Solid Rocket Engines: 1

EXTROVERT Space Propulsion 09

Solid Rocket Engines: 1


Solid rocket motors

spIρ

Solid rockets

are simpler and cost less than liquid-fueled rockets

have lower Isp than most liquids (~ 285 sec)

are more dense -> higher “density impulse” . So packaging

is easier.

Cannot be throttled or shut down during the flight (unless

pre-designed to do so)

Unlike liquid rocket engines, the fuel and oxidizer are premixed in solid rocket. The result is a rubbery solid that burns when heated.

Thrust is limited by nozzle size – not by pump capacity. Easy to get very high thrust for boosters.


Δknown V

Applications

- Missiles (acceleration, storage)- Booster, strap ons (high thrust per size)- Apogee kick motors

EXTROVERT Space Propulsion 09Star-Grained Solid Rocket Motor

http://www.nf.suite.dk/stargrain/ After 1 minute of burn


General configurationTE-M-364-4 is a 15,000 lb. thrust solid propellant motor developed for use as an upper stage. It is an enlarged version of the TE-M-364, one of a series of solid propellant motors that powered the workhorse USAF Burner I and Burner IIA upper stages to orbit scientific, weather, navigation, and communications satellites. The TE-M-364-4 powered the upper stages of the USAF Atlas boosters used to launch the Global Positioning System (GPS) satellites. It also was used as the second stage motor on USAF Thor vehicles that launched satellites of the Block 5D Defense Meteorological Satellite Program (DMSP) as well as the third stage motor on the Thor Delta launch vehicles. www.wpafb.af.mil/ museum/engines/eng62.htm

http://www.wpafb.af.mil/museum/outdoor/od10.htm

http://www.wpafb.af.mil/museum/engines/eng62.htm

http://www.wpafb.af.mil/museum/engines/eng62.htm


history.nasa.gov/ rogersrep/v1p56.htm

http://history.nasa.gov/rogersrep/v1p56.htm

EXTROVERT Space Propulsion 09“STS SRB motors

SRB motor: propellant mixture ammonium perchlorate (oxidizer, 69.6 percent by weight), aluminum (fuel, 16 percent), iron oxide (a catalyst, 0.4 percent), a polymer (a binder that holds the mixture together, 12.04 percent), and an epoxy curing agent (1.96 percent). The propellant is an 11-point star- shaped perforation in the forward motor segment and a double- truncated- cone perforation in each of the aft segments and aft closure. This configuration provides high thrust at ignition and then reduces the thrust by approximately a third 50 seconds after lift-off to prevent overstressing the vehicle during maximum dynamic pressure.”liftoff.msfc.nasa.gov/ Shuttle/About/detsrb.html

http://liftoff.msfc.nasa.gov/Shuttle/About/detsrb.html



“The SRBs are used as matched pairs and each is made up of four solid rocket motor segments. The pairs are matched by loading each of the four motor segments in pairs from the same batches of propellant ingredients to minimize any thrust imbalance. The segmented-casing design assures maximum flexibility in fabrication and ease of transportation and handling. Each segment is shipped to the launch site on a heavy- duty rail car with a specially built cover.”

liftoff.msfc.nasa.gov/ Shuttle/About/detsrb.html



“The forward section of each booster contains avionics, a sequencer, forward separation motors, a nose cone separation system, drogue and main parachutes, a recovery beacon, a recovery light, a parachute camera on selected flights and a range safety system.

Each SRB has two integrated electronic assemblies, one forward and one aft. After burnout, the forward assembly initiates the release of the nose cap and frustum and turns on the recovery aids. The aft assembly, mounted in the external tank/SRB attach ring, connects with the forward assembly and the orbiter avionics systems for SRB ignition commands and nozzle thrust vector control. Each integrated electronic assembly has a multiplexer/ demultiplexer, which sends or receives more than one message, signal or unit of information on a single communication channel.”




The nozzle is gimbaled for thrust vector (direction) control. Each SRB has its own redundant auxiliary power units and hydraulic pumps. The all-axis gimbaling capability is 8 degrees. Each nozzle has a carbon cloth liner that erodes and chars during firing. The nozzle is a convergent- divergent, movable design in which an aft pivot- point flexible bearing is the gimbal mechanism.

The cone- shaped aft skirt reacts the aft loads between the SRB and the mobile launcher platform. The four aft separation motors are mounted on the skirt. The aft section contains avionics, a thrust vector control system that consists of two auxiliary power units and hydraulic pumps, hydraulic systems and a nozzle extension jettison system.

Eight booster separation motors (four in the nose frustum and four in the aft skirt) of each SRB thrust for 1.02 seconds at SRB separation from the external tank. Each solid rocket separation motor is 31.1 inches long and 12.8 inches in diameter.




www.wstf.nasa.gov/.../ Explosion/HEBFTesting.htm

Solid rocket Explosion: Large fragmentscreated

http://www.wstf.nasa.gov/Hazard/Explosion/HEBFTesting.htm




www.aero.org/.../crosslink/ winter2003/08.html

Inertial Upper Stage

W. Paul Dunn

http://www.aero.org/publications/crosslink/winter2003/08.html

http://www.aero.org/publications/crosslink/winter2003/contributors.html%2301


Stinger Unofficial names/slang: n/a Function: To provide close-in, surface-to-air weapons for the defense of forward combat areas, vital areas and installations against low altitude air attacks. Date deployed: 1987 Contractor: General Dynamics /Raytheon Unit cost: $38,000 Length: 5' - 0" Wingspan: 3.5" Diameter: 0' - 0" (0.00m) Speed: Supersonic Weight at launch: 34.5 lbs (launcher w/ missile) Guidance: Fire-and-forget passive infrared seeker Range: approx. 1 - 8 km Engine: Dual thrust solid fuel rocket motor Warhead: High explosive

www.combatindex.com/.../ detail/mis/stinger.html

Stinger Man-Portable S-A Missile

http://www.combatindex.com/hardware/detail/mis/stinger.html


Solid Propellants

Double Base – molecules of fuel/oxidizer are mixed (e.g., gun powder dissolved in nitroglycerine) – oxygen in both (less common, more explosive)

Composite – Heterogeneous mixture of fuel, oxidizer and binder, plus some other additives – more common.


Fuels←Powdered Aluminium STS

Powdered Mg

PBAN RSRM←

HTPB ←

Binders

most popular now

The binder holds the entire formulation in a structurally sound propellant grain, under temperature and pressure variations, plus accelerations and vibration loads of flight.

Binders should have low density and energy of combustion, plus structural integrity using minimal binder volume.

“Solids Loading” = percentage the total propellant mass taken up by fuel + oxidizer. Usually > 90%Binders are usually long-chain polymers – keep the propellant powdersand crystals in a continuous matrix through polymerizing andcross-linking.

EXTROVERT Space Propulsion 09Oxidizers

• Ammonium Perchlorate (AP) – contains chlorine – acid rain • Ammonium Nitrate (AN) is more benign. But inherently low burning rate and a phase change near 30 deg. C.


Other Ingredients

Fixers (bonding agents): improve bond between oxidizer and binder

Curatives: increase rate of polymerization.

Plasticizer: improve physical properties at low temperatures

Darkening agents: reduce thermal radiation losses through translucent propellant

HMX: increases burning rate. Can cause detonations too.


Propellant Burning Rate

= ncr aP

( ) ( )=ln ln cr n aP

regression rate proportional to pressure to some nor

Regression Law – St. Robert’s Law

A “plateau” type burning rate law is more common, where n becomes closeto zero over a range of pressure.

Note: n has to be < 1 for stability

&,r m

cP

&m

> 1n< 1n


Grain cross sections to control burning

• End grain: neutral

• Internal Burning Tube: progressive

• Internal-External Burning Tube: neutral

• Rod and Tube: neutral

• Internal Burning Star: neutral

• Dog Bone: neutral

• Slots and Tube: neutral

• Slotted Tube: neutral

• Wagon Wheel: neutral

• Multiple Perforations: neutral

• Neutral thrust history generally gives the smallest inert mass since

the maximum and average pressures on the structure are nearly

the same with this. Else use regressive thrust profiles.


Simple Solid Rocket Analysis

In a solid rocket motor, the “chamber” pressure is related to the geometry and burn rate. Therefore we must know something about the geometry to find Pc (time) and thus thrust and Isp vs. time.

(Simple end-burner design)

= +. . .burn volume exitm m m

( ) ( )ρ ρ∂= +∂

.exitb p c cA t r V m

t

Then from conservation of mass:

.pm

(mass released from surface per unit time = mass added to growing chamber volume + mass exhausted)

rAb

pc .exitm

…..(1)

….(2)

ρpρc

Lweb

density of solid

density of gas in bore

propellant mass flow rate


= = =*

. .0

0 0

F c t Fsp

p p

T C P A C CIg

g m g mRecall

Δ= =

. .

*c t exitPPm mC

….(3)

( ) ρρ ρ∂ ∂

= + +∂ ∂ *c c t

b p c cV P AA t r V

t t C

= ncr aP

or

So

where (St. Robert’s Law) …..(4)

EXTROVERT Space Propulsion 09 Let us first assume:

0ctρ∂

=∂

1) in equilibrium

ρ ∂∂cVt ρc

( )bA t

= ncr aP ρ *,p C

( )ρ ρ= =* *

b c cnt P c p

A P PA rC aP C

2) is small (note that is a gas density)

3) is constant (end burner type design)

4) (n < 1) (and “a”, n, do not change over time)

then ..


(This is a steady-state approximation)

(steady-state lumped-parameter) where

ρ

→

→

→

→

*

3

c

n

P

P MPacm

saMPamC sKgm

Pay particular attention to units in (5) if you are going to use Humble’s table of “a” in Table 6.9.

We can use equation (5) to estimate the size of an end-burner for a desired Pc and performance.

………(5)

€

AbAt

= PcaPc

nC*ρ p


Example

= ←4cP Mpa target

=* 1500 /C m s=1.85fvC ≈1.2f(note for solids)

ρ = 31800 /P Kg m

[ ]= .3.40 cr P in cm/s (Pc =MPa)

=500,000VacT N =100secbt

, ,t bA A weblIf the desired and

Find and =

= 6500000 1.85 4 *10f t c

t

T C A P

N A Pa

tA

[ ]{ }( )= =

= =

.3.40 4 100sec

60.629 .6063

web b

web

l rt MPa

l cm m

=.0676m2


from (5)

[ ]{ }( )( )= ⎛ ⎞⎛ ⎞⎜ ⎟⎜ ⎟⎝ ⎠⎝ ⎠

.3 36

41 1.4 4 1500 / 1800 /

100 1*10

b

t

A MPam MPaA MPa m s Kg mcm Pa

=244.35b

t

AA

So ( )= 2244.35 .0676bA m

= 216.51bA m

andπ

= * 2bb

AD

=2.293bD m

=.132web

b

LD

If this geometry is unacceptable, we can change Pc and resize. For example, a higher Pc will make a longer, more slender solid rocket.


Time varying Burn Area

For a more general cross section (tube, star, wagon wheel, etc) we would expect the cross-sectional area or the total exposed burn area to change with time.

Given the initial geometry

= =nc

dxr aPdt

and

ρ−⎡ ⎤⎛ ⎞ ⎛ ⎞⎛ ⎞=⎢ ⎥⎜ ⎟ ⎜ ⎟⎜ ⎟⎝ ⎠⎝ ⎠⎝ ⎠⎣ ⎦

11*

61 1

100 1*10

nbc p

t

A m MPaP aCA cm Pa


But and Pc can change with time

Let and be fixed

Then

Time varying Burn Area

ρ*, , ,pa CIn the units we have been using for each, where typically assume and n are constant.

, ,b tA A

( )=b bA A xtA

( )ρ

−⎧ ⎫⎡ ⎤ ⎛ ⎞⎪ ⎪⎛ ⎞ ⎛ ⎞= ⎨ ⎬⎢ ⎥⎜ ⎟ ⎜ ⎟⎜ ⎟⎝ ⎠ ⎝ ⎠⎝ ⎠⎪ ⎪⎣ ⎦⎩ ⎭1*

61 1 1

100 100 1*10

nnb

pt

A xdx m m MPaa aCdt cm A cm Pa


( ) ( )ρ

−⎧ ⎫⎡ ⎤ ⎛ ⎞⎪ ⎪⎛ ⎞ ⎛ ⎞= − = ⎨ ⎬⎢ ⎥⎜ ⎟ ⎜ ⎟⎜ ⎟⎝ ⎠ ⎝ ⎠⎝ ⎠⎪ ⎪⎣ ⎦⎩ ⎭∫ ∫1*

60

1 1 1100 100 1*10

nx t nb

i ptxi

A xm m MPadx x x a aC dtcm A cm Pa

(at x=xf, t=tb)

for most complex shapes, we will need to integrate the R.H.S. numerically,and may be a complex calculation over multiple regions. ( )bA x

For a simple shape

where L= bore length

( ) π=2bA x xL

x


So

( )π ρ −

−

⎧ ⎫⎪ ⎪= ⎨ ⎬⎪ ⎪⎩ ⎭

∫ ∫1*

801

2100 1*10

nnx t

pn

txi n

aC Ldx a dtA

x

( )A t

( )c t fP t AC

bt = maxx x

Integrating this (left to HW) gives

X(t) - regression amount as a function of time and therefore

Thrust (t) =

And the total burn time, , can be calculated for when

Pc(t)

Solid Rocket Engines: 1

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