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#26 2017-06-23 12:38:07

GW Johnson
Member
From: McGregor, Texas USA
Registered: 2011-12-04
Posts: 2,543
Website

Re: Amateur solid-fueled rockets to *orbital* space?

Getting good thrust coefficient CF at low Pc in vacuum is why the vacuum Isp data in my old Pratt & Whitney handbook are for Pc = 100 psia,  while the sea level data are for 1000 psia.  Thrusters in vacuum are usually quite low chamber pressure,  so that the hardware is thinner and lighter. 

Same could be true of upper stage engines,  but that choice depends more on where that engine design came from: Spacex's second stage Merlin is just the first stage Merlin fitted with a much longer nozzle.  They both use the same chamber design. 

That low-pressure lightening-of-weight proves to be more significant than low pressure c* losses,  at least for the kind of equipment specifically designed to take advantage of the effect. 

Thruster expansion ratio is usually limited by packaging/installation issues than anything else,  and the nozzles are quite often scarf-cut reflecting the outer mold line of the vehicle.  This applies to mono- or bi-propellant thrusters.  Basic CF applies to liquids and solids and hybrids. 

I ran my spreadsheet version of thrust coefficient at a fixed Pc = 200 psia,  and reset Pe iteratively until I got Ae/At = 750.9 at Pe = 0.0108 psia.  The CF for Pe = Pamb at that expansion is 1.98047.  CF vac is 2.0226.  These were figured for specific heat ratio 1.20 and a 15 degree effective conical half angle to the nozzle,  pretty "typical" values. 

Thrust per unit throat area is Pc CF = 404.52 psi,  using the vacuum CF. 

Can't tell you what the thrust is,  without some way to set throat area.  Normally that is sized to get the thrust you want out of the stage,  or else it comes from an existing design.  Cannot tell you what Isp is until I know something about a 200 psia chamber c*. 

If for an aluminized solid based on AP-HTPB the 200 psia c* is near 4800 ft/sec,  then Isp = CF c* / gc ~ 302 sec.  If instead it was 4900 fps,  then Isp ~ 308 sec.  If it is nearer 5000 fps,  then Isp ~ 314 sec. 

Chamber c* is a power function of Pc of the form c* = k Pc^m,  where for solids m ~ .01,  although the variation of m from propellant to propellant is significant.  At 200 vs 1000 psia,  for m = 0.01,  we lose about 2% of our c*. 

If c* were known,  then you could figure the propellant flow through the throat per unit throat area:  w/At = Pc gc / c*.

Hope that helps. 

GW

Last edited by GW Johnson (2017-06-23 16:02:17)


GW Johnson
McGregor,  Texas

"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#27 2017-06-24 07:31:53

RGClark
Member
From: Philadelphia, PA
Registered: 2006-07-05
Posts: 430
Website

Re: Amateur solid-fueled rockets to *orbital* space?

GW Johnson wrote:

Getting good thrust coefficient CF at low Pc in vacuum is why the vacuum Isp data in my old Pratt & Whitney handbook are for Pc = 100 psia,  while the sea level data are for 1000 psia.  Thrusters in vacuum are usually quite low chamber pressure,  so that the hardware is thinner and lighter. 
Same could be true of upper stage engines,  but that choice depends more on where that engine design came from: Spacex's second stage Merlin is just the first stage Merlin fitted with a much longer nozzle.  They both use the same chamber design. 
...
GW

Do you know of examples of those low chamber pressure solid thrusters at Pc = 100 psia? For my upper stage of the all-solid launcher, I want to use one of the size of the Cesaroni motor used on the Up Aerospace suborbital rocket I mentioned in post #13 with a 412 pound fueled weight and 0.8 propellant fraction. This puts is it at about 80 pound empty weight and 320 pound propellant weight. If I got the chamber pressure smaller by a factor of 10, then I could reduce that empty weight to only 8 pounds, and the payload would increase by 72 pounds.

But this means I want to keep the same propellant weight, and use the same size case but only thinner wall thickness by a factor of 10. But I don't know if this is possible for a solid-fueled motor. For a liquid I could just reduce the propellant burn rate to reduce the chamber pressure, i.e., less propellant burn at any time, less pressurized gases produced at any time, therefore lower chamber pressure.

But decreasing the chamber pressure is not so easy to do with a solid. It won't work to just increase the case volume because that eliminates the advantage you wanted to get of having a smaller empty, i.e., case weight. You can't reduce the propellant amount either to get a larger volume for the central open section because you want the same size stage.

Is there some method to get the same amount of propellant to just burn at a slower rate? Another possible way it could work would be by having the propellant compressed to an extreme degree to only 1/10th it's usually volume. Don't know if that is technically feasible.

  Bob Clark


Nanotechnology now can produce the space elevator and private orbital launchers. It now also makes possible the long desired 'flying cars'. This crowdfunding campaign is to prove it:
Nanotech: from air to space.
https://www.indiegogo.com/projects/nano … 13319568#/

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#28 2017-06-24 12:41:41

GW Johnson
Member
From: McGregor, Texas USA
Registered: 2011-12-04
Posts: 2,543
Website

Re: Amateur solid-fueled rockets to *orbital* space?

You have to design the motor specifically to be a low-pressure motor.  Usually,  that will be a different grain design as well as a lower burn rate.  It is not possible to compress solid propellant;  it simply is what it is. 

Most of the propellants historically have not exhibited stable burn rate behavior under about 300 psia.  There are many different failure modes.  I have seen that limit go lower with AP-HTPB-type composites:  around 100 psia.  Very few work well below that,  but I have actually seen a few very fuel-rich formulations (very low Isp) work at 30-50 psia.

Burn rates with AP composites more-or-less fall in the 0.2 to 0.3 in/sec range at about 400 psia.  Ordinary ones have burn rates in the 0.3 to 0.4 in/sec range at 2000 psia (low exponent).  The "high rate" composites can have burn rates in the 0.8 to 1.0 in/sec range at 2000 psia (high burn rate exponent).  So there's room for a lot of burn rate tailorability by adjusting particle size distributions for the solids,  and in using iron-bearing burn rate catalysts.  Most of these propellants fall in the vicinity of 0.06-something lb/cu.in density. 

Sorry,  there's no simple "low pressure retrofit" to an existing motor design,  other than just making throat area bigger.  Each application is its own custom internal ballistic design.  That's just the nature of solids.  Grain design surface history is only part of that. 

There are limits on the web/motor radius ratio associated with grain thermal-structural,  limits on grain port to nozzle throat area ratio,  limits on allowable pressure drop down the bore,  etc.  There are also limits on the burn rates,  temperature sensitivity of that burn rate,  and thermal-structural properties (including very serious visco-elastic effects hot,  and glassy-transition brittleness cold). 

The low pressure motor you would like probably does not exist within that set of case dimensions,  but also probably could.  It would have to be specifically designed,  and its propellant specifically tailored to low burn rate in the lab,  so that it could use similar nozzle throat dimensions and expansion proportions.  Otherwise,  your exit is limited to the stage diameter,  period (excepting some sort of aerospike,  perhaps,  but I haven't seen much benefit to them in vacuum conditions over a physical bell).  Whether this low pressure design would have exactly the same case/propellant ratios is problematical,  but they probably could be made similar. 

As for low-pressure thrusters,  I have no specific example systems to name.  Such data could be searched-for. 

GW

Last edited by GW Johnson (2017-06-24 12:44:27)


GW Johnson
McGregor,  Texas

"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#29 2017-08-16 19:11:37

RGClark
Member
From: Philadelphia, PA
Registered: 2006-07-05
Posts: 430
Website

Re: Amateur solid-fueled rockets to *orbital* space?

GW Johnson wrote:

Lessee,  a 5 sec burn at a nominal 0.3 in/sec for AP-composites yields a web burned of 1.5 inches,  about 50% of the radius of a 6-inch motor.  That's definitely in the ballpark for the thrust-time curves you posted.  I cannot pin down a pressure from this,  as burn rates are tailorable from 0.2 to 1+ in/sec at 1000,  and pressure exponents are typically quite low (near 0.3). 
Reducing motor pressure from 1000 psia to around 500 psia,  with a 14.7 psia backpressure,  reduces pressure ratio from 60-ish to 30-ish,  which reduces CF from near 1.5 to nearer 1.4.  That reduces c* from just over 5000 ft/sec to around 4800-4900 ft/sec,  which is just about what I computed from the tabular thrust and impulse data from their propellant weight. 
I would hazard the guess their average chamber pressure is nearer 400-500 psia,  which brings down Isp into the range they quote.  Most of the military hardware I worked on had chamber pressures above 1500 psia,  and we typically got the higher Isp values. 
...
GW

I found this commercial solid motor with a similar sea level Isp as the Cesaroni solid rocket motors Pro150 I mentioned earlier:

Star 37.
Thiokol solid rocket engine. Total impulse 161,512 kgf-sec. Motor propellant mass fraction 0.899. First flight 1963. Solid propellant rocket stage. Burner II was a launch vehicle upper stage developed by Boeing for the Air Force Space Systems Division. It was the first solid-fuel upper stage with full control and guidance capability developed for general space applications.
AKA: Burner 2;TE-M-364-1. Status: First flight 1963. Number: 180 . Thrust: 43.50 kN (9,779 lbf). Gross mass: 621 kg (1,369 lb). Unfuelled mass: 63 kg (138 lb). Specific impulse: 260 s. Specific impulse sea level: 220 s. Burn time: 42 s. Height: 0.84 m (2.75 ft). Diameter: 0.66 m (2.16 ft).

Thrust (sl): 33.600 kN (7,554 lbf). Thrust (sl): 3,428 kgf.
http://www.astronautix.com/s/star37.html

So we might estimate the vacuum Isp of the Cesaroni to be in the Star 37's range of 260 s.

Using this as a vacuum Isp of a first stage and a 285 s vacuum Isp of second and third stage I get a three-stage solid rocket able to reach orbit, with not even particular high mass ratios of 5 to 1 for the stages:

standard_nozzles_ver2.jpg

With a payload of 7 kg:

standard_nozzles_results.jpg


And this with standard nozzles, no altitude compensation required.

The Cesaroni Pro150 retails for about $3,000 and in general the Cesaroni solids cost in the range of $100 per kg of the motor mass:

Cesaroni O8000 White Thunder Rocket Motor.   
$3,099.95
Product Information
Specification
Brandname:  Pro150 40960O8000-P                 Manufacturer:  Cesaroni Technology
Man. Designation:  40960O8000-P                    CAR Designation:  40960 O8000-P
Test Date:  4/10/2008                                     
Single-Use/Reload/Hybrid:  Reloadable              Motor Dimensions mm:  161.00 x 957.00 mm (6.34 x 37.68 in)
Loaded Weight:  32672.00 g (1143.52 oz)         Total Impulse:  40960.00 Ns (9216.00 lb/s)
Propellant Weight:  18610.00 g (651.35 oz)       Maximum Thrust:  8605.10 N (1936.15 lb)
Burnout Weight:  13478.00 g (471.73 oz)          Avg Thrust:  8034.50 N (1807.76 lb)
Delays Tested:  Plugged                                    ISP:  224.40 s
Samples per second:  1000                               Burntime:  5.12 s
https://www.csrocketry.com/rocket-motor … motor.html

So take the cost of the third stage, derived from the Cesaroni Pro1050, as $3,000, and the second stage 4 times larger as $12,000, and the first stage larger by an additional factor of 4 as $48,000. So $63,000 for a small sat launcher with a 7 kg payload to orbit.

  Bob Clark

Last edited by RGClark (Yesterday 02:44:14)


Nanotechnology now can produce the space elevator and private orbital launchers. It now also makes possible the long desired 'flying cars'. This crowdfunding campaign is to prove it:
Nanotech: from air to space.
https://www.indiegogo.com/projects/nano … 13319568#/

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#30 2017-08-16 19:23:03

louis
Member
From: UK
Registered: 2008-03-24
Posts: 2,299

Re: Amateur solid-fueled rockets to *orbital* space?

I've looked into this before...can't remember any real details but I think you could get one or two people off the surface of Mars with a few tonnes (maybe 5-8 tonnes) of fuel/propellant. Think it was 3 tonnes for the Moon...but technology moves on.


Let's Go to Mars...Google on: Fast Track to Mars blogspot.com

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#31 Yesterday 02:42:21

RGClark
Member
From: Philadelphia, PA
Registered: 2006-07-05
Posts: 430
Website

Re: Amateur solid-fueled rockets to *orbital* space?

louis wrote:

I've looked into this before...can't remember any real details but I think you could get one or two people off the surface of Mars with a few tonnes (maybe 5-8 tonnes) of fuel/propellant. Think it was 3 tonnes for the Moon...but technology moves on.

I thought of using small solid motors to do a Mars sample return mission. It would seem to cost a small fraction of NASA's $10 billion estimate for such a mission.

But the calculation in that last post was to suggest any university would have the capability of conducting their own smallsat launch to LEO.

  Bob Clark

Last edited by RGClark (Yesterday 02:46:20)


Nanotechnology now can produce the space elevator and private orbital launchers. It now also makes possible the long desired 'flying cars'. This crowdfunding campaign is to prove it:
Nanotech: from air to space.
https://www.indiegogo.com/projects/nano … 13319568#/

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#32 Yesterday 16:41:40

SpaceNut
Administrator
From: New Hampshire
Registered: 2004-07-22
Posts: 10,075

Re: Amateur solid-fueled rockets to *orbital* space?

http://nasawatch.com/archives/2017/08/g … -on-i.html

GAO: Surplus Missile Motors: Sale Price Drives Potential Effects on DOD and Commercial Launch Providers

GAO estimated that DOD could sell three Peacekeeper motors—the number required for one launch, or, a "motor set"—at a breakeven price of about $8.36 million and two Minuteman II motors for about $3.96 million, as shown below. Other methods for determining motor prices, such as fair market value as described in the Federal Accounting Standards Advisory Board Handbook, resulted in stakeholder estimates ranging from $1.3 million per motor set to $11.2 million for a first stage Peacekeeper motor.

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#33 Today 01:47:12

RGClark
Member
From: Philadelphia, PA
Registered: 2006-07-05
Posts: 430
Website

Re: Amateur solid-fueled rockets to *orbital* space?

SpaceNut wrote:

http://nasawatch.com/archives/2017/08/g … -on-i.html

GAO: Surplus Missile Motors: Sale Price Drives Potential Effects on DOD and Commercial Launch Providers

GAO estimated that DOD could sell three Peacekeeper motors—the number required for one launch, or, a "motor set"—at a breakeven price of about $8.36 million and two Minuteman II motors for about $3.96 million, as shown below. Other methods for determining motor prices, such as fair market value as described in the Federal Accounting Standards Advisory Board Handbook, resulted in stakeholder estimates ranging from $1.3 million per motor set to $11.2 million for a first stage Peacekeeper motor.


Article doesn't say how much these motors repurposed as orbital launchers could loft to orbit. Anyone know that?

  Bob Clark


Nanotechnology now can produce the space elevator and private orbital launchers. It now also makes possible the long desired 'flying cars'. This crowdfunding campaign is to prove it:
Nanotech: from air to space.
https://www.indiegogo.com/projects/nano … 13319568#/

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#34 Today 16:51:02

SpaceNut
Administrator
From: New Hampshire
Registered: 2004-07-22
Posts: 10,075

Re: Amateur solid-fueled rockets to *orbital* space?

you are in luck as these are already being used for the https://en.wikipedia.org/wiki/Minotaur_(rocket_family) derived from converted Minuteman and Peacekeeper intercontinental ballistic missiles.

https://en.wikipedia.org/wiki/LGM-118_Peacekeeper

Engine     Three-stage solid-fuel rocket.
First stage: 500,000 lbf (2.2 MN thrust) Thiokol SR118 solid-fuel rocket motor
Second stage: Aerojet General SR119 solid-fuel rocket motor
Third stage: Hercules SR120 solid-fuel rocket motor


https://en.wikipedia.org/wiki/LGM-30_Minuteman

Engine     Three-stage solid-fuel rocket engines
First stage: Thiokol TU-122 (M-55);
second stage: Aerojet-General SR-19-AJ-1;
third stage: Aerojet/Thiokol SR73-AJ/TC-1

Current https://en.wikipedia.org/wiki/Minotaur_V

The Minotaur V is a five-stage vehicle, and is designed to place up to 630 kilograms (1,390 lb) of payload into a geosynchronous transfer orbit, or 342 kilograms (754 lb) on a trans-lunar trajectory. It consists of a Minotaur IV+, with a Star-37 as a fifth stage.

Two variants are available, one with a spin-stabilized Star-37FM upper stage, and the other with a Star-37FMV capable of three-axis stabilization. The Star-37FMV upper stage is heavier, reducing payload capacity, but is more maneuverable.

  • First stage – SR-118 Engines    
    Thrust     1,607 kilonewtons (361,000 lbf)
    Burn time     83 seconds

    Second stage – SR-119 Engines
    Thrust     1,365 kilonewtons (307,000 lbf)
    Burn time     54 seconds

    Third stage – SR-120 Engines    
    Thrust     329 kilonewtons (74,000 lbf)
    Burn time     62 seconds

    Fourth stage – Star-48BV Engines    
    Thrust     64 kilonewtons (14,000 lbf)
    Burn time     84 seconds

    Fifth stage (Baseline) – Star-37FM Engines    
    Thrust     47.26 kilonewtons (10,620 lbf)
    Burn time     63 seconds

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