1998-10 r.lo m.d. cryogenic solid propellant grains

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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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    9

    Literature

    (1) R.E.Lo, , DFVLR-Stuttgart: "Technical Feasibili-

    ty of Chemical Propulsion Systems with very highPerformance", Proceedings of the XVIIIth Astro-

    nautical Congress, Belgrade, 25.-29.9.1967, pp.

    121-132

    (2) R.E.Lo, DFVLR-Lampoldshausen: "Quasihybrid

    Rocket Propulsion Systems (Quasihybride Rake-

    tenantriebe)", Raumfahrtforschung, Heft 4, April

    1970, in German.

    (3) R.E.Lo, DFVLR-Stuttgart: "Chemical Heating of

    Hydrogen by Tribrid Combustion (Chemische

    Wasserstoffaufheizung durch tribride Verbren-

    nung)", DGLR-Symposion 'Chemical Rocket En-

    gines', 21.March,1967 Munich, Chemie-

    Ingenieur-Technik (1967) 39, Heft 15, S. 923-

    927, in German

    (4) R.E.Lo, DFVLR-Lampoldshausen: "Chemical

    Heating of Hydrogen by Aluminum Combustion

    with Oxygen or FLOX (Chemische Wasserstof-

    faufheizung durch Verbrennung von Aluminium

    mit Sauerstoff oder FLOX)", DLR-Mitt. 70-03

    (Feb. 1970), in German

    (5) R.E.Lo, DFVLR-Lampoldshausen: "Theoretical

    Performance of the Rocket Propellant Combina-

    tion F2,02/LiH,Al/H2 and Simpler Subsystems

    (Theoretische Leistungen des Raketentreibstoff-

    systems F2,02/LiH,Al/H2 und einfacherer Teil-systeme)", DLR-Mitt. 69-21 (Dez.1969), in Ger-

    man

    (6) R.E.Lo, Berlin Univ.Techn.: "A Novel Kind of

    Solid Rocket Propellant", Aerospace Science and

    Technology, Elsevier, 1998

    (7) B.Voslamber, M.Voslamber.R.Lo, Berlin Un-

    iv.Techn: "Feasibility Study on Cryogenic Solid

    Rocket Boosters", DARA Project 50TT 9631, Fi-

    nal Report 2/1997 (in German)

    (8) D.Froning, Jr.,Roger E. Lo: Possible Revolutions

    in Rocket Propulsion", 49th Int. Astronautical

    Congress, Melbourne, Austr., Oct.02.1998, Ses-

    sion IAA.3.3., Co-operation and Competition inAdvanced Propulsion, IAA-98-IAA.3.303

    (9) Calc. by H.Adirim, ILR/ Berlin Univ.Techn,

    Jul.1998

    (10) P.G.Carrick, Phillips Lab, Edwards AFB: "Theo-

    retical Performance of High Energy Density

    Cryogenic Solid Rocket Propellants", AIAA-95-

    2892, 31st AIAA Joint Propulsion Conf., San Di-

    ego Jul.1995

    (11) W.F.Staylor, NASA-Langley R.C.: "Frozen Pro-

    pellant - A Booster Concept?", Astronau-

    tics&Aeronautics, Sept.1997, p.56-59

    (12) R.E.Lo, N.Eisenreich: "Modulare und kryogene

    Feststofftreibstze - eine neue Klasse chemischer

    Raketenantriebe" Deutscher Luft- und Raum-

    fahrtkongress, Bremen 5.-8.Okt.1998