1998-10 r.lo m.d. cryogenic solid propellant grains
TRANSCRIPT
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IAF-98-S.3.10
Modular Dissected Cryogenic Solid-Rocket Propellant GrainsR o g e r E . L o , B e r l i n U n i v . T e c h n . , A e r o s p a c e I n s t i t u t e < R o g -
e r . L o @ T U - B e r l i n . d e >
IntroductionThe concept of cryogenic solid-rocket propellant
(CSP-) grains - see (1) - is based on the idea that any
conceivable chemical rocket propellant combination
can be used as solid rocket propellant if the definition
of solid propellants is not restricted to ambient temper-
ature. Liquid bipropellants and tripropellants are thus
transformed into solid propellants, so are hybrids, slur-
ries and quasi-hybrids (i.e. solid mono- or bi-propellants with Hydrogen injection). Consider using
frozen Oxygen (SOX) instead of Ammonium Perchlo-
rate along with some solid hydrocarbon fuel (HC).
What results is a solid propellant SOX/HC with an Isp
roughly equal to the semicryogenic hybrid
LOX/Polyethylene that in turn corresponds to the well
known liquid combination LOX/Kerosene with an Isp
of 300s (E, 68:1) that exceeds any conventional solid
propellant.
Consider a high energy solid grain using
SOX/SH2. The first objection coming to mind against
such a concept or similar ones is the danger of combus-tion instability up to detonation. Many quasi-
homogeneous mixtures of cryogenic solid fuels and
oxidizers might be very well behaved, but this is a
concern for the majority of propellant combinations.
The cure lies in transforming deflagration into boun-
dary layer combustion by dissecting the grain into a
number of homogeneous subunits. Imagine cylindrical
elements with a central combustion channel of arbitrary
shape. In hybrid propulsion, a solid fuel is consumed
by having a fluid oxidizer flowing along its surface.
The oxidizer source is o u t s i d e the combustion
chamber. In the case of dissected solid propellant
grains the source is a melting and/or evaporating solidoxidizer located i n s i d e the same combustion
chamber. Therefore the arrangement can be called an
"internal hybrid burner". Alternating layers of oxidizer
and fuel yield multiple boundary layer combustion.
Obviously, the leading element in such a stack needs to
be an oxidizer rich gas generator module capable of
self sustained burning.
The concept of modular grains can be generalized to
compounds of modules of any suitable shape and com-
position. Elements can be 100% oxidizer or fuel as well
as oxidizer- or fuel-rich gas generators.
--------------------------------------------------------------------Copyright: 1998 by Roger E.Lo. Published by the American
Inst.of Aeronautics and Astronautics, Inc. with permission. Released
to IAF/IAA/AIAA to publish in all forms.
CSPs combine the advantages of liquid and solid
propulsion: high performance in terms of Isp and low
structural mass per unit total weight.
The scope of chemical propul-sion performance
Chemical propulsion is based on chemical reactions
that range from monopropellant decomposition to metal
combustion. A widely used convention divides attaina-
ble Isps into a high-energy and a low energy domain,
with a narrow medium energy domain around 300 s in
between (see Fig.1).
*At the lower end, liquid monopropellants form thebasis for hot gas AOC thrusters.
* Conventional solid propellants are quasi-homogeneous mixtures of solid fuels, oxidizers and
other constituents, stored in a container that later be-
comes the combustion chamber. Defining as one in-
dividual component every constituent that requires aseparate container, such propellants are solid mono-
propellants. Restriction to solid state at room temper-
ature sets tight limits to the choice of suitable matters.
As a result, even so called "high energy solid rocket
propellants" are no more than medium energy on the
absolute scale used in Fig.1. However, robust state-
of-the-art solid propellants are low-energy.
* Medium energy liquid bipropellants occupy the in-termediate realm, with storable hypergolics at the
lower, semi-cryogenic combinations (e.g.
LOX/hydrocarbon) at the upper end.
* High energy (HE) liquid bipropellants includeLOX/LH2 the propellant combination marking the
upper limit of chemical state-of-the-art propulsion
with 391s at 68:1 E. To go beyond this mark would
require using better oxidizers and/or fuels. Fluorine
and ozone are known to yield higher Isps than Oxy-
gen, but are not used because they are either too ag-
gressive or too explosive. There are several metals
that yield more combustion energy per unit mass of
products than Hydrogen does. Using them in bipro-
pellants requires slurries or gels that cause produc-
tion, storage and feeding problems. However, slurry
bipropellant combinations such as LOX/ Beryllium/Hydrogen would perform close to the very limit of
conventional chemical propulsion (1).
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Fig.1: Isp range of chemical propulsion
*Hybrids are liquid/solid bipropellants. Isp-performance covers a broad range from low- to high-
energy. Hypergolic combinations such as
HNO3/Amine-fuel (not shown in Fig.1) are found at
the lower end. Semi-cryogenic hybrids include envi-
ronmentally benign, medium energy combinations
(e.g. LOX/Polyethylene, PE, not shown). The solid
fuel in regular hybrids lends itself for adding metals.
Among others, the formation of the Oxides and Fluo-
rides of Beryllium, Lithium, Boron, and Aluminum
yields the highest known mass-specific heats of for-
mation. Powders of these metals can be mixed with
solid fuels or can be introduced as hydrides if suchexist in solid state. Thus, regular hybrids cover a
broad range on the scale of Isps.
Metal combustion does not make much sense without
molecular weight reduction. This can be achieved by
an excess of Hydrogen. Metal-rich solid propellants
used for Hydrogen heating are a special category of
bipropellants and were dubbed "Quasi-hybrids" (2).
Using extra Hydrogen in addition to some other fuel,
one obtains three component combinations also
called tripropellants. Hybrids with highly metalized
fuel grain are excellent means for heating excess Hy-
drogen (3) and were called tribrids. Most of the met-
als in question can easily reduce H2O under combus-
tion chamber conditions. Hence it can be demonstrat-
ed that the combustion products at maximum Isp mix-
ture ratios are free of Hydrogen combustion productseven in very complex mixtures (4)(5). Using liquid
Ozone rather than LOX and BeH2 rather than Be
(both are endothermic compounds) one obtains prob-
ably the highest known Isp of conventional chemical
propellants.
* The concept of modular dissected cryogenic solid-rocket propellant grains (CSPs) provides a means for,
in principle, realizing a l l propellant combinations
as solid propellant motors. This includes, under cer-
tain provisions, propellants that are hypergolic liquids
at ambient temperature. If Hydrogen is to be used,
temperatures well below 14K (25R) are required.This kind of cryogenic solids is the main subject of
investigations at Berlin University of Technology's
Aerospace Institute (6)(7)(8).
Super High Energy Propellants (SHEPs) employ
metastable chemical species. Known examples are
atomic Hydrogen and other atomic species, free radi-
cals and excited species. If storable, they could push
the upper limit of chemical propulsion by a factor of 4
to 5 beyond the bounds shown in Fig.1. Matrix isola-
tion is one method of storing SHEPs. In the case of
some of the more sensitive species it is believed that
matrix isolation needs extremely low temperaturesalong with strong magnetic fields. Cryogenic solids are
being studied at USAF in the High Energy Density
Matter Program, which considers solid Hydrogen
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doped with SHEPs and burned in hybrid manner with
liquid Oxygen, see (10). Such elements could easily be
included in the modular CSPs suggested in this article.
Special cryogenic high energypropellantsS O X / S H 2
CSPs are much colder than liquids and hence con-
tain less energy resulting in lower Isp performance.
However, a more detailed calculation shows that the
missing heats of liquefaction are small compared with
the overall heat content and therefor the differences are
small. Fig.2 shows the situation for SOX/SH2: even
when subcooled to 2K (3,6R), the difference to
LOX/LH2 performance is only 0,5%!
3760
3770
3780
3790
3800
3810
3820
3830
Liqui d Soli d, 14 K Soli d, 2K
Frozen Flow
Equilibrium Flow
Fig.2: Mass specific impulse [m/sec] of O2/H2 propel-
lants at optimum mixture ratio, standard 68:1 equili-
brium expansion (7)
The following
Tab.1 contains further data about SOX/SH2 Isps
as function of mixture ratio. As Fig.3 shows, the opti-
mum mixture ratio is where one is used to find it when
Oxygen and Hydrogen are liquid.
O/F Ispvac,E Isp,E vg.density vol. spec. Isp, E
1,00 3374 3167 0,146 461
2,00 3857 3617 0,207 748
3,00 4034 3766 0,262 987
4,00 4093 3804 0,312 1186
5,00 4091 3787 0,357 1353
6,00 4053 3734 0,399 1488
7,00 3980 3649 0,436 1593
8,00 3858 3529 0,471 1663
9,00 3714 3400 0,504 1712
10,00 3366 3117 0,533 1750
Tab.1: Theoretical performance of SOX/SH2 propel-
lants at 68:1 equilibrium expansion (9)
1 2 3 4 5 6 7 8 9
3300
3400
3500
3600
3700
3800
3900
F
Isp[m/s]
O/F
E
Fig.3: Mass-specific impulse of SOx/SH2 as function
of O/F mass ratio at 68:1 equilibrium expansion (7)
In general, the density of solids is higher than the one
of liquids. This is also true for Oxygen and Hydrogen
(see Fig.5). As a consequence, the slight Isp reduction
found for CSPs is more than offset when volume spe-
cific Isps are compared (see Fig.6).
C S P s i n s t e a d o f s l u r r i e s
Frozen Hydrogen provides for easy mixing with denser
materials. There is no danger of separation as with
slurries. Propellants of the type O2/Al/H2 or O2/Mg/H2
can therefor be used as monopropellants (defined as
explained above), by using solid Oxygen elements
along with frozen Hydrogen / metal suspensions.
In the following Tab. 2 are shown results of calcu-lations with such propellants. The amount of metal in
the fuel was chosen between 45 to 90%. However, the
ratio of Oxygen to Aluminum was always kept in equi-
valence according to the formation of Al2O3 due to the
earlier demonstration in (3)(4)(5) that this is optimal. It
should be noted that the exhaust products near the Isp
maximum around 60 - 65% Al (corresponding to a
molar ratio of about 8, see Fig.4) are exclusively made
up of solid Al2O3 and molecular Hydrogen. There is no
H2O in the equilibrium at all, hence the term "Chemi-
cal heating of Hydrogen" (as opposed to nuclear heat-
ing, for instance).
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As can be seen in Fig.1 and Tab. 2, there is only a
small gain compared with the metal-free system. Now
how about density? Fig.5 shows average specific pro-
pellant masses over mixture ratio, Fig.6 the correspond-
ing volume specific Isps (i.e. Isp times specific mass).
%ALin fuel
%H2in fuel
Mass ratioH2/Al
Molar ratioH2/Al
O/F[kg SOX /
kg fuel]
Isp,E[m/s]
Ispvac ,E[m/s]
vol. spec.Isp E
Avg. prop.density
[g|cc]
90,00 10,00 0,11 1,50 0,80 3011 3308 2416 0,8028
81,80 18,20 0,22 3,00 0,73 3472 3801 1864 0,5365
75,00 25,00 0,33 4,50 0,67 3728 4077 1541 0,4136
69,23 30,77 0,44 6,00 0,62 3837 4160 1344 0,3425
64,29 35,71 0,56 7,50 0,57 3872 4172 1147 0,2962
60,00 40,00 0,67 9,00 0,53 3861 4145 1017 0,2636
56,25 43,75 0,78 10,50 0,50 3822 4095 915 0,2395
52,94 47,06 0,89 12,00 0,47 3774 4038 833 0,2209
50,00 50,00 1,00 13,50 0,45 3722 3978 767 0,2061
47,37 52,63 1,11 15,00 0,42 3670 3918 712 0,194145,00 55,00 1,22 16,50 0,40 3616 3859 666 0,1841
Tab. 2: Theoretical performance of SOX/(SH2+Al) CSPs with equiv. ratio of SOX/Al. (9)
Mol H2/Al
Isp,v
ac,E
[m/s]
2000
2500
3000
3500
4000
4500
0 2 4 6 8 10 12 14 16 1
I_sp_vac_E
I_sp_E
Fig.4: Isp (E, 68:1) of SOX/(SH2+Al) CSPs with
equiv. ratio SOX/Al; (9)
121086420
0,0
0,2
0,4
0,6
0,8
1,0
Avg.
Prop.
Density[g/cc]
Mass Ratio: Ox./total Fuel
SOX/SH2
LOX/LH2
45%Al
75% Al
90% Al
Fig.5: Average propellant density of O2/H2 propellants
in solid and liquid state (9)
121086420
0
1000
2000
3000
Isp,vol[kg/s,m
2]
Mass ratio O/(H2Al)
SOX/(SH2+Al)
SOX/SH2
LOX/LH2
Fig.6: Volume-specific impulse of SOX/(H2Al) as
function of O/F mass ratio at 68:1 equil. expansion; (9)
0,90,80,70,60,50,4
0
1000
2000
3000
Mass ratio O/(H2+Al)
Isp,vol[kg/s,m
2]
45
60
50
75
90 % Al
Fig.7: Volume-specific impulse of SOX/(SH2 +Al) as
function of O/F mass ratio at 68:1, equil. expansion;
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Oxygen in equivalence to Aluminum. %Al refer to fuel
composition; (9)
Fig.7 shows the aluminized part in more detail. Density
and hence Isp,vol skyrocket within a small O/F range,
because in comparison with Hydrogen, Aluminumneeds ten times less Oxygen for combustion.
The effect of more and more metal combustion is to
replace more and more Oxygen, rather than Hydrogen.
It takes 8kg Oxygen for burning 1kg of Hydrogen, but
less than 0,9 kg for 1kg of Aluminum. Compared on a
constant total impulse basis, it would therefor take
quite aluminum rich fuel mixtures in SOX/(SH2+Al)
propellants if the average density of SOX/SH2 were to
be equaled.
Thus the heating of Hydrogen by metal-combustion
rather than Hydrogen-combustion is a process with
much higher efficiency in terms of propellant massconsumption. As Tab. 3 shows, several metals release
more than twice as much energy per unit Oxygen as
Hydrogen does. Ironically, this is exactly the reason,
why metal-combustion used in thermal Hydrogen pro-
pulsion is not a very good means for improving average
propellant density.
Prod.
(solid)Hf
[Mj/kg]
A-wght
metal
M-
wght
prod.
MJ/kg-
metal
MJ/kg
O2
BeO 23,948 9,013 25,013 66,462 37,439
Li2O 19,971 6,94 29,88 42,992 37,296
B2O3 18,338 10,82 69,64 59,014 26,606
Al2O3 16,412 26,98 101,96 31,012 34,862
MgO 14,947 24,32 40,32 24,780 37,666
H2O g 13,431 1,008 18,016 120,029 15,124
Tab. 3: Energy released by the formation of oxides
S O X / f r o z e n H y d r o c a r b o n s
Obvious candidates for improving on the density Isp ofSOX/SH2 are hydrocarbon fuels. The following data
(Tab. 4) show SOX with frozen Kerosene as an exam-
ple.
LOX/RP1 SOX/SRP1
O/F Isp
68:1,E
avg. spec.
mass
vol. spec.
Isp E
Isp
68:1,E
avg. spec.
mass
vol. spec.
Isp E
1,00 2015 0,924 1863 1884 0,976 1838
2,00 2844 0,989 2812 2771 1,046 2899
2,50 2944 1,009 2970 2898 1,068 3094
3,00 2908 1,024 2979 2876 1,085 3121
4,00 2765 1,047 2895 2737 1,110 3038
5,00 2637 1,063 2803 2609 1,127 2941
6,00 2522 1,074 2709 2492 1,140 2840
7,00 2414 1,083 2615 2382 1,150 2738
8,00 2314 1,090 2522 2279 1,157 2637
9,00 2221 1,096 2433 2185 1,163 2542
10,00 2136 1,100 2350 2099 1,168 2453
Tab. 4: Isps and average densities of Oxygen/Kerosene in liquid and solid state;
based in part on extrapolated values by (9)
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2000
2200
2400
2600
2800
3000
3200
0,0 2,0 4,0 6,0 8,0 10,0
MR [-]
I_sp_vac_E
I_sp_E
Fig.8: Standard Isp 68:1, E and Isp,vac.
of SOX/SRP1 (9)
Boundary layer combustion insolid propulsion
The concept of dissected solid-rocket propellant
grains is based on the idea that the degree of separation
is a perfect means for controlling the rate of propellant
consumption in solid rockets. Example: consider a
cylindrical grain of SOX followed by a cylindrical
grain of hydrocarbon, both with a central combustion
channel (not necessarily of circular shape). Without
other means, this 2-piece "Internal Hybrid" arrange-
ment would not be able to sustain combustion. As the
degree of dissection is increased by increasing the
number of alternating oxidizer and fuel elements, the
regression velocity attainable after ignition would, to a
certain degree, approach the value expected for a mix-
ture of the two reactants. Dissected grains offer a new
degree of freedom for controlling combustion rate, that
may not have been required with conventional propel-
lants ("Sandwich grains") but might be essential for
cryogenic solids.
Fig.9 shows a motor with modular CSP grain de-
sign. (All dimensions in this figure are arbitrary). The
grain is a stack of alternating oxidizer- and fuel ele-
ments. The cross section of the central combustionchannel can have any shape and change along the grain.
Gasgenerator ModuleIgniter- and sustainermodule
Fuel Element
Oxidizer Element
Fig.9: Solid rocket motor using a modular CSP grain in
disk-stack design
The concept of modular dissected solid-rocket pro-
pellant grains is based on the idea that for any solid
propellant combination there exist subcritical mixture
ratios that do not require separation for well behaved
combustion. Propellant elements with such mixture
ratios
offer a further degree of freedom for obtaining adesired value of regression rate (combustion veloci-
ty)
offer themselves as modules for ignition as well asfor sustaining and enhancing combustion (not nec-
essarily using the same propellants as the bulk of
the modular grain composite).
The element at the front end of the stack in Fig.9 is
formed by such a fuel rich gas generator composition,
capable of self-sustained burning. It serves as igniterand sustainer (of course, the arrangement of oxidizing
and reducing elements might just as well be the other
way round). If required, there is another self sustained
burner at the end, serving as what with hybrids is called
"turbulator". In between, there is the dissected stack
with multiple boundary layer combustion.
There are many more geometrical solutions other
than sandwiches, e.g. wedge shaped arrangements or
cigarette burners. A "mixedness parameter" has been
defined for their characterization, see (6).
To ensure proper operation, they are all subject to
the following boundary conditions and requirementsconcerning their operational parameters:
1. The time average over-all mixture ratio (= "tankmixture ratio") determines the overall equivalence
ratio. It is defined as
O/ F to t 1
tc (O/ F)xmaxt0
c
t dt M oxMfu
where
tc : cut-off time (i.e.: c = combustion duration)
(O/F)xmax
: instantaneous over-all mixture ratio
Mox : Total amount of oxidizer present in ele-
ments and gas generators
Mfu : Total amount of fuel present in elements
and gas generators
xmax: rear edge of grain, maximum value of the co-
ordinate of linear length x
Desired condition: O/ F tot = some specific value2. Instantaneous over-all mixture ratio, i.e. O/F ratio
of gases passing rear edge of grain (of course relat-
ing to the composition of their sources before com-
bustion)
(O/F)xmax(t) = mfox(t) / mffu(t)
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where
mf: mass flows with
mfox(t) = ii1
n ox
A i(t)r i ii1
ngg
mfox,gg (t) and
mffu(t) = ditto
where
nox , nfu , ngg: number of respective oxidizer-, fuel-
or gas generator elements
i: number of individual elements (e.g. number 1 =
gg, 2, 4, 6 = oxidizer, 3, 5, 7 = fuel)
i: specific mass of element i
Ai(t): combustion surface area of element i at time t
r i : average regression velocity (= combustion rate
per unit area) between linear co-ordinate xi,o (front
edge of element i) and xi,e (rear edge)
mfox,gg(t), mffu,gg(t): mass flow contribution of gas
generator i at time t according to its primary com-
position
Desired condition: the instantaneous (O/F) over-all
mixture ratio should at all times be as close to the
desired value as possible:
1
t c (O / F)xmax
t0
t c
(t)dt = (O/F)tot
3. The average regression velocity of element i ex-tending between xi,o (front edge) and xi,e (rear
edge)
r i =1
x i ,e x i,0ri (x)dx
x i,0
x i,e
where
ri(x): local regression velocity of element i, that
must satisfy the basic hybrid heat balanceri(x) = q (x) / (ihi)
where
q (x): local heat flow per unit area
hi : specific heat of sublimation of element i
Desired condition:
ri(xi,e) = ri+1(xi,o)
i.e. adjacent elements should keep a smooth commonedge. This desire should actually by quite easy to
meet, because
steps do not normally form in hybrids since they en-
hance regression of protruding material
regression velocities can be matched by either ma-
nipulating the lengths of individual elements or
their total number.
Performance of cryogenic solidrockets
A general feasibility study of Cryogenic Solid
Boosters (CSBs) sponsored by the German Aerospace
Research Center DLR (2) was finished in 1996. Thebasic assumption was that the above concepts are feasi-
ble and it aimed at identifying specific problem areas
and their possible solutions. Boosters were investigated
with solid Hydrogen (SH2) and Oxygen (SOX) as a
sample propellant combination. As a result, no insur-
mountable problems were found in the areas of cooling
equipment and its operation during fabrication and
launch operations; neither were there problems with
thrust to weight ratio of uncooled but insulated CSBs
that leave their terrestrial cooling equipment at the
launch pad. The pressure variation of the melting points
of SOX (and its allotropic modifications) and of SH2
appeared manageable.
Precautions will be necessary with respect to me-
chanical stability and the influence of off-design condi-
tions. While many frozen liquids are quite sturdy (see
(11)) special supporting and enclosing measures are
suggested (see (12)).
On the other hand, under specific assumptions, very
substantial performance gains were calculated for
ARIANE V and the US-STS if the conventional solid
boosters were replaced by cryogenic ones (SOX/SH2).
As is shown in
Tab.5, the lift-off mass of the boosters of both space
transportation systems could be cut down to about 2/3
of their present mass.
As was already pointed out in the section on Isp-
calculations, the use of metalized SOX/(Al,SH2) solid
propellants does not really improve on the metal-free
combination. A just slightly lesser mass would have to
be bought by using about 60% of Aluminum in the fuel
(see Tab.6). Whether such amounts could be burned
with sufficient efficiency, must be doubted.
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STS-SRB STS-SRB
CSP-
Ersatz
ARIANE
5-EAP
ARIANE
5-EAP
CSP-
Ersatz
Isp [m/s] 2634,9 3438,52 2678,13 3438,52
pc [bar] 62,4 20 ~30 20
pc/pe - 66 ~77 66
O/F 2,3 5 ~2,3 5
M-prop [Mg] 504 349,525 237 167,870
Itot [MNs] 1320 1320 634,7 634,7
cut-off time [s] 124 120 130 130
Thrust [kN] 11500 5300
M-tank [Mg] 91,000 54,411 38,000 26,207
Lift-off [Mg] 595 403,936 275 194,077
Tab.5: STS and ARIANE V boosters compared with
their CSP-Ersatz; after (7)
The results were optimized for best mixture ratio ofSOX/SH and for optimum chamber pressures chosen,
see Fig.10
400
410
420
430
440
450
460
470
480
490
500
0 20 40 60 80 100 120
Chamber pressure [bar]
Fig.10: Mass of SOX/SH2 SRB-Ersatz Boosters at opt.
O/F over chamber pressure (7)
Aluminum addition does not replace Hydrogen, ra-
ther, it replaces Oxygen. This can be seen in the follow-
ing example series of results, refering to ARIANE V
EAPs: the non-metalized EAP-booster uses 32340 kg
of SH2 and 132590kg of SOX. The weight minimized
metalized booster of same total impulse burns 67150kg
Aluminum with 60470kg of Oxygen (less than half the
previous amount!), heating but not oxidizing 37300kg
of Hydrogen (15% more!).
%Al Mp [t] Vp[m3]
90,00 210,9 262,8
81,80 182,9 340,7
75,00 170,3 412,0
69,23 165,5 472,5
64,29 164,0 553,8
60,00 164,5 624,2
56,25 166,2 694,1
52,94 168,3 762,0
50,00 170,6 828,0
47,37 173,0 891,8
45,00 175,6 954,1
0,00 164,9 521,0Tab.6: Propellant mass and volume of boosters of
ARIANE V size (Itot ~635 MNs) with SOX/(Al,SH2)
cryo-solid propellants
Feasibility demonstrationOngoing work aims at the experimental demonstra-
tion of the feasibility of cryogenic solid grain combus-
tion and is considered as the second step in a four step
procedure leading to the acquisition of the technology
of modular, dissected solid propellant grains.
SummaryAs a conclusion, the concept of Modular Dis-
sected Cryogenic Solid Rocket Propellants opens a
whole area of new chemical propulsion research and
development. Chemical propellant combinations of the
highest known values of Isp can be used in solid rocket
motors because deflagration is replaced by boundary
layer combustion in an arangement that was dubbed
"multilayer internal hybrid combustion".
The concept was accepted as "potentially revolu-tionizing propulsion technology" at the recent Ad-
vanced Propulsion Workshop of the IAA (Jan.1998 at
El Segundo, Cal., see 3). In the framework of the APW
a web site was established for describing and discuss-
ing the concept in public (http://www.aero.org/apw/).
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