct that flaw somehow? Seems like they'd have to swap out the solid core for a liquid one, and that might even allow them to probably get rid of the strap-ons too.
Gee, I didn't realize that the cryogenic upper stage for GSLV-Mk3 was a different technology (gas generator) as compared to GSLV-Mk2 (staged combustion). I'd always thought both rockets were using the same cryogenic engine technology, and that therefore the Mk-3 was just an evolution on the configuration of the Mk-2.
If the staged combustion in GSLV-Mk2 is not being replicated in the GSLV-Mk3, then what further platforms will make use of the staged combustion technology? It seems to me that if ISRO has taken the trouble to develop staged combustion technology, that it should want to use it on a launch vehicle that isn't handicapped.
To me, that means that either GSLV-Mk2 should be replaced with a better-designed rocket, or else the design flaw in Mk2 should somehow be corrected.
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#35
by
antriksh
on 06 Jan, 2014 02:18
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I was reading that ISRO is working on uprating the cryogenic engine from its current 73kN to 90kN thrust:
http://isp.justthe80.com/launchers/isro-s-cryogenic-upper-stage-cus#TOC-More-Powerful-Cryogenic-Engine
That still doesn't correct the design flaw of carrying the dead weight of the 1st-stage core post-burnout. But I guess ISRO doesn't mind because of the cheapness of developing GSLV-Mk1/2.
Is there any opportunity to correct that flaw somehow? Seems like they'd have to swap out the solid core for a liquid one, and that might even allow them to probably get rid of the strap-ons too.
Gee, I didn't realize that the cryogenic upper stage for GSLV-Mk3 was a different technology (gas generator) as compared to GSLV-Mk2 (staged combustion). I'd always thought both rockets were using the same cryogenic engine technology, and that therefore the Mk-3 was just an evolution on the configuration of the Mk-2.
If the staged combustion in GSLV-Mk2 is not being replicated in the GSLV-Mk3, then what further platforms will make use of the staged combustion technology? It seems to me that if ISRO has taken the trouble to develop staged combustion technology, that it should want to use it on a launch vehicle that isn't handicapped.
To me, that means that either GSLV-Mk2 should be replaced with a better-designed rocket, or else the design flaw in Mk2 should somehow be corrected.
The CUS already reached 90 kn in this flight. it started with 75 and reached 90, and then again back to 75. Staged combustion cycle technology development experience will help in the semi-cryo engine under development and probably for future high rated cryo engines for RLV. Once LVM3 is available, ISRO will concentrate on building only >3 ton com sats and so role of GSLV mk2 will be limited to lofting GSO imaging stas and inter-planetary missions.
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#36
by
cave_dweller
on 06 Jan, 2014 02:46
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That still doesn't correct the design flaw of carrying the dead weight of the 1st-stage core post-burnout. But I guess ISRO doesn't mind because of the cheapness of developing GSLV-Mk1/2.
Is there any opportunity to correct that flaw somehow? Seems like they'd have to swap out the solid core for a liquid one, and that might even allow them to probably get rid of the strap-ons too.
They could decrease the fuel capacity in the boosters to avoid the 1st stage post burn out dead weight problem. Though I am not sure if the weight decrease in L440 chambers will compensate for loss of thrust duration.
What is the weight of the 1st stage core shell casing?
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#37
by
sanman
on 06 Jan, 2014 03:13
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They could decrease the fuel capacity in the boosters to avoid the 1st stage post burn out dead weight problem. Though I am not sure if the weight decrease in L440 chambers will compensate for loss of thrust duration.
What is the weight of the 1st stage core shell casing?
Or alternatively, they could increase the size of the 1st-stage core, to make it last longer, which would improve capacity rather than lowering it! It would also allow that near-worthless 2nd-stage to be eliminated. Call that the GSLV Mk-2B.
Come to think of it - why didn't they just do that from the start?
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#38
by
sanman
on 06 Jan, 2014 03:26
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The CUS already reached 90 kn in this flight. it started with 75 and reached 90, and then again back to 75. Staged combustion cycle technology development experience will help in the semi-cryo engine under development and probably for future high rated cryo engines for RLV. Once LVM3 is available, ISRO will concentrate on building only >3 ton com sats and so role of GSLV mk2 will be limited to lofting GSO imaging stas and inter-planetary missions.
Ah, that's right - saw those comments about the variable thrust being cranked up to hit 90.
What about restart capability? I read that this engine has that, but I guess it wasn't tested on this flight.
I'd read that the next big focus for LPSC will be on Semi-cryo engines, and so this development program will result in larger boosters for lower stages. So this seems to be the opposite of everyone else's engine development path. Everybody else did semi-cryo first, because that's easiest and gets you flying sooner - certainly more practical from a business perspective. India seems to have approached development from the opposite side, which seems to be what's taking it so long in getting into the commercial launch market.
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#39
by
cave_dweller
on 06 Jan, 2014 04:25
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Or alternatively, they could increase the size of the 1st-stage core, to make it last longer, which would improve capacity rather than lowering it! It would also allow that near-worthless 2nd-stage to be eliminated. Call that the GSLV Mk-2B.
Come to think of it - why didn't they just do that from the start?
I thought about this. I figured it would be much easier to manipulate the fuel content in the strap on boosters rather than modifying the engineering of the solid core.
Solid fuels tend to be heavier and have shorter specific impulse.
ISRO could maintain the same stage design layout.
Decrease fuel capacity in Stage-1 strap on boosters.
Increase fuel capacity in either Stage 2 or Stage 3 and maintain current weight configuration but achieve better performance.
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#40
by
cave_dweller
on 06 Jan, 2014 04:38
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IMO, I think ISRO would be better off developing a Semi-cryogenic 1st stage + Cryogenic 2nd Stage configuration instead of trying to improve solid core and hypergolic engine performance.
Which is what I believe will be configuration of ULV.
A single Semi-cryogenic core 1st stage.
Need more thrust? Add semi-cryogenic boosters. Or increase the fuel capacity of 1st stage and use engine clusters.
A single cryogenic 2nd stage. Need more thrust? Increase fuel capacity and/or use engine clusters.
I think ISRO should move away from hypergols for rocket engines (and just limit using hypergols to satellites, spacecrafts and such). Expensive, unstable, not easily transportable, needs a lot of careful handling.
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#41
by
sanman
on 06 Jan, 2014 04:44
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#42
by
cave_dweller
on 06 Jan, 2014 05:36
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Yep, the evolution path is ULV->RLV->TSTO->Air Breathable TSTO
The problem India has is not so much the lack of capability. But that of bootstrapping overhead and industry support
India is ways behind in materials technology which was one of the biggest factors in delays related to ICE (Indian Cryogenic Engine).
Its a chicken and egg problem. For industry to be sustainable there has to be enough volume (or number of flights from ISRO). For ISRO to have a high volume of flights the materials should be affordable.
Today it is solved by government funding. But that has its limits which is manifested as lengthened delays.
With ICE, India might have achieved point of critical mass where it can attract enough flight customers to make it a thriving industry.
It would be good for India to be another competitor in the space. This will have the dual benefit of economic benefits AND making spaceflight cheaper by competition and reflexive volume.
I feel really happy for India!!
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#43
by
antriksh
on 06 Jan, 2014 10:57
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may be because CE7.5 is staged combustion and others are either gas generator or expander cycles
Asking again: would you provide your source?
Low scientific literacy, among society as well as policy makers elected by such a society - has ALWAYS been, and still is THE security threat - to socio-economic prosperity, and consequently to life and property.
A simple search can provide the answer
read about CUS:
http://www.isro.org/gslv-d3/pdf/GSLV-D3_GSAT-4%20Brochure.pdf
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#44
by
quanthasaquality
on 06 Jan, 2014 11:10
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I think India's rocket program was quite smart. Start out with the 4 stage PSLV. A 4 stage rocket can have high design margins. Second stage engine was tech transfered from Ariane 4. Third and fourth stages replaced by hydrogen upper stage purchased from Russia, until indigenous copy is made. First stage supports strap on boosters. Larger solid first stages with tighter margins are gradually made (s-200). Then make a rocket twice as large, with double engines. Keep the rocket team plodding along for a few decades. India then has a rocket for the neglected ~5 ton to LEO market, and a 10 ton to LEO rocket for more mainstream satellites, requiring only a small industrial base. Quite clever.
Then India works on a new 20 ton hydrogen engine. Why not just use 2 GSLV mk 2 hydrogen engines?
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#45
by
AJA
on 06 Jan, 2014 11:47
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K Radhakrishnan does some market prognosticationAnswering a question on the commercial potential of the GSLV with India’s own cryogenic engine, Dr. Radhakrishnan said that while the trend now was to build 5.5-tonne satellites, there was a niche market for two-tonne communication satellites. Satellites weighing 3.2 tonnes to 4.5 tonnes were being built in the world. “So there is a set of satellites” which could be put into orbit by the GSLV.
And... some (recycled?) information on international collaborations..
India has been invited to take part in experiments onboard the International Space Station, Dr. Radhakrishnan said.
The ISRO and NASA would jointly build a satellite with synthetic aperture radar for earth observations. A.S. Kiran Kumar, Director, Space Applications Centre of the ISRO, Ahmedabad, was working with the Jet Propulsion Laboratory for preparing the project report related to the spacecraft, he said.
A simple search can provide the answer
read about CUS: http://www.isro.org/gslv-d3/pdf/GSLV-D3_GSAT-4%20Brochure.pdf
This discussion is taking place in
two threads.
Anyway,
you said it was the D-
5 brochure, and now you do a bait-and-switch quoting the D-
3 brochure. I found the ISRO press release anyway. Btw... if a search is simple, simpler still is for the OP to include the link in the original post.
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#46
by
edkyle99
on 06 Jan, 2014 13:50
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They keep referring to the cryogenic stage... I assume this is a HydroLox stage?
ISRO calls its previously existing hypergolic liquid stages "storable". It calls its LOX/LH2 stage "cryogenic". It calls its LOX/kerosene R&D effort "semi-cryogenic".
- Ed Kyle
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#47
by
sanman
on 06 Jan, 2014 14:55
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Regarding the chicken-and-egg hurdle, I wonder if ISRO could have progressed faster by engineering a cryogenic engine based on LCH4 instead of LH2? At least LCH4 has an industrial base to support it, unlike LH2, and so moving in that direction isn't a leap out into empty space like LH2 is.
Even with the planned RLV/TSTO, they should still consider LCH4, because of its high Isp, and plus it's not as hard to work with as LH2.
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#48
by
akula2
on 06 Jan, 2014 20:13
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By the way, ISRO GSLV D5 happens to be first orbital launch of 2014
My somewhat belated congratulations to all
I'm intrigued to learn:
a) what's the total cost of his mission combined with the delay (cost of leak)?
b) will there be any Return on Investment when total cost considered?
More importantly, the fuel tank material has been changed, fully phasing out the traditional but corrosion-prone aluminium-zinc combine, called AFNOR 7020. The new alternative, aluminium-copper alloy called AA2219, is now the material for all PSLV and GSLV tanks.
c) I look at it as a Managerial lapse. Just wondering, will there be any disciplinary action of any sort taken? If yes, to what level/extent?
I understand Space launches are complex but that gaffe could have been avoided - saving loads of time/resources/money. Please someone answer, thank you.
Lastly, it was a good writeup by William Graham
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#49
by
cave_dweller
on 06 Jan, 2014 20:17
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Regarding the chicken-and-egg hurdle, I wonder if ISRO could have progressed faster by engineering a cryogenic engine based on LCH4 instead of LH2? At least LCH4 has an industrial base to support it, unlike LH2, and so moving in that direction isn't a leap out into empty space like LH2 is.
Even with the planned RLV/TSTO, they should still consider LCH4, because of its high Isp, and plus it's not as hard to work with as LH2.
Oxidizer: LOX Oxidizer Density: 1.140 g/cc.
Oxidizer Freezing Point: -219 deg C.
Oxidizer Boiling Point: -183 deg C.
Fuel LCH4:Fuel Density: 0.424 g/cc.
Fuel Freezing Point: -184 deg C.
Fuel Boiling Point: -162 deg C.
Fuel LH2:Fuel Density: 0.071 g/cc.
Fuel Freezing Point: -259 deg C.
Fuel Boiling Point: -253 deg C.
It appears Liquid Methane offers better density. But its only 80% Hydrogen (4 out of 5 atoms). So there is the overhead of carrying 20% dead weight (Carbon). Though there is the benefit of not having to cool it as aggressively as H2. I am also not sure about the efficiency of the combustion cycle.
By going the LOX+LH2 route, it would help maintain a conservative approach (proven technology, low risk) as opposed to going directly for LOX+LCH4 and going through the overhead of learning a new technology with an unknown risk.
Regardless, I think India would do well to develop its materials and human resources base. For instance only US, Russia and China have Titanium smelters. Saudi Arabia is building one.
Also, I think India should also consider open/modular satellite systems design. For instance, a single large satellite in GSO orbit can serve the needs of many regions instead of each region launching its own small satellite in the same orbit. An open & modular approach will bring efficiency of design, re-use, lower costs and will also drive the need for heavier launch vehicles. ISRO has demonstrated these intentions with the inclusion of experiments from multiple nations on Chandrayaan-1. They could extend the same to commercial services as well.
With the Indian economy developing combined with judicious approach to developing resources and industries could certainly enable exciting developments in Indian Space Technology in the future.
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#50
by
chota
on 07 Jan, 2014 01:33
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Disciplinary action should be the last thing to do for a developing program IMHO. Reason for failure looks silly when retrospected.
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#51
by
cave_dweller
on 07 Jan, 2014 04:24
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By the way, ISRO GSLV D5 happens to be first orbital launch of 2014
My somewhat belated congratulations to all
I'm intrigued to learn:
a) what's the total cost of his mission combined with the delay (cost of leak)?
b) will there be any Return on Investment when total cost considered?
More importantly, the fuel tank material has been changed, fully phasing out the traditional but corrosion-prone aluminium-zinc combine, called AFNOR 7020. The new alternative, aluminium-copper alloy called AA2219, is now the material for all PSLV and GSLV tanks.
c) I look at it as a Managerial lapse. Just wondering, will there be any disciplinary action of any sort taken? If yes, to what level/extent?
I understand Space launches are complex but that gaffe could have been avoided - saving loads of time/resources/money. Please someone answer, thank you.
Lastly, it was a good writeup by William Graham
R&D tends to be the process of making mistakes, learning from them, experiencing accidents, training to avoid them and other growing up pains. Its the process of maturation.
India doesn't have as strong a manufacturing economy comparable to most western economies or even China. So when a mistake or an accident does happen, the time to correct it is longer. That is just the fact of life and nature of development in emerging nations.
Cryogenic technology is a key enabler and a great leverage a nation can posses. It demonstrates technological capability and maturity. So, immediate ROI is one thing. However there are many other intangible and ripple effects through out the economy and human resource development that are not readily apparent but very much a causal outcome.
However, from a purely economic/balance sheet standpoint:
Cost of current mission: Rs 365 cr (US $56.15 million @ Rs 65/$1)
Cost of GSLV D5: Rs 220 cr (US $33.84 million)
Cost of GSAT 14: Rs 145 cr (US $22.30 million)
India currently pays Rs 500 cr (ESA/Arianespace - US $76.93 million ) for 3.5-Tonne satellite. Cost of Satellite not included.
India has to pay foreign service providers either in US $ or EU currency. This has the effect of worsening the current account deficit since India has unfavorable balance of payments owing to the fact that India imports more than it exports.
This then limits large payment items such as satellite launches to essential services.
This also limits the kind of satellites India can launch since there is a serious risk of technology exposure when satellites are launched by external parties.
You can imagine how the above two can have a debilitating effect on overall technological development due to high cost of satellite transport + unfavorable foreign currency position.
With maturation of GSLV and related technologies, India would gain the capability to launch heavier and more capable satellites from its home ground for a much lower cost and develop its economy, human resources and technology.
Even with all cumulative failures of GSLV, time delays etc., this is an essential technology and India has done extremely well than almost all of the predecessors.
Keep in mind Cryogenic technology has been in general human consciousness for over 80 yrs now. It took US & Russia 30 yrs of active thinking to develop theirs.
India started in 1990s with a tattered economy and bootstrapped the entire process from scratch at a pittance.
India has done well. In my opinion, exceedingly well given the resources and the challenges.
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#52
by
akula2
on 07 Jan, 2014 09:09
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Also, I think India should also consider open/modular satellite systems design. For instance, a single large satellite in GSO orbit can serve the needs of many regions instead of each region launching its own small satellite in the same orbit. An open & modular approach will bring efficiency of design, re-use, lower costs and will also drive the need for heavier launch vehicles. ISRO has demonstrated these intentions with the inclusion of experiments from multiple nations on Chandrayaan-1. They could extend the same to commercial services as well.
With the Indian economy developing combined with judicious approach to developing resources and industries could certainly enable exciting developments in Indian Space Technology in the future.
One of my focus areas or target audience.
The bigger push for reforms are needed for proactive Private Sector participation. Whether ISRO supports or not, I'm actively working on a solid plan for 2020; currently closing on the location minus Bangalore or Chennai. If everything goes well one Indian state will have dedicated three Satellites (not small at all). Perhaps a few valuable services could be extended to 1 or 2 neighboring states. I foresee an incredible demand from a few states which are high on the Indian GDP rankings.
IMO, Govt. and ISRO could learn a lot from India's vibrant Automobile sector success story.
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#53
by
akula2
on 07 Jan, 2014 11:01
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Disciplinary action should be the last thing to do for a developing program IMHO. Reason for failure looks silly when retrospected.
I think you got rattled
I meant, there must be some kind of Internal Memo etc to the concerned staff. Also, remarks should go into their service records. If you look at various International organizations in different verticals (including NASA or anyone), you'd be really surprised. I'm NOT saying there won't be ANY failure or it's unacceptable.
Fact is, any Govt. Organization is accountable because of the Taxpayer money. Anyone could be called for a hearing to seek a testimony or for any serious lapse/accident/losses. It happens in US and few other Space nations. NASA isn't excluded.
Look at how FDA and India's Drug regulator work. It takes a lot of planning, time, energy and money to operate a company in US compared to India. It's just that many Indians are not used to hear such things, but once people move abroad it teaches a lot.
R&D tends to be the process of making mistakes, learning from them, experiencing accidents, training to avoid them and other growing up pains. Its the process of maturation.
India doesn't have as strong a manufacturing economy comparable to most western economies or even China. So when a mistake or an accident does happen, the time to correct it is longer. That is just the fact of life and nature of development in emerging nations.
I concur with your views.
I run a few manufacturing non-public companies in three nations, about 1600 work in Pharma/Bio and Medical Electronics sectors. My 4th company would be in Aerospace domain (Communications etc). Actually, I'm well aware of R&D and its associated costs, risks etc. Many Pharma companies adhere to stringent FDA/EU norms; if any violate they've exclusive powers to penalize and/or ban a plant. I guess you might have read the latest Ranbaxy's Mohali plant news:
http://www.fda.gov/newsevents/newsroom/pressannouncements/ucm368445.htmWhat I mean to emphasize is, no one is above accountability. I personally know a few NASA Project Managers who take pride for successfully delivering projects within shoe string budgets and tight time frames. Imagine, if ISRO can save only 100 crore each year (1 billion INR) they can spend that on offering student fellowships, promote STEM at colleges, investing in latest infrastructure like building new Xeon/Tesla based Hybrid Super Computers and so on. I don't believe if anyone tells me there is ZERO loss of money (2013-14 budget is 6792 crore). Now, don't misconstrue me as if I'm against ISRO spending
On some post I wished that ISRO should get at least $3 billion each year (or 18,600 crore) to pursue various missions.
Having said, all it takes is Effective Governance and right Policies across the Govt. and Corporate spectrum. What's happening in Delhi heralds a new dawn for India. As an IIT Alumnus, I'm proud to be associated with AAP as a NR donor/supporter. The change will come...
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#54
by
antriksh
on 07 Jan, 2014 13:07
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#55
by
vyoma
on 07 Jan, 2014 14:47
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I've some questions on CE 7.5 and GSLV Mk2 in general:
1) Does it have restart capability?
2) Variable thrust or fixed thrust?
3) GSLV D5 had 3.4m diameter fairing. Will ISRO test 4m diameter fairing in subsequent launches? If so, would it require any design/structural changes in lower stages, or is it just plug-n-play?
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#56
by
sanman
on 07 Jan, 2014 15:59
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1) Yes, but I don't think it was flight-tested during D5
2) Yes, and that was indeed flight-tested during D5
3) The new acoustic testing facility is meant to evaluate the vibrational and structural loads associated with changing stuff like the fairing size. It seems doubtful they'd redesign the rest of the rocket to accommodate a new fairing that doesn't work well with the existing rocket.
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#57
by
antriksh
on 07 Jan, 2014 23:59
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I've some questions on CE 7.5 and GSLV Mk2 in general:
1) Does it have restart capability?
2) Variable thrust or fixed thrust?
3) GSLV D5 had 3.4m diameter fairing. Will ISRO test 4m diameter fairing in subsequent launches? If so, would it require any design/structural changes in lower stages, or is it just plug-n-play?
1) Nope
2) Variable. 7.5 to 8.2 (would like to correct that 9 ton not reached yet)
3) D6 will test 4m fairing. No structural changes required, or plug-n-play. Exhaustive analysis done using simulation and wind tunnels.
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#58
by
kanaka
on 08 Jan, 2014 00:03
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I've some questions on CE 7.5 and GSLV Mk2 in general:
1) Does it have restart capability?
2) Variable thrust or fixed thrust?
3) GSLV D5 had 3.4m diameter fairing. Will ISRO test 4m diameter fairing in subsequent launches? If so, would it require any design/structural changes in lower stages, or is it just plug-n-play?
1) Nope
2) Variable. 7.5 to 8.2 (would like to correct that 9 ton not reached yet)
3) D6 will test 4m fairing. No structural changes required, or plug-n-play. Exhaustive analysis done using simulation and wind tunnels.
What does it take to achieve restart capability in the space?
what modifications are needed to attain 9 ton ?
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#59
by
antriksh
on 08 Jan, 2014 12:50
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I've some questions on CE 7.5 and GSLV Mk2 in general:
1) Does it have restart capability?
2) Variable thrust or fixed thrust?
3) GSLV D5 had 3.4m diameter fairing. Will ISRO test 4m diameter fairing in subsequent launches? If so, would it require any design/structural changes in lower stages, or is it just plug-n-play?
1) Nope
2) Variable. 7.5 to 8.2 (would like to correct that 9 ton not reached yet)
3) D6 will test 4m fairing. No structural changes required, or plug-n-play. Exhaustive analysis done using simulation and wind tunnels.
What does it take to achieve restart capability in the space?
what modifications are needed to attain 9 ton ?
advanced igniter and new hardware.
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#60
by
sanman
on 08 Jan, 2014 16:55
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Hmm, so it doesn't have the restart capability yet? Is that planned for the future?
Likewise, if thrust is improved upto 90kN, are there any plans to take it higher - like say, to 100kN?
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#61
by
cave_dweller
on 09 Jan, 2014 08:56
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What does it take to achieve restart capability in the space?
what modifications are needed to attain 9 ton ?
I believe this is a fairly complex and expensive problem. An additional ignition setup would mean additional complexity and weight. Also, there is a problem of "boiling off" due to resulting increased volume during burn offs.
That is, if you run the motor for 60 seconds. And shut it off, you'll have to compensate for the volume evacuated by that 60 seconds of fuel. Which would mean the cryogenic setup would have to further cool the propellants to maintain fuel density for appropriate flow mass rate or nominal thrust when re-ignited.
This issue also exists when the engine doesn't have to be restarted. But I believe is handled by the turbo pump. And results in gradual loss of thrust and is acceptable when executing a full burn.
Hypergolic fuels are better suited for in-space applications (orbital maneuvers etc). Since they do not need to be cooled to maintain fuel density and also because they do not need an additional ignition setup. Hypergolic fuel and oxidizer auto-ignite on contact.
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#62
by
cave_dweller
on 09 Jan, 2014 09:06
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Hmm, so it doesn't have the restart capability yet? Is that planned for the future?
Likewise, if thrust is improved upto 90kN, are there any plans to take it higher - like say, to 100kN?
I think it should be easy to increase the thrust simply by using engine clusters.
However the problem may be one of maintaining the aero-dynamic and weight character of the flight.
Which should again be fairly straight forward.
I think the biggest mystery of cryogenic engines are fuel tank material science (weight vs strength) and turbo and turbo booster pump design and materials.
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#63
by
antriksh
on 09 Jan, 2014 12:30
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Hmm, so it doesn't have the restart capability yet? Is that planned for the future?
Likewise, if thrust is improved upto 90kN, are there any plans to take it higher - like say, to 100kN?
Restart capability will be on CE20's upgraded versions. That wont be required if LVM3 is ready by 2017-2018.
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#64
by
sanman
on 14 Jan, 2014 04:32
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Restart capability will be on CE20's upgraded versions. That wont be required if LVM3 is ready by 2017-2018.
I don't understand - CE20 is the cryo engine stage for LVM3, isn't it?
So you're saying the later versions of it will have restart capability, as part of the LVM3 launch stack.
Why won't CE20 restart capability be required if LVM3 is ready by 2017-18?
Also, I was looking at this:
http://www.b14643.de/Spacerockets_1/India/GSLV/Gallery/GSLV-2_D5.htm
It says "In development: Payload Assist Modul (PAM-G) for direct injection to orbit"
So in what circumstances is it used and why?
EDIT: I found mention of it from you in another thread:
http://forum.nasaspaceflight.com/index.php?topic=30037.msg1003555#msg1003555Still - what does direction injection into orbit mean? I thought everything gets directly injected into orbit. You mean like without any orbit-raising maneuvers required?
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#65
by
rnataraja
on 14 Jan, 2014 04:33
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Regarding the chicken-and-egg hurdle, I wonder if ISRO could have progressed faster by engineering a cryogenic engine based on LCH4 instead of LH2? At least LCH4 has an industrial base to support it, unlike LH2, and so moving in that direction isn't a leap out into empty space like LH2 is.
Even with the planned RLV/TSTO, they should still consider LCH4, because of its high Isp, and plus it's not as hard to work with as LH2.
At a broader level than described here. How well has private industry aspired to support the ISRO's efforts in India? What is missing in them today?
If someone who was inspired(like me) by the grand but secret univers and space, and are very serious about satisfying the desire by contributing to advancements and be part of an aspiring private industry. What is a good area?
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#66
by
ss1_3
on 14 Jan, 2014 14:51
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Still - what does direction injection into orbit mean? I thought everything gets directly injected into orbit. You mean like without any orbit-raising maneuvers required?
Most certainly, they are referring to direct injection into GEO orbit as against the current capability of launching payload into GTO first and then performing on board engine burns to get into GEO.
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#67
by
antriksh
on 14 Jan, 2014 15:30
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Restart capability will be on CE20's upgraded versions. That wont be required if LVM3 is ready by 2017-2018.
I don't understand - CE20 is the cryo engine stage for LVM3, isn't it?
So you're saying the later versions of it will have restart capability, as part of the LVM3 launch stack.
Why won't CE20 restart capability be required if LVM3 is ready by 2017-18?
Sorry I didnt put it clearly. CE7.5 wont have restart capability, but CE20 will have that in its upgraded version (not the one that will be demonstrated in 2016-17)
Restart capability will be on CE20's upgraded versions. That wont be required if LVM3 is ready by 2017-2018.
It says "In development: Payload Assist Modul (PAM-G) for direct injection to orbit"
So in what circumstances is it used and why?
Still - what does direction injection into orbit mean? I thought everything gets directly injected into orbit. You mean like without any orbit-raising maneuvers required?
Yep no orbit raising maneuver by the satellite.
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#68
by
johnxx9
on 14 Jan, 2014 16:51
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Some interesting tidbits! It's from 2009. One change is that even the basic MkII configuration has flown with 40 tons of propellant in the second stage.
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#69
by
sanman
on 14 Jan, 2014 20:42
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What does this Payload Assist Module (PAM) use? What fuel is it based on? What engine does it use - something like Liquid Apogee Motor? So it's like an uppermost stage, but it's not cryogenic - or is it? Why use a non-cryo engine on top of a cryo engine stage? That would make no sense, unless it's meant for deferred use and needs to avoid boil-off.
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#70
by
johnxx9
on 15 Jan, 2014 07:05
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What does this Payload Assist Module (PAM) use? What fuel is it based on? What engine does it use - something like Liquid Apogee Motor? So it's like an uppermost stage, but it's not cryogenic - or is it? Why use a non-cryo engine on top of a cryo engine stage? That would make no sense, unless it's meant for deferred use and needs to avoid boil-off.
It'll be used for MEO missions, to place navigation satellites directly into 19,000km circular orbit. It is derived from the upper stage of PSLV. It's been in development for some time. Uses hypergolic propellants.
Heres it being tested.
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#71
by
sanman
on 15 Jan, 2014 17:01
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So since MEO is less than the GTO that the Cryogenic Upper State can deliver to, then why use this PAM along with the Cryogenic Upper Stage? It sounds like you'd want to use this PAM thing in lieu of the Cryogenic Upper Stage, if you're only trying to reach MEO. Or maybe this PAM thing can boast your payload capacity to MEO and GEO?
Can this PAM thing help with BEO missions to the Moon or Mars, etc?
I wonder what kind of lift capacity it can give to these places.
Although it seems rather strange to me that you'd use a liquid hypergolic stage on top of a lighter cryo stage. Doesn't that sort of defeat the purpose of having a cryo stage? What does the liquid hypergolic stage give you that the cryo stage doesn't?
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#72
by
Lars_J
on 15 Jan, 2014 19:13
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Although it seems rather strange to me that you'd use a liquid hypergolic stage on top of a lighter cryo stage. Doesn't that sort of defeat the purpose of having a cryo stage? What does the liquid hypergolic stage give you that the cryo stage doesn't?
Hypergolic systems are easier to make long endurance stages from. They don't need to be cooled, they are easier to ignite. This is useful for final orbit insertion several hours or days into the mission. Most hydrogen cryogenic stages have a lifetime of a few hours at most.
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#73
by
sanman
on 15 Jan, 2014 20:19
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Ahh, so this hypergolic stage is meant for deferred use.
So if you were designing a lunar lander, would that likewise be done best as hypergolic too?
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#74
by
baldusi
on 16 Jan, 2014 00:10
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Although it seems rather strange to me that you'd use a liquid hypergolic stage on top of a lighter cryo stage. Doesn't that sort of defeat the purpose of having a cryo stage? What does the liquid hypergolic stage give you that the cryo stage doesn't?
The current CE-7.5 engine can't restart. And for circularization (be it GSO or MEO) you'd need at least the circularization burn. And i doubt that the Cryo stage has more than a couple of hours of life. Hypergolics are storable and can make multi hour missions.
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#75
by
Lars_J
on 16 Jan, 2014 15:21
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Ahh, so this hypergolic stage is meant for deferred use.
So if you were designing a lunar lander, would that likewise be done best as hypergolic too?
Yes. I think all lunar landers (manned and unmanned) have been hypergolic.
There have been proposals for cryogenic landers (Altair), but even then the ascent stage has been hypergolic.
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#76
by
cave_dweller
on 16 Jan, 2014 19:52
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Ahh, so this hypergolic stage is meant for deferred use.
So if you were designing a lunar lander, would that likewise be done best as hypergolic too?
Yes. I think all lunar landers (manned and unmanned) have been hypergolic.
There have been proposals for cryogenic landers (Altair), but even then the ascent stage has been hypergolic.
Are hypergolics easier to store in space than LOX+LH2?
How do hypergolics react to the extreme temperatures of space?
And how do cryogenics react?
How are they sheltered (if at all)? Always keep it oriented away from Sun?
How do they manage to isolate "hot" components from "cold" components of the space craft?
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#77
by
Lars_J
on 16 Jan, 2014 19:57
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Ahh, so this hypergolic stage is meant for deferred use.
So if you were designing a lunar lander, would that likewise be done best as hypergolic too?
Yes. I think all lunar landers (manned and unmanned) have been hypergolic.
There have been proposals for cryogenic landers (Altair), but even then the ascent stage has been hypergolic.
Are hypergolics easier to store in space than LOX+LH2?
How do hypergolics react to the extreme temperatures of space?
And how do cryogenics react?
I'm no expert, but cryogenics obviously have to be kept cold. Hypergolics are less sensitive, but you still have to prevent them from freezing or becoming too hot. It is just an easier temperature range to manage through use of insulation and heaters. So yes, easier to store in space.
http://en.wikipedia.org/wiki/Hypergolic_propellant
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#78
by
sanman
on 16 Jan, 2014 22:14
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In some thread someone commented that cryogenic restart capability is impinged upon by the fact that when you burn off some of your cryo propellants, some of the remaining propellants will be more likely to go into vapor phase. Has anybody ever tried a bladder approach, whereby the propellant container shrinks in volume to keep the remaining propellant compressed? Obviously the elastic force of your bladder has to exceed the vapor pressure of the volatile.
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#79
by
vyoma
on 17 Jan, 2014 01:13
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In some thread someone commented that cryogenic restart capability is impinged upon by the fact that when you burn off some of your cryo propellants, some of the remaining propellants will be more likely to go into vapor phase. Has anybody ever tried a bladder approach, whereby the propellant container shrinks in volume to keep the remaining propellant compressed? Obviously the elastic force of your bladder has to exceed the vapor pressure of the volatile.
Centaur upper stage uses stainless steel balloon tank, which doesn't have any structural integrity on its own. LH2/LOX are pumped to keep balloons inflated.
But, I guess restart capability is achieved a bit differently. There are two problems to solve during cryo restart:
1) Vapors in tanks and fuel lines, which expand rapidly and can damage tanks.
2) In zero-g, liquid LH2/LOX will be "floating" in tanks. So when engine is restarted, fuel may not flow into fuel lines and turbopumps.
To address these issues before restart:
1) Centaur vents vapors into Space, before starting main cryo engine. This results in some loss of fuel.
... we are going to take this Atlas valve, scale it down a little bit, mount the same valve on the liquid hydrogen tank as the liquid oxygen tank, and we’re going to command the valve to open when you have to vent.
The solution was to design a vent fin or snout on the nose fairing that extended about 50 inches from the tank, just far enough away to keep the hydrogen gas from igniting along the hot surface of the vehicle.
2) Centaur uses small conventional rocket motors (H2O2 powered ullage rockets) to accelerate a bit in order to settle down "floating" liquid fuel at the bottom of tanks. This makes sure that fuel indeed flows into fuel lines when main cryo engine is restarted.
The new knowledge of the ullage area also led to greater confidence in managing the propel-lants during the coast phase. A Reaction Control System (RCS) was designed to provide acceleration just before venting. The system consisted of a hydrogen peroxide supply bottle, lines, valves, and small vernier motors called thrusters. Before the tank pressure reached a certain point, small thrusters in the tank were fired to reposition the propellants in the aft end of the vehicle.
Then a valve could be opened to relieve pressure without danger of loss of liquid propellant. The same thrusters were fired in preparation for starting the engine, so that liquid rather than gaseous hydrogen and oxygen entered the pumps.
Sources:
http://history.nasa.gov/SP-4230.pdfhttp://www.spacelaunchreport.com/aclv3cb.htmlI guess balloon tanks might also help in keeping fuel settled down, by deflating when fuel is consumed (speculating).
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#80
by
Steven Pietrobon
on 17 Jan, 2014 03:12
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Yes. I think all lunar landers (manned and unmanned) have been hypergolic.
For the actual propulsion just before landing, yes that is true. Surveyor did use a solid stage (a Thiokol TE-364 motor) in a staged descent with the actual landing using storable propellants (monomethyl hydrazine hydrate or MMH-H2O and MONO-10 or 90% N2O4 and 10% NO). The planned Soviet LK lander was to use a Block D kerolox stage in staged descent, with the LK lander itself using UDMH and N2O4 for the actual landing.
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#81
by
sanman
on 17 Jan, 2014 03:13
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In some thread someone commented that cryogenic restart capability is impinged upon by the fact that when you burn off some of your cryo propellants, some of the remaining propellants will be more likely to go into vapor phase. Has anybody ever tried a bladder approach, whereby the propellant container shrinks in volume to keep the remaining propellant compressed? Obviously the elastic force of your bladder has to exceed the vapor pressure of the volatile.
Centaur upper stage uses stainless steel balloon tank, which doesn't have any structural integrity on its own. LH2/LOX are pumped to keep balloons inflated.
...
I guess balloon tanks might also help in keeping fuel settled down, by deflating when fuel is consumed (speculating).
Yeah, an elastic bladder would keep contracting as propellant is used up.
What about glassy amorphous metal alloys? For example, there are certain metal alloys which have highly elastic properties. If you could make balloon tanks out of such elastic alloys and then pump them up full of LOX and LH2 so that they were under tension, they could then contract as the propellants were consumed.
http://phys.org/news/2011-07-japanese-material-scientists-superelastic-alloy.html
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#82
by
baldusi
on 17 Jan, 2014 03:21
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H2 is liquid at about 20K, LOX is around 93K. At those temperatures, very little materials are elastic. Not to mention the bad manners of the H2 molecules to get everywhere they can, generating embritelment and free H2 everywhere.
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#83
by
hop
on 17 Jan, 2014 04:09
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I guess balloon tanks might also help in keeping fuel settled down, by deflating when fuel is consumed (speculating).
No, the centaur "balloon" tanks are pressure stabilized, meaning they use pressure to maintain their shape, they aren't like elastic balloons. Non-cryogenic systems often do have an internal bladder, but as baldusi says, this is harder with cryogenics.
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#84
by
AJA
on 17 Jan, 2014 09:02
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In some thread someone commented that cryogenic restart capability is impinged upon by the fact that when you burn off some of your cryo propellants, some of the remaining propellants will be more likely to go into vapor phase. Has anybody ever tried a bladder approach, whereby the propellant container shrinks in volume to keep the remaining propellant compressed? Obviously the elastic force of your bladder has to exceed the vapor pressure of the volatile.
This issue isn't restricted to cryogenics. ANY fuel tank'll face this issue. Turn a bottle of water upside down.. the efflux isn't smooth -- precisely because water falling out results in a vacuum at the top of the bottle... and periodically, the atmospheric pressure breaks surface tension in the lower free surface enough to allow some air to form a bubble and gush in.
So what do they do? They take the lightest gas known - compress it vastly, and tack it on to the rocket. They add a valve whereby this gas is slowly vented into the fuel tanks as the tank drains. (As long as the rocket's thrusting and undergoing acceleration, there is a well defined free surface within the tank) This keeps the pressure on top of the liquid, and "should" keep flow smooth (given other favourable circumstances). When you work out the arithmetic, you also make sure to set the super-incumbent gas pressure in the empty (or at-cutoff) fuel tank, catered to entirely by your pressurant gas - is greater than the vapour pressure of your propellant at the ambient temperatures likely to be encountered. The greater you make it, the lower the boil-off (but lower payload fraction).
Anyway, you do all that, and then you realise all sorts of things are catching fire, and exploding. So you swap out the lightest known gas, with the second lightest known gas
, while singing -- about your first iteration at a solution --
TL;DR - They use Helium to pressurise the tanks as they're evacuated.
Balloon tanks would be awesome, but aside from material concerns and challenges, you'd also have dynamic stability issues... how d'you ensure isotropic contraction? No one here has had a party balloon they inflated and let go of - fly straight and true
Then again, those party balloons don't have active guidance.
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#85
by
baldusi
on 17 Jan, 2014 10:46
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Two points:
1) On your H2 tank, you can use the lightest gas. But normally this is done by heating a tiny bit just enough to get gaseous H at enough pressure (I think it's around 0.33MPa/50psi or so).
Membranes are used, but mostly on spherical or nearly so tanks. Think of a how would you put a membrane on a very thin and long tank, like the first stage tanks. You'd waste a lot of mass and the membrane would have to move from the top to the bottom, probably rubbing against itself. And elastomers are not known for their low coefficient of friction.
On a related note when a membrane is used, the other side of the tank is pressurized with He from a bottle with a regulator. This gives a constant pressure for pressure fed engines and work in any direction. The valves from the bottle to the tank was what got stuck during SpaceX's CRS-2, and that's why they couldn't activate the RCS. The tank's lacked the He that actually gave the pressure to the pressure fed engines.
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#86
by
sanman
on 17 Jan, 2014 13:37
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Balloon tanks would be awesome, but aside from material concerns and challenges, you'd also have dynamic stability issues... how d'you ensure isotropic contraction? No one here has had a party balloon they inflated and let go of - fly straight and true Then again, those party balloons don't have active guidance.
Well, the tank doesn't itself have to be the superstructure or fuselage of the rocket. The tank can be a flexible internal enclosure inside the superstructure. Maybe it could be a flexible accordion type of structure, to ensure it contracts in one axial direction. Boron Nitride, by the way, has shown itself capable of forming a quasi-hexagonal mesh similar to that in Graphene and Carbon Nanotubes, but the BN structure is capable of buckling in an accordion-like manner, in a linear direction. That's one reason why BN nanotubes are being researched for tank armor, because this buckling characteristic offers energy-absorbing or damping properties. It seems like these intrinsic properties would still be available even at cryogenic temperatures.
http://www.academia.edu/2718708/Compressive_Buckling_of_Boron_Nitride_Nanotubes_with_Hydrogen_Storage
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#87
by
AJA
on 17 Jan, 2014 19:17
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Well, the tank doesn't itself have to be the superstructure or fuselage of the rocket. The tank can be a flexible internal enclosure inside the superstructure. Maybe it could be a flexible accordion type of structure, to ensure it contracts in one axial direction.
What about the evacuated space between the inside of your aero-shell and the outside of your flexible tank wall? Aero-shells need to be pressurised to maintain their shape and rigidity. What's going to keep your aero-shell from collapsing because of the pressure differential? Also, why would your flexible tank continue to passively shrink when there's an evacuated zone outside it? Any active system that performs mechanical work to squeeze the balloon is going to cost mass. What's that? You'd use Helium? Uhh....
Unless you want to think about entraining ambient air, that's what you'd probably converge on. But capturing ambient air would require you to siphon some off, and would cost you drag, shock, vibration, heating, and is still entrained mass that you're going to be carrying uphill. Not to mention the aerodynamic problems caused by the flow, when it encounters a possibly anisotropic shrinking tank.
In any case, I wasn't talking about the problems with the aerodynamic interactions as much as I was referring to the mass distribution upsets caused by the deviations in the contraction of the balloon; throwing the rocket off balance, and moving its centre of gravity around unpredictably. That would be a GNC nightmare.
Boron Nitride, by the way, has shown itself capable of forming a quasi-hexagonal mesh similar to that in Graphene and Carbon Nanotubes, but the BN structure is capable of buckling in an accordion-like manner, in a linear direction. That's one reason why BN nanotubes are being researched for tank armor, because this buckling characteristic offers energy-absorbing or damping properties. It seems like these intrinsic properties would still be available even at cryogenic temperatures.
That paper measures the change in the buckling behaviour of Boron nanotubes... when the nanotubes are subjected to an external compressional load. It says nothing about the direction of such buckling (accordion etc.), or the reversibility. This way.. you can buckle ANY material to shrink the volume, provided you ensure that it doesn't open up cracks that allow the LH2 out.
Plus, they're measuring the buckling in ISOLATED BN NANOtubes. Not BN Nanotubes packed so closely enough to form a seal impenetrable to Hydrogen molecules. But let's assume that a ring of BN nanotubes, spaced far enough apart to still exhibit the properties of nanotubes, rather than form a bulk solid structure -- i.e. stacked like the pillars of Stonehenge are still able to trap Hydrogen in the inner area because of a phenomenal adsorption power. Let's also assume that you can stack successive Stonehenges on atop the other to form a macroscopic cylinder -- or you're able to get a nanotube that's several metres in length; forgetting the difficulties of manufacturing a MULTI-STOREY structure out of NANO-tubes. (Boron Nitride btw -- atleast in the Wurtzite mineral form -- is harder than diamond and possibly the hardest substance known to man. Well...short of the exotic matter in stellar cores I guess).
For this structure to work as balloon tank, you'd have to ensure that the shrinkage happened only from top down... and that all your nano-tubes are aligned, and that it happens in the same direction, as opposed to the bottom half the tank shrinking down, and the top half shrinking up -- leading to a tear. You'd have to ensure this on a microscopic scale.
But...if you did all that though, I'll tell you one more thing you could do. You could get rid of all the wiring that needs to pass through that stage. The Boron Nitride nanotubes could act as fibre-optics!
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#88
by
sanman
on 17 Jan, 2014 23:43
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Gee, I didn't know that propellant tank pressurization inflates the aeroshell. Why would you even need that in space though, where there's no external ambient pressure? I thought that's the only relevant context for this discussion - trying to restart your cryo engine in space.
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#89
by
vyoma
on 22 Jan, 2014 15:04
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Here are two insightful articles on GSLV Mk2, GSLV Mk3, and Indian and Russian cryogenic engines.
An interview with S. Ramakrishnan, Director, VSSC:
http://www.frontline.in/cover-story/will-be-able-to-repeat-the-success/article5590227.eceAn interview with K. Radhakrishnan, Chairman, ISRO:
http://www.frontline.in/cover-story/gslv-mkiii-the-next-milestone/article5596588.eceWhat elements of the Mark II cryogenic engine and stage, which fired GSLV-D5, still retain the legacy of Russian engine technology and design? How much of it is truly indigenous and how much of it relies on the Russian heritage?
Basically, both engines use the “staged combustion cycle”. That is one approach compared with gas generator cycle, which we are using for the C20 engine to be used in GSLV MkIII. There are several other differences, conceptually also, especially the igniter system that we are using, which is totally different from what has been used in the Russian engine. [In MkII liquid oxygen (LOX) and gaseous hydrogen (GH2) are ignited by pyrogen-type igniters in the pre-burner as well as in the main and steering engines during initial stages, as against pyrotechnic ignition in the Russian engine.] But in the staged combustion cycle, similarities can be found in the way the engine is started and the steering engines are used for controllability.
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#90
by
vyoma
on 01 Oct, 2015 06:15
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http://timesofindia.indiatimes.com/city/hyderabad/IICT-develops-polymer-for-Isros-cryogenic-engine/articleshow/49164649.cmsHYDERABAD: In a significant development in rocket science, the Indian Institute of Chemical Technology (IICT) has developed a polymer for use in cryogenic rocket by the Indian Space Research Organisation (Isro).
Use of polymer instead of aluminium or other alloys in fuel plumbing and tubes in the cryogenic engine drastically reduces the overall weight of the rocket. The advantage of it is that the rocket can carry a higher payload. The weight of fuel plumbing and tubes made of polymer is just 10 per cent of those made of aluminium or other alloys.
"As of now, Isro is importing the polymer from the US. But we now have developed the Fluorinated Ethylene Propylene (FEP) polymer," said Shekharam Tammishetti, senior principal scientist and head, polymers and functional materials division, CSIR-IICT. He told TOI that the technology will be transferred to Isro.
In the cryogenic engine, the fuel that is used is oxygen and hydrogen. When the two elements are burnt together, they produce a lot of energy. At that temperature, only FEP polymer can withstand the heat.
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#91
by
russianhalo117
on 01 Oct, 2015 06:18
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http://timesofindia.indiatimes.com/city/hyderabad/IICT-develops-polymer-for-Isros-cryogenic-engine/articleshow/49164649.cms
HYDERABAD: In a significant development in rocket science, the Indian Institute of Chemical Technology (IICT) has developed a polymer for use in cryogenic rocket by the Indian Space Research Organisation (Isro).
Use of polymer instead of aluminium or other alloys in fuel plumbing and tubes in the cryogenic engine drastically reduces the overall weight of the rocket. The advantage of it is that the rocket can carry a higher payload. The weight of fuel plumbing and tubes made of polymer is just 10 per cent of those made of aluminium or other alloys.
"As of now, Isro is importing the polymer from the US. But we now have developed the Fluorinated Ethylene Propylene (FEP) polymer," said Shekharam Tammishetti, senior principal scientist and head, polymers and functional materials division, CSIR-IICT. He told TOI that the technology will be transferred to Isro.
In the cryogenic engine, the fuel that is used is oxygen and hydrogen. When the two elements are burnt together, they produce a lot of energy. At that temperature, only FEP polymer can withstand the heat.
FEPvis made in Houston and is transported under train car lease designation DOWX
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#92
by
vyoma
on 30 Oct, 2015 12:45
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#93
by
vyoma
on 30 Oct, 2015 12:46
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Can mods please change this topic title to: GSLV MkII & cryo stage discussion thread
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#94
by
input~2
on 30 Oct, 2015 14:34
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Can mods please change this topic title to: GSLV MkII & cryo stage discussion thread
Done!
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#95
by
vyoma
on 04 Sep, 2016 22:12
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Models of Vikas, CE-7.5 and CE-20 engines.
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#96
by
vyoma
on 05 Sep, 2016 21:45
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VSSC, ISRO recently issued a tender (Aug 2016) for supplying CFRP elements for GSLV MkII cryogenic stage:
http://www.isro.gov.in/sites/default/files/tenders/pt197.pdf This RFQ provides the details of realization of various CFRP elements required for cryo upper stage of GSLV launch vehicle. These are made of carbon epoxy prepreg and Aluminium honeycomb core materials and is fabricated by hand layup and Autoclave curing.
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#97
by
sanman
on 07 Sep, 2016 18:26
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Since SpaceX has achieved such great improvements through propellant densification, I wonder if ISRO has looked at this as a way to improve its launch vehicle performance? Anybody know?
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#98
by
vineethgk
on 08 Sep, 2016 03:45
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GSLV has a 20-year past and a long way aheadFrom now on the target is for two GSLV launches a year, which means a launch every six months. It took us a year between the last one and this. We want to improve that
There is an increasing demand for the GSLV. We are looking at possible opportunities for it to provide commercial launches, just as the PSLV has done.
Yes, even for full launches. A few discussions are going on. We have to wait for the talks to firm up.
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#99
by
worldtimedate
on 09 Sep, 2016 05:18
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Asked whether the space organisation is now comfortable with the indigenous cryogenic stage, a complex system compared to solid and earth-storable liquid propellant rocket stages, S. Somanath, Director, Liquid Propulsion System Centre, ISRO, said the scientists were very confident about it.
The cryogenic stage has settled into a system today. After the failure of the first stage, we identified the problems, conducted very detailed analysis and studies. Lots of tests were done simulating actual conditions, and they were very successful. We have mastered the technology, Mr. Somanath said.
He said ISRO was developing another engine, C-25, that will be twice as powerful as the current one.
ISRO now expects the GSLV to pick up business like the PSLV.
Source :
ISRO eyes Venus mission--- [ --- ]
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#100
by
vyoma
on 09 Sep, 2016 05:43
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#101
by
vineethgk
on 09 Sep, 2016 05:57
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http://www.deccanchronicle.com/business/economy/090916/isro-eyeing-330-billion-market.html
Isro officials said they were also planning to enhance the GSLV’s [GTO] payload capacity — from 2,600 kg to 2,800 kilos.
That's news for me. Would that imply that ISRO does intend to keep GSLV-II running for a considerable period even after LVM3 is ready for operational flights?
EDIT: It brings another question to my mind. Are there many payloads in the 2-2.5 tonne range in the international GTO launch market that GSLV-II could tap into?
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#102
by
sanman
on 09 Sep, 2016 14:04
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http://www.deccanchronicle.com/business/economy/090916/isro-eyeing-330-billion-market.html
Isro officials said they were also planning to enhance the GSLV’s [GTO] payload capacity — from 2,600 kg to 2,800 kilos.
That's news for me. Would that imply that ISRO does intend to keep GSLV-II running for a considerable period even after LVM3 is ready for operational flights?
EDIT: It brings another question to my mind. Are there many payloads in the 2-2.5 tonne range in the international GTO launch market that GSLV-II could tap into?
If you check ISRO's launch roster, you'll see there are plenty of flights slated for Mk-II, even after LMV3 comes into service. Maybe it's an easier/cheaper vehicle for them to fabricate, since Mk-II is similar to the PSLV architecture.
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#103
by
Stan Black
on 11 Sep, 2016 13:52
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So, I went through the available documentation and tried to figure out what payload was assigned to each GSLV. I used the outcome budget and the five-year plans.
Some documents here:-
http://forum.nasaspaceflight.com/index.php?topic=32023.msg1569474#msg1569474
Period 2001-2002
GSLV-D1 GSAT-1
GSLV-D2 GSAT-2
GSLV-D3 GSAT-3
GSLV-C1 INSAT-3D
GSLV-C2 GSAT-4
GSLV-C3
Period 2005-2006
GSLV-D1 GSAT-1
GSLV-D2 GSAT-2
GSLV-D3 GSAT-4
GSLV-F01 GSAT-3
GSLV-F02 INSAT-4C
GSLV-F03 INSAT-3D
GSLV-F04 INSAT-4D (GSAT-5)
GSLV-F05 INSAT-4E (GSAT-6)
GSLV-F06
11th Five Year Plan 2007-2012
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001 Mk. I
GSLV-D2 GSAT-2 launched 08.05.2003 Mk. I
GSLV-D3 GSAT-4 Mk. II
GSLV F01-F10 ten launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03 INSAT-3D
GSLV-F04 INSAT-4CR
GSLV-F05 GSAT-5
GSLV-F06 GSAT-9
GSLV-F07 GSAT-6
GSLV-F08 Chandrayaan-2
GSLV-F09 GSAT-12
GSLV-F10 GSAT-13
Notes:-
GSLV-F04 launched September 2007.
Period 2008-2009
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
GSLV-D3 GSAT-4 CUS-03 (A6)
GSLV F01-F10 ten launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03 INSAT-3D
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05 GSAT-5
GSLV-F06 GSAT-9 (identical to INSAT-4C)
Period 2009-2010
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
CUS-02
GSLV-D3 GSAT-4 CUS-03 (A6)
CUS-04 (A7)
CUS-05
GSLV F01-F16 sixteen launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03 GLONASS
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05 GSAT-5
GSLV-F06 INSAT-3D
GSLV-F07 GSAT-6
Notes:-
Reference to GSLV-F16.
Period 2010-2011
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
GSLV-D3 GSAT-4 CUS-03 (A6)
CUS-04 (A7)
CUS-05
GSLV F01-F16 sixteen launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05 GSAT-6
GSLV-F06 GSAT-5P
GSLV-F07 INSAT-3D
GSLV-F08 GSAT-7
Notes:-
No reference to GSLV-F03.
GSLV-D3 with GSAT-4 launched April 2010.
GSLV-F06 with GSAT-5P launched December 2010.
Period 2011-2012
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
GSLV-D3 GSAT-4 launched 25.04.2010 CUS-03 (A6)
GSLV-D4 GSAT-4R CUS-05 (A8)
GSLV-D5 GSAT-6
GSLV F01-F16 sixteen launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05 GSAT-7
GSLV-F06 GSAT-5P launched 25.12.2010
Notes:-
Only year with reference to a GSLV-D4.
12th Five Year Plan 2012-2017
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
GSLV-D3 GSAT-4 launched 25.04.2010 CUS-03 (A6)
GSLV-D4
GSLV-D5 GSAT-14
GSLV-D6 GSAT-6
GSLV F01-F16 sixteen launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05
GSLV-F06 GSAT-5P launched 25.12.2010
GSLV-F07
GSLV-F08 Chandrayaan-2
GSLV-F09 GSAT-9
GSLV-F10 GISAT
GSLV-F11 GSAT-6A
GSLV-F12 GSAT-7A
GSLV-F13 INSAT-3DR
Notes:-
No reference to GSLV-D4, GSLV-F05 or GSLV-F07 (GSLV-F03 already dropped).
Period 2012-2013
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
GSLV-D3 GSAT-4 launched 25.04.2010 CUS-03 (A6)
GSLV-D4
GSLV-D5 GSAT-14 CUS-05 (A8)
GSLV-D6 GSAT-6 CUS-06 (A9)
GSLV F01-F16 sixteen launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05
GSLV-F06 GSAT-5P launched 25.12.2010
GSLV-F07
GSLV-F08 Chandrayaan-2 CUS-07 (A10)
Notes:-
Launch of GSAT-7 using procured launch services.
Period 2013-2014
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
GSLV-D3 GSAT-4 launched 25.04.2010 CUS-03 (A6)
GSLV-D4
GSLV-D5 GSAT-14 CUS-05 (A8)
GSLV-D6 GSAT-6 CUS-06 (A9)
GSLV F01-F16 sixteen launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05
GSLV-F06 GSAT-5P launched 25.12.2010
GSLV-F07
GSLV-F08 Chandrayaan-2 CUS-07 (A10)
Notes:-
Launch of INSAT-3D using procured launch services.
GSLV-D5 with GSAT-14 launched January 2014.
Period 2014-2015
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
GSLV-D3 GSAT-4 launched 25.04.2010 CUS-03 (A6)
GSLV-D4
GSLV-D5 GSAT-14 launched 05.01.2014 CUS-05 (A8)
GSLV-D6 GSAT-6 CUS-06 (A9)
GSLV F01-F16 sixteen launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05
GSLV-F06 GSAT-5P launched 25.12.2010
GSLV-F07
GSLV-F08
Notes:-
Reference to GSLV-F07.
Period 2015-2016
Notes:-
GSLV-D6 with GSAT-6 launched August 2015.
Period 2016-2017
Rocket Payload Upper Stage
GSLV-D1 GSAT-1 launched 20.04.2001
GSLV-D2 GSAT-2 launched 08.05.2003
GSLV-D3 GSAT-4 launched 25.04.2010 CUS-03 (A6)
GSLV-D4
GSLV-D5 GSAT-14 launched 05.01.2014 CUS-05 (A8)
GSLV-D6 GSAT-6 launched 27.08.2015 CUS-06 (A9)
GSLV F01-F16 sixteen launch vehicles
Rocket Payload Upper Stage
GSLV-F01 GSAT-3 launched 20.09.2004
GSLV-F02 INSAT-4C launched 10.07.2006
GSLV-F03
GSLV-F04 INSAT-4CR launched 02.09.2007
GSLV-F05 INSAT-3DR launch 07.2016 CUS-07 (A10)
GSLV-F06 GSAT-5P launched 25.12.2010
GSLV-F07
GSLV-F08
GSLV-F09 GSAT-9 launch 03.2017 CUS-08
Notes:-
GSLV-F05 with INSAT-3DR launched September 2016.
See also:-
https://web.archive.org/web/*/http://www.isro.org
http://forum.nasaspaceflight.com/index.php?topic=1173.msg727300#msg727300http://forum.nasaspaceflight.com/index.php?topic=1173.msg567502#msg567502
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#104
by
vineethgk
on 11 Sep, 2016 14:16
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Thanks!
So the future launches that are planned for GSLV appear to be :
SouthAsiaSat
GSAT-9
Chandrayaan-2
GISAT-1
GSAT-7A
GSAT-6A
NISAR
Mars-2(?)
Anything I missed?
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#105
by
vineethgk
on 26 Sep, 2016 09:00
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Now that ISRO could demonstrate multiple engine restarts in its hypergolic upper stage, what would be the key challenges it would face in replicating the same in its cryogenic upper stages? I guess having an ignition system that works for multiple burns is one challenge, as it does not use hypergols. Also, would the engineering difficulties be different in CE-7.5 (SCC) compared to CE-20 (GGC)? Any insights are appreciated.
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#106
by
cave_dweller
on 27 Sep, 2016 23:05
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Now that ISRO could demonstrate multiple engine restarts in its hypergolic upper stage, what would be the key challenges it would face in replicating the same in its cryogenic upper stages? I guess having an ignition system that works for multiple burns is one challenge, as it does not use hypergols. Also, would the engineering difficulties be different in CE-7.5 (SCC) compared to CE-20 (GGC)? Any insights are appreciated.
Would have to use hypergols or battery to ignite LH2/LOX. Which raises the question, what is the current ignition mechanism in CE 7.5? And what will it be in CE 20 engine?
Alternately, (I don't believe this has been attempted before) could also let the stage coast in the orbit and use solar power to spark the fuel. But this would require either co-incidental orientation of the stage with sunlight OR specifically re-orienting the stage to receive sunlight -- which means it would require some extraneous thrust.
Another trade-off to consider is where to put the fuel. In the launch vehicle or on the satellite?
In the case of Mangalyaan, being that PSLV didn't have the thrust to develop the velocity to deliver the craft directly into an earth escape orbit (similarly in the case of Chandrayaan-1), the choice was made to store the fuel in the satellite/craft and use a combination of thruster burns + Earth gravity to develop velocity and ultimately escape.
This brings forth another question, could a larger satellite be used as a launch vehicle for a cluster of nano satellites that don't have self sufficient orientation/navigation thruster/booster engines?
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#107
by
worldtimedate
on 28 Sep, 2016 20:34
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As I expected, since GSLV MK-II's Cryogenic Upper Stage does NOT have multiple restartable capability, it will launch the Chandrayaan-2 into a parking orbit similar to GTO.
Here is the report from Frontline Magazine Science Section
G. Nagesh, Project Director, Chandrayaan-2, said the orbiter, the lander and the rover were together called the composite module. The GSLV-Mk II will first place this composite module in an orbit of 170 km by 19,500 km, called earth-parking orbit. From there, with the help of the liquid engines in the orbiter, we will take Chandrayaan-2 to the moon’s orbit of 100 km, he said. It is exactly the same as Chandrayaan-1's orbit. Once Chandrayaan-2 (that is, the composite module) is in the lunar orbit, ISRO will beam commands to it for the lander to fly out of the orbiter.
Source :
Cryogenic gains for GSLV--- [ --- ]
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#108
by
sanman
on 28 Sep, 2016 22:44
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As I expected, since GSLV MK-II's Cryogenic Upper Stage does NOT have multiple restartable capability, it will launch the Chandrayaan-2 into a parking orbit similar to GTO.
Here is the report from Frontline Magazine Science Section
G. Nagesh, Project Director, Chandrayaan-2, said the orbiter, the lander and the rover were together called the composite module. The GSLV-Mk II will first place this composite module in an orbit of 170 km by 19,500 km, called earth-parking orbit. From there, with the help of the liquid engines in the orbiter, we will take Chandrayaan-2 to the moon’s orbit of 100 km, he said. It is exactly the same as Chandrayaan-1's orbit. Once Chandrayaan-2 (that is, the composite module) is in the lunar orbit, ISRO will beam commands to it for the lander to fly out of the orbiter.
Source : Cryogenic gains for GSLV
--- [ --- ]
I've always heard everyone say that the CUS was designed to have restart capability, but that it's just that ISRO hasn't tried to flight test it yet - maybe they don't want to risk trying it out on a politically sensitive prestige mission like Chandrayaan-2?
http://space.stackexchange.com/questions/3505/does-isros-cryogenic-upper-stage-have-restart-capability
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#109
by
K210
on 29 Sep, 2016 05:35
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According to previous annual reports the CE-7.5 is designed to be restartable
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#110
by
worldtimedate
on 29 Sep, 2016 05:43
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I agree with Sanman that GSLV MK-II CUS might have multiple restart capability like all the CUS of other space fairing nations such as USA, Japan, ESA and China. But ISRO has yet to test and master it. Going by the recent PSLV C35 launch of 8 Satellits into different orbits, I am sure ISRO is capable of emulating that feat in GSLV MK-II CUS. Just as PSLV tested the multiple restart feature of upper stage twice before launching satellites, GSLV MK-II has to do the same thing. As for the Chandrayaan II mission, it would be unwise of ISRO to take such risk in doing the multiple restart feature to inject it to Lunar Orbit.
--- [ --- ]
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#111
by
baldusi
on 29 Sep, 2016 18:46
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According to previous annual reports the CE-7.5 is designed to be restartable
I didn't knew that the KVD-1 was ever restartable. In fact, I don't think that any non hypergolic Russian staged combustion engine is restartable. Start-up is particularly delicate issue for staged combustion engines, the hypergolic gas generator Vikas is trivial to restart in comparison.
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#112
by
sanman
on 30 Sep, 2016 17:32
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So I want to apologize and correct:
CUS is not restartable at present.
(That answer came right from the horse's mouth)
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#113
by
worldtimedate
on 31 Jan, 2017 18:06
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Recently in the Indian Science Congress in a speech Mr. Somnath of LPSC Director has mentioned that GSLV MK-II payload cababily would be
incrementally increased to 3 Ton or even more.
Umamaheshwaran, the GSLV MK-II project director already said last year that GSLV MK-II payload capability
would be increased to 2.8 Ton.
Source :
Cryogenic gains for GSLVIf that happens, that would be a monumental achievement for the naughty boy GSLV MK-II that gave the ISRO Launch Vehicle Team lot of headache until it was successfully launched for 3 consecutive times with the Indigenous Cryogenic Engine CE-7.5. Please check the Following Video.
Indian science congress 2017 Live Stream 4 : ( Speech by Mr. S. Somnath, LPSC Director )
--- [ --- ]
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#114
by
vineethgk
on 19 May, 2017 19:46
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Late bloomer GSLV is small cheerThe late bloomer may even be a short-lived intermediate rocket instead of being ISRO’s primary satellite vehicle as it was planned, as a few ISRO old-timers and industry watchers privately suggest.
The GSLV is caught in a glaring mismatch: it cannot lift India’s bigger satellites; and the size that it can lift is out of fashion and does not make economic sense.
While ISRO was perfecting the GSLV and falling behind schedule with the rocket’s crucial cryogenic stage, it progressed on the spacecraft side and upgraded the communication satellites to 3,000-plus kg in 2005. This was done to pack more punch (or transponders) per spacecraft. It would be roughly 24 regular transponders for 2,000 kg; 36 transponders for 3,000 kg and 48 transponders in a four-tonner.
Replying to a query from The Hindu, Gagan Agrawal, analyst with the U.S.-based space industry consulting firm Northern Sky Research, said: “The communications satellite market is consistently looking at payload sizes greater than four tonnes and the question remains whether the GSLV or [the bigger] MKIII can cater to the market [yet.] ”
As painted in the report, I would have been equally skeptical of GSLV Mk2's long-term career prospects until a few months back. But the recent reports of ISRO looking at building 2-tonne class all-electric satellites, and issuing tenders for the supply of EPS towards that goal, might save the GSLV after all. That possibility doesn't appear to have been taken into account in this news item. Moreover, rather than the prospect of ISRO throwing in the towel on Mk2, recent news reports have quoted ISRO sources hinting that they plan to increase the GTO payload capability of the launcher to 3 tonnes by incremental upgrades as well.
Incidentally, after reading reports of the all-electric, 2300kg SES-15 that was launched by Soyuz from CSG yesterday, I'm quite hopeful that we might see the 'naughty boy of ISRO' flying for some time. An EPS-enabled satellite of similar class can pack a decent number of transponders and can change the fortunes of GSLV Mk2 in the coming years, at least in the domestic market. There might be many requirements for which launching a smaller EPS comsat on a cheaper GSLV Mk2 rocket makes better sense than doing a heavier one on LVM3/GSLV Mk3. And then there is the advantage of a higher flight rate as well owing to it's PSLV heritage.
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#115
by
seshagirib
on 10 Jun, 2017 16:42
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So I want to apologize and correct:
CUS is not restartable at present.
(That answer came right from the horse's mouth)
On the contrary:
See this youtube video
"AIS 2017 A2 Dr V Narayanan LPSC 14 Feb 17"
Dr. Narayanan has confirmed the built in restart capability for both CUS-7 and C20. (question and answer session from 25:55 onwards ). He also said that the capability will be tested after the MK-3 flight.
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#116
by
sanman
on 11 Jun, 2017 00:14
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So I want to apologize and correct:
CUS is not restartable at present.
(That answer came right from the horse's mouth)
On the contrary:
See this youtube video
"AIS 2017 A2 Dr V Narayanan LPSC 14 Feb 17"
...
Dr. Narayanan has confirmed the built in restart capability for both CUS-7 and C20. (question and answer session from 25:55 onwards ). He also said that the capability will be tested after the MK-3 flight.
I hate to push back, but the source I asked was just as authoritative as Dr Narayanan. So, maybe this then sounds like some threshold must have been crossed between Sep-30-2016 and Feb-14-2017 in order for both statements to be true? Maybe it achieved certification for restartability during those 5 months?
Prior to asking, I too had always heard people on the net saying that the CE-7.5 was restartable, but I then asked anyway, just to be sure.
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#117
by
demonslayer
on 01 Jul, 2017 17:48
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So I want to apologize and correct:
CUS is not restartable at present.
(That answer came right from the horse's mouth)
On the contrary:
See this youtube video
"AIS 2017 A2 Dr V Narayanan LPSC 14 Feb 17"
...
Dr. Narayanan has confirmed the built in restart capability for both CUS-7 and C20. (question and answer session from 25:55 onwards ). He also said that the capability will be tested after the MK-3 flight.
I hate to push back, but the source I asked was just as authoritative as Dr Narayanan. So, maybe this then sounds like some threshold must have been crossed between Sep-30-2016 and Feb-14-2017 in order for both statements to be true? Maybe it achieved certification for restartability during those 5 months?
Prior to asking, I too had always heard people on the net saying that the CE-7.5 was restartable, but I then asked anyway, just to be sure.
It's written in the book From fishing Hamlet to red planet that the kvd-1 was meant to be restartable. Since they didn't want to take the two-burn risk, they decided to run it uprated first and then later at normal thrust to sort of simulate a two burn.
Sent from my Nexus 6 using Tapatalk
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#118
by
sanman
on 02 Jul, 2017 09:46
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It's written in the book From fishing Hamlet to red planet that the kvd-1 was meant to be restartable. Since they didn't want to take the two-burn risk, they decided to run it uprated first and then later at normal thrust to sort of simulate a two burn.
But a real re-start means re-lighting the stage , which hasn't yet been tested in flight, correct? I wonder how much they've tested it on the ground.
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#119
by
demonslayer
on 04 Jul, 2017 06:17
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It's written in the book From fishing Hamlet to red planet that the kvd-1 was meant to be restartable. Since they didn't want to take the two-burn risk, they decided to run it uprated first and then later at normal thrust to sort of simulate a two burn.
But a real re-start means re-lighting the stage , which hasn't yet been tested in flight, correct? I wonder how much they've tested it on the ground.
No the restart hasn't been tested on the ground. Restart capability not only requires re-ignition ability but also the gas bottles need to be at an adequate pressure level in order to ensure the required propellant flow into the preburner. Not to mention the propellant loss due to boil off.
Sent from my Nexus 6 using Tapatalk
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#120
by
K210
on 28 Mar, 2018 07:03
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The next generation Vikas engine developed by the Liquid Propulsion Systems Centre (LPSC) is being flown for the first time. LPSC director V Narayanan told Express that the improved engine would give a significant advantage in terms of enhancing payload capability. “Usually, the chamber pressure is 58 bar, but with the use of high-thrust Vikas engine, we will achieve 62 bar, which is a 6% increase in thrust that gives us 70 kgs of additional payload gain in this mission. Right now, we are going to use the high-thrust Vikas engine only in the second stage. Basically, we are validating it. For Chandrayaan-2 mission, we will be using five such engines aiming for a payload gain of around 250 kgs,” Narayanan said.
Narayanan said this would be the best way of mission planning and optimum utilisation of propellants. “All these new things are being done keeping lunar mission in the mind and ISRO’s bigger game plan to increase GSLV payload capability. For Chandrayaan-2, we are formulating a perfect combination. The four strap-ons and second stage will be boosted with high-thrust Vikas engines; cryogenic upper stage will be loaded with enhanced propellants of 15 tonnes instead of current 12.8 tonnes and will be operated with 9.5 tonne thrust compared to the present 7.5.”
Source:
http://www.newindianexpress.com/states/tamil-nadu/2018/mar/28/with-eye-on-lunar-mission-isro-to-test-high-thrust-vikas-engine-1793608.htmlISRO confirms that GSLV F-10 will be first fully upgraded GSLV. All four L40 strapons will have high thrust vikas engines and well as the second stage. The upper stage will have propellant loading increased to 15 tons and thrust of CE-7.5 engine will be increased from present 7.5 tons to 9.5 tons.
Payload capability to GTO for GSLV has been increased by 70kg for GSLV F-08 mission to bring the GTO capability of GSLV up from 2.25 tons to 2.32 tons. With additional upgrades on GSLV F-10 mission later this year this will further boost GTO capability by 250 kg to bring GSLV GTO capability to 2.57 tons.
In short:
Current GSLV GTO capability: 2.25 tons
GSLV F-08 GTO capability: 2.32 tons
GSLV F-10 GTO capability: 2.57 tons
ISRO final goal for GSLV : 3.25 tons to GTO
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#121
by
vineethgk
on 28 Mar, 2018 11:17
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The next generation Vikas engine developed by the Liquid Propulsion Systems Centre (LPSC) is being flown for the first time. LPSC director V Narayanan told Express that the improved engine would give a significant advantage in terms of enhancing payload capability. “Usually, the chamber pressure is 58 bar, but with the use of high-thrust Vikas engine, we will achieve 62 bar, which is a 6% increase in thrust that gives us 70 kgs of additional payload gain in this mission. Right now, we are going to use the high-thrust Vikas engine only in the second stage. Basically, we are validating it. For Chandrayaan-2 mission, we will be using five such engines aiming for a payload gain of around 250 kgs,” Narayanan said.
Narayanan said this would be the best way of mission planning and optimum utilisation of propellants. “All these new things are being done keeping lunar mission in the mind and ISRO’s bigger game plan to increase GSLV payload capability. For Chandrayaan-2, we are formulating a perfect combination. The four strap-ons and second stage will be boosted with high-thrust Vikas engines; cryogenic upper stage will be loaded with enhanced propellants of 15 tonnes instead of current 12.8 tonnes and will be operated with 9.5 tonne thrust compared to the present 7.5.”
Source: http://www.newindianexpress.com/states/tamil-nadu/2018/mar/28/with-eye-on-lunar-mission-isro-to-test-high-thrust-vikas-engine-1793608.html
ISRO confirms that GSLV F-10 will be first fully upgraded GSLV. All four L40 strapons will have high thrust vikas engines and well as the second stage. The upper stage will have propellant loading increased to 15 tons and thrust of CE-7.5 engine will be increased from present 7.5 tons to 9.5 tons.
Payload capability to GTO for GSLV has been increased by 70kg for GSLV F-08 mission to bring the GTO capability of GSLV up from 2.25 tons to 2.32 tons. With additional upgrades on GSLV F-10 mission later this year this will further boost GTO capability by 250 kg to bring GSLV GTO capability to 2.57 tons.
In short:
Current GSLV GTO capability: 2.25 tons
GSLV F-08 GTO capability: 2.32 tons
GSLV F-10 GTO capability: 2.57 tons
ISRO final goal for GSLV : 3.25 tons to GTO
If the payload gain is only about 250 kg even with HTVEs in L40 and CUS upgrades in F10 flight, achieving 3.2-3.3 tonnes as they claimed earlier would appear a long shot. My guess is that either Narayanan was referring only to the effect of addition of HTVEs in L40 and GS2 when he quoted the 250kg figure, or the F10 flight wouldn't have all the major upgrades that they plan to do eventually.
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#122
by
TheVarun
on 28 Mar, 2018 18:25
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Just finished watching the televised launch of the Insat 3Dr on Gslv on Sept 8/2016. One of the Isro scientists mentioned that vehicle had an upgraded Cryogenic engine/ stage for the flight. So the C-12 has been uprated at least once. Any idea what kind of improvement was made then?
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#123
by
K210
on 31 Mar, 2018 10:18
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Just finished watching the televised launch of the Insat 3Dr on Gslv on Sept 8/2016. One of the Isro scientists mentioned that vehicle had an upgraded Cryogenic engine/ stage for the flight. So the C-12 has been uprated at least once. Any idea what kind of improvement was made then?
GSLV F05/INSAT-3DR did have a slightly improved upper stage.
Improvements were:
- C12.5 stage dry mass reduction
- Maximum thrust of CE-7.5 increased from 109% to around 112%
Changes resulted in GTO capacity rising from 2.2 ton to 2.25 ton
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#124
by
K210
on 29 Aug, 2018 12:49
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Umamaheswaran R., now Associate Scientific Secretary, ISRO, told me in October 2016 that GSLV-MkII can put 3.2 tonnes into an initial orbit of 180 km by 20,000 km.
Yes. We enhanced the GSLV-MkII with high thrust engines and so on. With the enhancement, the number is 2.7 tonnes into GTO [geostationary transfer orbit]. Now, 3.2 tonnes has become 3.8 tonnes. And 22,000 km has become 37,000 km. This combination cannot be launched by GSLV-MkII.
Source:
https://www.frontline.in/science-and-technology/article24801393.eceGSLV MK-2 now has GTO capability of 2.7 tons. That is quite a jump from previous 2.25 ton capability.
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#125
by
sanman
on 30 Aug, 2018 16:43
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#126
by
worldtimedate
on 30 Aug, 2018 21:28
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GSLV-F11 cryo engine hot test successfulThe Indian Space Research Organisation (Isro), on Thursday said it has successfully completed the hot test of cryogenic engine for the GSLV-F11, which is scheduled to launch the GSAT-7A in November this year.
A hot test is a ground test conducted to check for the safety and also whether or not all design parameters are met. Unlike a cold test where all the propellants are checked for, a hot test involves firing of the engine in test conditions.
The test, conducted at the Isro Propulsion Complex in Mahendragiri on August 27, was carried out for 200 seconds during which the engine operated in the nominal and 13% uprated thrust regimes.
"All the propulsion parameters during the test were found satisfactory and closely matched with predictions. For the first time, indigenously developed copper alloy is used in this engine," Isro said.
The cryogenic engine will now be integrated with the propellant tanks, stage structures and associated fluid lines to realise the fully integrated flight cryogenic stage.
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#127
by
worldtimedate
on 12 Sep, 2018 19:22
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Indigenisation of Copper-Chromium-Zirconium-Titanium Alloy for Cryo and Semi-Cryo engines- A Success StoryRolling Mill LayoutCopper Alloy (Cu-0.5Cr-0.05Ti-0.05Zr) is an important and vital item required for cryogenic/semi-cryogenic engines for the realisation of thrust chamber inner shell and injector face plates of Cryogenic Upper Stage (CUS) engine for GSLV Mk-II, CE20 engine for GSLV Mk-III and Semi-Cryo (SC) stage. This is also required for the Steering Engines (SE) of CUS engine, Gas generator of CUS & CE20 engines, injectors, pre-burner and pyro components of SC engine.
These projects require Copper Alloy plates, rods and forgings of various dimensions. For plates, thickness requirement range from 12 mm to 18 mm and width of 850 mm. Rods and forgings of this alloy are also required with diameters ranging from 30 mm upto 300 mm.
Vacuum Induction Melting Furnace 1000kgIndigenisation efforts were made through NFTDC, Hyderabad for CUS, CE20 and SC. Melt capacity was augmented to 1000 kg and plate rolling mill capable of 1500 mm width was established for meeting the project requirements. All required products using this alloy have been successfully realized for CUS, CE20 and Semi-Cryo projects. The hot test of the CUS engine using this copper alloy for 200 sec in the nominal and 13% uprated thrust regime was carried out at IPRC, Mahendragiri. This engine will power the cryogenic stage of GSLV Mk-II, which is scheduled to launch GSAT-7A in November this year.
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#128
by
TheVarun
on 17 Sep, 2018 23:13
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Can someone give an approximation for the number of parts in a cryogenic engine/stage like the C-12? Just approximate. Would it be in the thousands or tens of thousands? I'm asking because I want to determine the degree of difficulty a country like India( or any developing one or even developed) would experience in developing such an engine and stage. Also, what degree of indigenisation of the C-12 would have been realistically achieved until now, and what more could be reasonably expected? Again, given the presumably very high number of parts and components.
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#129
by
K210
on 18 Nov, 2023 09:58
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Going over GSLV launch history you have to wonder why ISRO keeps this launch vehicle operational. Out of 15 launches there have been 6 failures of varying degrees putting the failure rate at almost 50%. Now they have LVM3 there is no reason to keep such a unreliable launch vehicle flying.
That high failure rate is largely due to the developmental nature of the GSLV Mk1. In fact there are only 2 failures of the GSLV Mk2 to date (D3 mission in 2010 and F10 in August 2021). For those who don't know the Mk2 variant of this rocket is the operational version which has been flying since 2010.
The GSLV Mk1 as you probably know used imported Russian cryogenic upper stages which apart from not delivering the promised performance resulted in all kinds of design problems the worst of which caused the GSLV F06 RUD over the bay of bengal in December 2010.
I think if we view GSLV as more of a technology testbed of sorts we can have a more positive view of the entire GSLV program. GSLV was the first launch vehicle of ISRO to fly liquid fueled strapons, first to have a cryogenic upper stage, first to employ hot staging and of course first to achieve the Geostationary Transfer Orbit (GTO). First in something is always hard and prone to failure and hence why we should appreciate GSLV's various contributions to the Indian Launch Vehicle program.
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#130
by
Steven Pietrobon
on 19 Nov, 2023 02:13
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GSLV was the first launch vehicle of ISRO to fly liquid fueled strapons, first to have a cryogenic upper stage, first to employ hot staging and of course first to achieve the Geostationary Transfer Orbit (GTO).
The first payload sent into GTO was by PSLV, carrying METSAT 1 on 12 September 2002 (target orbit was 180x36,000 km according to the press kit). The previous GSLV Mk.I launch on 18 April 2001 had a performance shortfall and sent its payload into sub-GTO with a 32,000 km apogee.
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#131
by
zubenelgenubi
on 19 Nov, 2023 18:46
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Moderator:
I merged several GSLV MkII threads. 🚀
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#132
by
Zed_Noir
on 20 Nov, 2023 10:01
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Going over GSLV launch history you have to wonder why ISRO keeps this launch vehicle operational. Out of 15 launches there have been 6 failures of varying degrees putting the failure rate at almost 50%. Now they have LVM3 there is no reason to keep such a unreliable launch vehicle flying.
<snip>
I think if we view GSLV as more of a technology testbed of sorts we can have a more positive view of the entire GSLV program. GSLV was the first launch vehicle of ISRO to fly liquid fueled strapons, first to have a cryogenic upper stage, first to employ hot staging and of course first to achieve the Geostationary Transfer Orbit (GTO). First in something is always hard and prone to failure and hence why we should appreciate GSLV's various contributions to the Indian Launch Vehicle program.
Think it is better for ISRO to phased out the GSLV Mk2 quickly. Since even the LVM3 appears to have an end date on the horizon.
A future "LVM3" with semi-cryogenic core replacing the current hypergolic LVM3 core is a new vehicle even if ISRO stated otherwise.
Unless maintaining the 2 low volume production lines is political expediency. ISRO should consolidated their near future launches with the LVM3 to reduce production infrastructure footprint and get scale of economy with more production volume.
Interesting factoid - the strapped-on hypergolic boosters on the GSLV Mk2 have a longer burn time and higher ISP than the solid first stage.
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#133
by
K210
on 22 Nov, 2023 04:16
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GSLV was the first launch vehicle of ISRO to fly liquid fueled strapons, first to have a cryogenic upper stage, first to employ hot staging and of course first to achieve the Geostationary Transfer Orbit (GTO).
The first payload sent into GTO was by PSLV, carrying METSAT 1 on 12 September 2002 (target orbit was 180x36,000 km according to the press kit). The previous GSLV Mk.I launch on 18 April 2001 had a performance shortfall and sent its payload into sub-GTO with a 32,000 km apogee.
GSLV was the first launch vehicle of ISRO to
attempt to achieve GTO. PSLV was never designed or intended to fly to GTO orbits.
The performance shortfall in GSLV D1 was around 60 m/s or about 0.6% of the intended injection velocity which lead to a 3000-4000 km lower apogee than intended.
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#134
by
K210
on 22 Nov, 2023 04:18
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This paper goes over GSLV D1 mission in detail
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#135
by
zubenelgenubi
on 29 Nov, 2023 16:09
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My opinions:
Re: difficult GSLV MkII development
Hindsight is 20/20. 👀 👓 Congratulations. 🎊 🏅
Re: GSLV MkIII/LVM3 development
Very optimistic development schedule collided with reality, resulting in multi-year delays.
How many launches remain for GSLV? Is there an intended last launch? Or does commonality with the quite successful PSLV imply that there is no end-date yet?
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#136
by
K210
on 30 Nov, 2023 08:23
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How many launches remain for GSLV?
At least 10-12 more launches. GSLV isn't being retired anytime soon.
- 4 GSLV launches for remaining NVS satellites (NVS-2 to NVS-5)
- 1 GSLV launch for INSAT-3DS
- 3 GSLV launches for IDRSS satellites (Data relay satellites for human spaceflight mission)
- 1 GSLV launch for GISAT-2
- 1 GSLV launch for MOM-2 (TBD)
- 1 GSLV launch for Venus Orbiter Mission (TBD)
- 1 GSLV launch for NISAR
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#137
by
K210
on 31 Dec, 2023 09:47
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While GSLV's GTO payload capacity is quite low it is capable of placing more than 3 ton into 700km SSO. This is about double what the PSLV can launch into SSO in its most powerful configuration. Maybe GSLV can be used in future launches to launch heavier SSO payloads beyond what the PSLV can handle.
On some post I wished that ISRO should get at least $3 billion each year (or 18,600 crore) to pursue various missions.
