\
THE COLLEGE OF AERONAUTICS
CRANFIELD
PROPERTY REQUIREMENTS FOR LIQUID
ROCKET PROPELLANTS
by
NOTE NO. 9 7 . November. 1959. T H E C O L L E G E O P A E R O N A U T I C S C R A N F I E L D P r o p e r t y r e q u i r e m e n t s f o r l i q u i d r o c k e t p r o p e l l a n t s b y -E. M. Goodger, M.Sc.(Sng.), Ph.D., A.M.I.Mech.-E., A.F.R.Ae.S., P.Inst.Pet.
An analysis is made of the properties necessary for liquid rocket propellants to give effective performance with acceptable handling. Typical propellants are examined against this requirement background, and their relative suitability assessed,
1, Introduction 1 2, Performance Characteristics 1 2.1. Reaction Energy 1 2.2. Specific Impvilse 4 2.3. Combustion Characteristics 6 3, Handling Characteristics 7 3.1. Volume Impulse 7 3.2. Vapour Pressure 7 3.3. Thermal Stability 7 3.4. Viscosity-Temperature Relationship 7 3.5. Heat Capacity 8 3.6. Combustion Characteristics 8
3.7. Uniformity, Availability, Cost 8 3.8. Storage, Transfer, Transport 8
4, Overall Suitability 9 4.1. Hydrocarbons 9 2f. 2. Liquid Hydrogen 9 4.3, Sliorries 10
4.4, Hydrides 11
4.5, Oxidants 13 5, Conclusions 14 6, Acknowledgements 15 7, References 16 Appendix 17 Tables 1, 2, 3 Figures1, Introd\Action
In the literatijre, there is a tendency to deal individually with rocket propellants, and to list their properties in the manner of a catalogue. This presentation is necessary eventually, vrhen final
details have to be settled after a choice of propellants has been made, but selection may be simplified by the comparison of individioal properties
over a range of propellant types. Furthermore, it is convenient to subdivide these properties into those which determine the performance of the rocket engine, and those which influence handling. In this Note, the following definitions are implied :
-Performance Characteristics - Factors with a major influence upon the project design and propulsive effectiveness of the rocket engine. Handling Characteristics - Factors v/ith a major influence upon
the mechanical detail design of the rocket engine, the fuel system, the vehicle, and ground-handling equip-ment, and on the economics of s\jpply, In general terms, performance is the attraction in the search for improved propellants, whereas handling imposes the problems. In many cases, it is convenient to take liquid oxygen and a conventional hydrocarbon fuel of the kerosine type for reference purposes, so that the extent of improvements may be assessed quantitatively.
This Note has been prepared with the following objectives : -i To ccmpare the performance character-ist-ics of typ-ical l-iqu-id
rocket propellants,
il To compare the handling characteristics of these propellants, iii To assess the overall suitability by examining the propellants
ind ividually.
2, Performance Characteristics 2,1, Reaction Energy
Since most propulsive systems are based upon a chemical reaction, a high reaction energy per unit mass of reactant is the first essential. Generally, the process is one of oxidation, although alternative reactions
such as fluorination sire both possible and desirable,
The gravimetric heat quantities released by the oxidation of a number of elements are charted in Pig. 1, and hydrogen is seen to have
far and away the highest heating value, at 28,700 CHJ/lb. ether promising
Introduction 1
Performance Characteristics 1
2.1. Reaction Energy 1
2.2. Specific Impulse 4
2.3. Combustion Characteristics 6
Handling Characteristics 7
3.1. Volume Impulse 7
3.2. Vapour Pressure 7
3.3. Thermal Stability 7
3.4. Viscosity-Temperature Relationship 7
3.5.
Heat Capacity 8
3.6. Ccmbustion Characteristics 8
3.7. Uniformity, Availability, Cost 8
3.8. Storage, Transfer, Transport 8
Overall Suitability 9
4.1. Hydrocarbons 9
4.2. Liquid Hydrogen 9
4.3. Slurries 10
4.4. Hydrides 11
4.5. Oxidants 13
Conclusions 14
Acknowledgements 15
References 1é
Appendix 17
Tables 1, 2, 3
Figures
1. Introduction
In the literature, there is a tendency to deal individxially with rocket propellants, and to list their properties in the manner of a catalogue. This presentation is necessary eventually, w^hen final
details have to be settled after a choice of propellants has been made, but selection may be simplified by the comparison of individual properties
over a range of propellant types. F\jrthermore, it is convenient to subdivide these properties into those which detemdne the performance of the rocket engine, and those which influence handling. In this Note, the following definitions are implied :
-Performance Characteristics - Factors with a major influence upon the project design and propulsive effectiveness of the rocket engine. Handling Characteristics - Factors with a major influence upon
the mechanical detail design of the rocket engine, the fuel system, the vehicle, and ground-handling equip-ment, and on the economics of supply, In general terms, performance is the attraction in the search for improved propellants, whereas handling imposes the problems. In many cases, it is convenient to take liquid oxygen and a conventional hydrocarbon fuel of the kerosine type for reference pixrposes, so that the extent of improvements may be assessed quantitatively.
This Note has been prepared \yith the following objectives :-i To compare the performance character:-ist:-ics of typ:-ical l:-iqu:-id
rocket propellants,
il To compare the handling characteristics of these propellants, iii To assess the overall suitability by examining the propellants
individually.
2, Performance Characteristics 2,1, Reaction Energy
Since most propulsive systems are based upon a chemical reaction, a high reaction energy per lonit mass of reactant is the first essential. Generally, the process is one of oxidation, although alternative reactions
such as fluorination are both possible and desirable,
The gravimetric heat quantities released by the oxidation of a number of elements are charted in Fig, 1, and hydrogen is seen to have far and away the highest heating value, at 28,700 CHU/lb. Other promising
elements, in descending order of oxidation energy, are beryllium, boron, lithium, carbon, aluminium, silicon, and magnesium. Since difficulties may be encountered with the use of hydrogen alone, these alternative
elements must also be considered, for use either as additives or as hydrogen carriers.
In chemistry, the reaction energy is defined as the heat added to the products in order to regain the standard state sifter a reaction with the reactants originally at the standard state. An exothermic reaction, therefore, is indicated by a negative sign. -The reaction energy requirement may thus be written as max. neg. ZH , where the superscript represents the standard state of 25 C. and 1 atmosphere. Reaction energy may be expressed in terms of the heats of formation of
the reactants and of the resulting products
:-A H° = S :-A H° Products - S:-A H° Reactants.
It follows that max, neg. A H calls for products of max. neg. AH„ and reactants of max. pos, AH-,
The heat of formation is derived from the following processes :-Elemental Molecules 3. Gaseotis Atoms ^Conipound Molecules
(at standard state)^-^. .,J '^—-v '' (^* standard state) Atomisation Heat Bond Energy
Endothermic Exothermic Positive Negative
Thus, A H° = 2 A H° - 2 D ( X - Y ) ,
where A H represents atomisation (or sublimation) heat required to convert the elemental molecules into gaseous atoms, and D ( X - Y ) represents the bond energy required to dissociate an X-Y bond into gaseous atoms.
As shown in Fig. 2, the requirement of max. neg. A H now develops into the four requirements of min. and max, pos. 2 A H for the products and reactants respectively, and max. and min. pos. 2 D ( X - Y )
for products and reactants respectively, \Tith individual hydrocarbons, values of these quantities included in Fig. 2 and plotted in Fig. 3 begin to shoViT the fundamental requirements of molecular structure for high reaction energy. These may be summarised as low molecular vreight with numerous multi carbon-carbon bonds, yet a high hydrogen content producing large concentrations of H^O.
With hydrocarbon fuels produced commercially from petroleum feedstock, as shoivn in Pig^ 4, gravimetric lieatinp; value is found to fall by about 1C^ for a 35/^ rise in specific gravity from 0,7 to 0,95. The combined effects result in a rise in the voliAmetric heating value of about 25%. An increased content of hydrogen, therefore, reduces the specific gravity and raises the gravimetric heating valtie until the ultimate is reached in hydrogen itself. The specific gravity of liquid hydrogen at its boiling peine, of -253''C. is 0.071. In airborne
considerations, mass ic: a vJtal factor, so that the gravimetric energy requirement directs fuel selection towards la\\' specific gravity fuels, and liquid hycirogen. Y.'ith high-speed vehicles operating v/ithin a gaseous atmosphere, however, aerodynamic drag imposes a minimum volume requirement. In these applications, therefore, a heavy fuel may be advantageous, and specific gravity can be increased appreciably by loading liquid fuel with finely-pulverised solid to form a slurry. This arrangement
offers a practicable compromise between the attractive high density and the difficult pumpability of the solid alone.
The importance of replacing carbon as a hycirogen carrier is emphasised by the realisation that, in aviation kerosine, carbon contributes about 86% of the mass of the fuel, yet only 62fu of the energy. Hydrides are formed by boron, lithium, silicon, and magnesium in addition to carbon. Hydrides of nitrogen are also worth attention since the nitrogen can be utilised as a temperature'-controlling diluent. Fuels of high-energy interest, therefore, resolve;themselves into the
follovidng groups :
-i The l-iydrocarbons ii Liquid hydrogen
iii Slurries of beryllium, boron, lithium, carbon, alijiminium, silicon, and magnesium in a suitable high-energy carrier fluid.
iv Hydrides of beryllium, boron, lithium, silicon, magnesium and nitrogen.
The heating values plotted against specific gravity in Fig, 5
indicate dramatically the improvements available v/ith the non-hydrocarbon high-energy fuels. Arbitrary curves have been constructed to connect the gravi metric, and volumetric, heating values of hydrogen, boron, and their compounds. The trends v/ith specific gravity are similar to those noted v/ith petroleum fuels. A 60/4O mass ratio boron-kerosine slurry is included, together with ammonia, hydrazine, and a typical alcohol.
Since both gravimetric and vo].umetric heating values are significant for high-speed applications v/ithin a gseous atmosphere, the product
of these two heating values can be cairpared v/ith that for aviation kerosine, and expressed as the 'Performance Index', i.e.
where K for aviation kerosine ( A V T U R ) is 840 x 10 approximately in consistent units. Fig, 6 shows values of perfonnemce index plotted against specific gravity. Arbitrary curves have been constructed to connect the hydrogen-carbon and hydrogen-boron compounds, A more representative fuel for reference purposes might be RP-1, which is a pure, light-cut kerosine of low aranatic content, prepared especially for rocket applications. However, its properties are not far removed from Avtur, and its performance index is approximately 1,02,
2,2, Specific Impulse
Combusticn in the absence of diluent gases results in temperatures of 3000 C. or more, compared with about 2000 C. met in most
iiir-breathing engines. The thermodynamic behaviour of the product gases in their passage through the propelling nozzle (Pig. 7) now beoccoes
significant, and the performance of the propellants cannot be assessed upon the reaction energy alone. An overall assessment becomes possible by appljring the Bernouilli esqaression to the gas flow through the nozzle, and calculating the exit velocity, V , This leads to an expression for the specific impulse, I , which is fhe thrust produced with unit mass consumption rate of propellants (see Appendix).
T 1 2 y G T„ I' / „ / „ N y-1 1
^sp = 32rr7r ^)•y^ i r ° | J - ( p / P c ^ — J
v/here p = nozzle exit pressure (= atmospheric pressure under optimum conditions),
p = canbustion chamber pressure.
y = ratio of specific heats for propelling gases T = temperature of combustion chamber gases
c
M = mean molecular weight of propelling gases
The requirements for h i ^ specific impulse may now be written as follov/s
:-i M:-in. p / p For a g:-iven p , th:-is represents a max, p , wh:-ich (see Pig,8) is a function mainly of nozzle throat area and of
propellant consumption rate.
ii Min. V The value of y is minimised by increasing the (see Pig.9) concentration of monatomic molecules in the
Gt-nerally, the level of combustion temperature
follov/s that of the heating value, as shov/n in Fig. 1, Some discrepancies occur, hov/ever, (e.g. Ti)due to the relative values of specific heat of the resulting prod-ucts. The hydrocarbons show a striking example of this (Pig. 10) where an increased heating value is seen to result in a reduced T due to the influence of the greater quantity of water vapour produced. Hence Ï is a propellant function.
This requires large concentrations of light products (e.g. v/ater vscpoxxc), and is therefore a propellant function.
Mixture ratio is an additional variable which produces a change in a derived parameter from a combination of changes in individual parameters. Figs. 11A and B show mixture ratio curves for tv/o typical propellant pairs. Generally, although the maximum non-dissociated temperature occurs at the stoichiometric ratio, the influence of dissociation, together v/ith the relative values of product heat capacities, is such that the maximum dissociated temperature lies on the fuel-rich side of stoichiometric, 'Jith fuels of high hydrogen content, therefore, producing large concentrations of v/ater vapour, the reducing value of M v/ith fuel enricliment brings the specific impulse peak further over on the fuel-rich side. V.'ith these fuels, also, the
lov/ specific gravity compared with most oxidants tends to move the peak of the volume impulse curve back towards stoichiometric.
Prom the data presently available, a general direct relationship is found to exist betv/een peak specific impulse and CHU/lb propellant mixture. This relationship is reflected in the fall in peak specific impulse v/ith increased specific gravity, shov/n in Fig. 12, Because of the inclusion of specific gravity in the impulse term, a plot of voli;ime impulse shows a closer direct relationship with specific gravity, as in Fig. 13A. The large-scale replot in Fig. 13B indicates a difference of about ^ betv/een shifting and frozen equilibrium values of peak
volume impulse.
Since it is difficult to predict the specific iinpulse from a single characteristic property of a fuel, comparisons have been made in the form of charts (Figs. 14 and 15) vri.th the fuels arranged in descending order of peak specific impulse and peak volume impulse respectively. Arbitrary curves have been drawn to connect these values for four different oxidants, and for fluorine. The charts therefore also indicate the effectiveness of oxidants, and show that maximum specific irnpu].se is produced with iiydrogen and fluorine as
a propellant pair, and maximum volume impulse with hydrazine and fluorine. iii Max. T c (see Pig.9) iv Min. M (see Fig.9)
2,3. Combustion Characteristics
Since liquid propellants are invariably sprayed into the combustion chamber, properties which provide for fine atomisation and rapid
vaporisation are necessary for a compact flame and a short chamber length. These properties may be listed as follows
:-i. Spray Formation
Specific Gravity - Slight, direct, effect on droplet penetration, Viscosity - Direct effect on both droplet diameter and
penetration.
SiiTface Tension - Direct effect on droplet diameter and penetration.
ii Spray Vaporisation
Volatility - Direct effect on vaporisation rate Thermal Conductivity - Direct effect on vaporisation rate, Specific Heat - Inverse effect on vaporisation rate, The required fuel properties, therefore, are a low viscosity, surface tension, and specific heat, and a high volatility and thermal conductivity,
Once flame has been initiated, and the supply of propellant vapour assured, the length of the combustion chamber can be minimised by means of a high flame speed. Flame speeds of the hydrocarbons lie at a very low level, ranging from about 2 to 4 ft/sec. when b\iming in air at atmospheric pressure, about ten times this v/hen burned in oxygen, and increased further when burning under high pressure. The flame speeds of hydrogen, and the boron hydrides, are considerably greater than those of the hydrocarbons.
The mixture range of inflammability is noraially associated with flame speed, A vd.de inflammable range is a valuable property where changes in propellant flow rate are lilcely to be encountered since flame stability is improved, and flame-out avoided. The inflammable range of hydrogen is very much v/ider than that of the hydrocarbons,
For extended operation, clean combustion is essential, with the absence of nozzle fouling by deposit build-up. 17ith most fuels, this is achieved by control of mixture ratio to avoid over-enrichment of fuel, \7hen condensation and freezing points of combustion products are high, the danger exists of liquid and solid particles forming in the propelling gas stream, and depositing on the internal surfares of the
with boron fuel, and v/ith slurries. Aluminium, with an oxide freezing teiTiperature of about 2000°C. , is particularly difficiilt in this respect, and Fig. l6 shov/s the extent of deposit accumulated in a 2 in. bore duct during a 30 second test burning a 30^^ aluminiiffli-kerosine sltirry in air at a low fuel/air ratio (Ref,7).
3, Handling Characteristics 3.1. Volume Impulse
Probably the foremost requirement for rocket operation within a gaseous atmosphere is a maximum volume impxilse in order to minim-ise fuel system bulk and aerodynamic drag. Since this property is so closely dependent upon specific impu].se, its discussion has been included in the Performance section. The associated controlling property, specrlfic gravity of the propellant mixtijre, is seen to have a general direct relationship v/ith volume impiolse (Pig.. 13) so that
a high specific gravity is generally desirable for atmospheric operation. 3.2. Vapour Pressure
High-speed operation within a gaseous atmosphere leads to high biilk fuel temperatures due to kinetic heating. Thermal insulation
or refrigeration may be necessary. Vapour pressure rises with temperature at an increasiig rate (Fig.l7) and must be either resisted by the
strength of the tank structure, or relieved by tank venting. The farmer incurs v/eiglit penalties, whereas the latter gives rise to vapour loss, particularly at high altitude. A low vapour pressure is required, therefore, and Ref. 8 suggests a maximum limit of 20 p,s.i,a. at 230 C. 3.3. Thermal Stability
Prolonged exposure of hydrocarbon fuels to kinetic heating leads to the formation of a sediment of insoluble oxidation products v^hich tend to filter blockage. Maximum thermal stability is essential for atmospheric operation due to kinetic heating, and for spatial operation due to radiation heating. The thermal stability of hydrocarbon fuels is currently assessed in a heated filter rig, and a thermally-stable kerosine (JP 6) is specified.
3.4. Viscosity-Temperature Relationship
A minimum viscosity-temperature dependence is desirable in order to retain pumpability at low temperatures (e.g. starting conditions in arctic regions, or low-speed operation at high altitxode) yet provide at high temperatiure a viscosity level sufficient to lubricate metering valves in the control system. The fall in viscosity with temperature rise for some rocket propellants is shown in Fig. 18,
3.5. Heat Capacity
A high specific heat, together with a high latent heat of vaporisation, permit smaller heat transfer surfaces and lower flow rates of coolant,
Sane values of specific and latent heats are included in Table 1. 3.6. Combustion Characteristics
Fuels bvirning vsrith high luminosity, caused by glowing particles in the flame, are liable to form smoke, and to promote metal fatigue due to radiant heat transfer to the combustor walls. The aromatics fall
into this category. One valuable property which helps to ensure ignition, flame stability, and relight at altitude, is the spontaneous ignitibility of the propellants on contact in the liquid phase. This phenomenon is termed 'hypergolic' ignition, and is exhibited by hydrazine and HTP, aniline and nitric acid, and some other propellant pairs. It simplifies the detail design of the rocket engine by the elimination of electrical^ ignition equipment, but inevitably, increases the risk of fire in handling. A vaJ-uable development is the system of 'thermal' ignition, in v/hich
HTP is decctirposed catalytically in the combustion chamber, the
de-composition temperature being sufficient to ingite kerosine spontaneously. This preserves the convenience of spontaneous ignition within the chamber, without the increased fire danger outside it.
3.7. Uniformity. Availability. Cost
Greater constancy of purity and quality permits smaller tolerances in design and hence saves weight. Specification limits ensure uniformity of product quality, but limits can be set only as uniformity of supplies, control of processes, and reproducibility of test techniques permit. Cost of the finished product depends upon availability of supplies, and complexity of processing. Beryllium, for example, is in such small supply that its consumption as a rocket propellant is impracticable. Boron, on the other hand, exists in nature mainly in the oxide form as a borate. Since this is also the resxiltant combustion product, reaction energy must be supplied during processing in order to produce elemental boron, and this adds to the cost.
3.8. Storage. Transfer. Transport
For ease of handling, propellants should exist in •'"he liquid
phase at normal temperatures. Gases require either pressurised containers which are bulky and heavy, or cryogenic equipnent which tends to be
expensive and fragile. Solids provide maximum density, but are difficult to pump. Propellants should be inert and stable, with no physiological hazards or corrosive tendencies. Stability at extremes of temperature is required, v/ith maximum fire safety. These requirements dictate a boiling point well above ambient, and a freezing point well below, compatibility vrith a wide range of structural materials, a high flash point, a high spontaneous-ignition tanperature, and a high cracking
(decomposition) temperature, Seme of these properties are plotted for petroleum fuels in Fig. 19. Acetylene, with a high heating value/lb
is rejected due to its detonability when pressurised, ^» Overall S;uitability
4,1 . Hy.'JJlQQ^gibQns
Petroleum fuels offer the immediate attractions of high
availability and relatively lov/ cost, coupled v/ith reasonable perform-ance and an extensive backlog of handling experience (Ref.8). Aviation kerosine and RP-1 represent the optimum types of rocket fuel v/ithin the commercial product ra^nge, but certain individual hydrocarbons, although of reduced availa.bility, offer overall iriprovements. Conclusions drav/n from such an overall assessment (Ref ,9) suggest that the polycyclic naphthenio type hydrocarbons provide the best
compromise of high hydrogen content, high densily, low vapour pressure, good thermal stability, and good lov/ teniperature characteristics. The materials examined include decalin, phenantlireno, fluoranthene, and isopropyl bicyclohexyl. The supply situation would be eased by specifying petroleum fuels rich in polycyclic naphthenes rather thrji the pure hydrocarbons alone,
4.2, Liquid Hydrogen
Plydrogen is seen to exhibit the ultim:ite in gravimetric heat content and, since it realises a fairly high conibustion temperature v/ith products of lov molecular weight, it also produces a maximum
specific iiiipiilse. With its attractive combustion chrracteristics of high flame speed, v«ade mixture range of inflai,xic;xbility, and clean buiTiing, hydrogen represents the optimum in performance,
On tlie handling side, the vary low density of hj'-drogen renders the liquid phase essential for aircraft use, and the lov/ density of liquid liydrogen coropletely eliminates the suioeriorily in impulse when expressed on a voliome basis. The low boiling point of liquid
hydrogen incurs problciris of vapour pressiure and of material suitability. The vapour loss when stored in a double-v/alled vacuum container is of the order of l^o per day. Shielding v/ith a layer of liquid nitrogen (b.p, -196 G,) can reduce this to 0,01 T^'o per day. The latter type of vessel is cor;iparatively fragile, and v/ill not v/ithstand pressiJre
differences in excess of about 20 p,s,i. For storage in aircraft, attention is tijrning more towards the 'polar vacuum' teclmique, v/hero the vacuum chamber is filled with a powdered mixtixre of altuninium vrith
silica, 'Perlite', or other solid. This reduces the radiation heat loss, which is proportional to the fourth power of the temperature difference, but increases slightly the conduction heat loss. ''JThen paired with liquid oxygen, the oxidant miglat be used to replace liquid nitrogen for shieldir.g purposes to give a compact, but higlily inflammable, installation,
Most f e r r o u s a l l o y ' s and many o t h e r m e t a l s undergo r e c r y s t a l -l i s a t i o n and become e m b r i t t -l e d betv/een ambient t e m p e r a t u r e and
- 2 5 3 C , , a l t h o u g h t h e i r t e n s i l e s t r e n g t h may i n c r e a s e . M e t a l s v/hich a r e c o n s i d e r e d s u i t a b l e f o r use v/ith l i q u i d hydrogen a r e c o p p e r , 1 8 - 8 s t a i n l e s s s t e e l , i n c o n e l , ;3nd monel, Nat^ural r u b b e r i s u s e l e s s a t t h e s e low t e m p e r a t u r e s , b u t p . t . f . e , r e t a i n s r e s i s t a n c e t o s h o c k , An i n t e r e s t i n g s i d e e f f e c t i s t h e s o l i d i f i c a t i o n of oxj^gen i f p e r m i t t e d
t o approach t h e s u r f a c e of l i q u i d h y d r o g e n . The oxygen c i y s t a l s s i n k i n t h e l i q u i d hydrogen t o form a h i g h l y iiiflanï.iable s l u n y . On puinping, i t i s found t h a t t h e c r y s t a l s c o l l e c t a c h a r g e of s t a t i c e l e c t r i c i t y , and t h a t tiie s p a r k s produced d u r i n g f r a c t u r e of t h e c i y s t a l s a r e c a p a b l e of i n i t i a t i n g a v i o l e n t e x p l o s i o n .
The low v i s c o s i t y of l i q u i d hydrogen p o m d t s h i g h pvunping
e f f i c i e n c i e s . B a l l b e a r i n g s have b e e n o p e r a t e d s u c c e s s f u l l y a t h i g h s p e e d s , u s i n g l i q u i d hydrogen a s t h e l u b r i c a n t . The h i g h s p e c i f i c h e a t i s a t t r a c t i v e f o r c o o l a n t pvirposes, b u t the low b o i l i n , g p o i n t means t h a t the f u e l r a p i d l y v a p o r i s e s . Since hydrogen i s the e a r t h ' s n i n t h most abundant e l e m e n t , a v a i l a b i l i t y i s no p r o b l e m , and t h e p r o d u c t i o n of gaseous hydrogen h a s developed on a vvide s c a l e ,
L i q u e f a c t i o n demands f a i r l y e x t e n s i v e equipment due t o t h e l a r g e h e a t e x t r a c t i o n n e c e s s a r y . Hydrogen e x i s t s i n b o t h an o r t h o and a p a r a form, and t h e n a t u r a l o r t h o - p a r a c o n v e r s i o n i n t h e l i q u i d p h a s e r e l e a s e s more t h a n s u f f i c i e n t energy t o b o i l off t h e c o n v e r t e d p a r a h y d r o g e n . F o r l o n g - t e r m s t o r a g e , the p a r a form i s e s s e n t i a l , and c a n be a c h i e v e d by c a t a l y s i s ,
4 . 3 • S l u r r i e s
I n s p e c t i o n of F i g . 20 shoiTs t h a t the i n t r o d u c t i o n of s o l i d m a t e r i a l s i n t o a hydrocarbon c a r r i e r f l u i d r e s u l t s i n a p r o g j ' e s s i v e reduction i n s p e c i f i c ir/rpulse e x p r e s s e d on a f u e l mass b a s i s , and a s l i g h t i n c r e a s e o n l y i n s p e c i f i c impulse e x p r e s s e d on a b a s i s of m i x t u r e w i t h a i r . ITnen b a s e d on a i r a l o n e , t h e c o n c e n t r a t i o n of s o l i d i s s e e n t o be e f f e c t i v e i n i n c r e a s i n g p e r f o r m a n c o , so t h a t slvirry f u e l s a p p e a r t o be a t t r a c t i v e f o r r a m j e t a p p l i c a t i o n s . F o r r o c k e t o p e r a t i o n i n a gaseous a t m o s p h e r e , volume impulse i s s i g n i f i c a n t , and c a l c u l a t i o n s made ivitli t h e f u e l - a i r r e s u l t s from F i g , 20 shov/
an i n c r e a s e of 12,75/^ on volurae impulse f o r t h e 8C^ n a g n e s i u m - o c t e n e s l u r r y :
-PraiDollant Mi?^i!r^®-X°i^ïïP. .iEJEMiSS,
L b , s e c , / g a l l o n of l i q u i d p r o p e l l a n t S t o i c , o c t e n e - 1 & a i r 1,420 S t o i c . 8C^ b o r o n / 2 0 ^ o c t e n e - 1 & a i r 1,5^7
S t o i c , 8C^ al\.iminiun/20^ o c t e n e - 1 & a.ir 1,552 S t o i c , 8C^ maf5.iesiuiiy2C^ o c t e n e - 1 &. a i r 1,601
Similar improvements v/ould be exj^ected vTith oxjrgon as the
oxidant,
Combustion t e s t s reported i n Ref. 11 shov/ed tlicit
magnesiim-hydrocarbon s l u r r i e s burned r e a d i l y and e f f i c i e n t l y v/ithout deposition
even under conditions v.hore the l i q u i d hydrocarbon alone would not
b u m . Successful f l i g l i t t e s t s were a l s o conducted with a sitiall ramjet.
Boron s l u r r i e s burned l e s s r e a d i l y than petroleum-based j e t f u e l s , and
gave r i s e to objectionable deposits i n the combustor. Aluminium
s l u r r i e s generally suffer from deposition t r o u b l e s , vmless
porous-VTalled combustors are used,
On the handlin.g s i d e , the maximum mass cap-centration of s o l i d i s
l i m i t e d to about 50 or 6C% from considerations of pumpability and of
s e t t l e m e n t . Research a t Cranfield i n t o the preparation and physical
behaviour of a 50^ mass s l u r r y of powdered alundrium, of s i x micron
average p a r t i c l e s i z e , i n avio.tion kerosine showed t h a t suspension
s t a b i l i t y can be achieved by the incorporation of 0,5^ m^iss of
aluminium octoate v/hich, on h e a t i n g , increases the v i s c o s i t y of the
kerosine by the formation of a g e l ( R e f . 1 2 ) . Since the r e s u l t i n g
s l u r r y i s d i f f i c u l t t o pump or spray, a surface-active agent must sJlso
be added to a c t as a d i s p e r s a n t . One c o n s t i t u e n t of the s u r f a c e
-active agent dissolv^js i n the k e r o s i n e , and the remainder a t t a c h e s
i t s e l f to the surface of the s o l i d p a r t i c l e s , so r e l i e v i n g the i n t o r
-f a c i a l t e n s i o n , and promoting -f l u i d i t y . O-f the materials t e s t e d a t
Crajifield, polyoxyethylene s o r b i t o l mono-oleate was found to be
s a t i s f a c t o r y when used i n a mass c a i c e n t r a t i o n of 0 , 1 ^ i n the above
slvury,
Abrasion of pumps and systems would be a severe problem v/ith
s l u r r i e s , p£irticularly v;hcn using boron v/hich i s second only to diamond
i n hardiiess.
4.4. isil^SÜ^
Of the possible replacements for carbon in the hydrocarbons (i.e, carbon-hydrides), only beryllium, boron, aaid lithium offer a greater energy content. The heating value of beryllium hydride is some éO^ higher than that of aviation kerosine, but beryllium is not a pre^cticable fuel in viev/ of its exti-eme scarcity, toxicity, and the demands of the nuclear power industry. The heating value of lithivim hydride is little better than that of aviation kerosine. The boron hydrides (i,e, boranes) of interest have been shovm to be diborrine
(B^H/-) , pentaborane (Bp-Hg), eoid decaborane (B. JI . ) , which exist in
the gaseous, liquid, and solid phases respectivGQy at normal teniperatures, Diborane condenses at -92 G,, pcntaborone freezes at -4é,7 C, and
boils at 63 C,, whereas decaborane melts at 99.4 C, Boron itself melts at about 2200°C.
Gaseous diborane i g n i t e s spontaneously i n a i r , and i s r e a d i l y
explosive. I t s v i o l e n t r e a c t i o n i n the presence of moisture malces
handling d i f f i c u l t , but offers p o s s i b i l i t i e s of underwater propulsion,
As with many boron coiripounds, diborane i s t o x i c , and has a disagreeable
odour.
As a l i q u i d f u e l , pentaborane i s more a t t r a c t i v e from the handling
viewpoint. I t does not i g n i t e spontaneously i n a i r when cool, but
may explode when h o t . I t i s l i a b l e to decorr^ose on standing, giving
off hydrogen gas and forming heavier boranes which are s o l i d but
dissolve i n the parent pentaborane. This doconiposition i s slew even
i n the presence of moisture, and r e s u l t s i n b o r i c a c i d . Pentaborane
also i s t o x i c . Decaborcne i s s o l i d and has a lovrcr performance, but
i t i s l e s s t o x i c and has a low vapour pre,r>siire. Most s t r u c t u r a l metals,
except lead and copper, appear t o be compatible vn.th the boranes.
Mild s t e e l cannot be used f o r storage v e s s e l s due t o embrittlement
following the I'elcase of hydrogen gas,
I n viev/ of tliese p r o b l e n s , a valuable expedient i s the compounding
of carbon Vvdth boranes t o produce carboboranes. The carbon i s added
i n the form of alkj'-l groups, such as the etliyl r a d i c a l . Performance
suffers s l i g h t l y , but allq^'-lation c r e a t e s a s t a b i l i s i n g e f f e c t and
reduces the t o x i c i i y of the combustion gases. I n /mierica, tlie O l i n
-Mathieson Chemical Corp. produces carboboranes under the designation
HEP, i , e . higli energy f u e l s , whereas the Gallery Chemical Go.
designation i s HiCal,
I n ramjet a p p l i c a t i o n s , fuels of a pyrophoric (spontaneously
i g n i t i b l e in a i r ) nature are a t t r a c t i v e , since t h e i r inflammable
mixture range i s much v/idcr tlian t h a t of hydrocarbon f u e l s , and they
b u m a t lovTer p r e s s u r e s . These p r o p e r t i e s pcrr.iit the use of smaller,
l i g h t e r , and more r e l i a b l e ramjets, capable of ecotiomic high-speed
operation a t very high a l t i t u d e s . They may a l s o be used as turbojet
i g n i t i o n f l u i d s , and as a n t i - s c r e e c h a d d i t i v e s i n rocket engines.
The pyrophoric natvire of some boron compounds, although giving r i s e
t o handling d i f f i c u l t i e s , i s a t t r a c t i v e pcrformance-v/ise, and t e s t s
are proceeding v/ith t r i e t h y l borane
(TPIB), as well as t r i e t l i y l
aluminium (TEA) and trimethyl aluminium (ÊÏA) .
Of th^ nitrogen h y d r i d e s , xoseful ijerformance i s exhibited by
ammonia ( N H , ) , hydrazine (NpH, ) , unsymmetrical dimethyl hydrazine
or UDMH (N„C J I o ) , and 'Hyd^me' , v;tiich i s a éO/40 mass mixture of
UDMFI and diethylenetriamine (DET) . In g e n e r a l , nitrogen hydrides
are t o x i c , but t h e i r conibustion products are i n e r t . They are
inflammable, and form explosive mixtures v/ith a i r . Ammonia has a high
vapour pressiore, but a. hi,gli s p e c i f i c h e a t .
Hydrazine is spontaaieously ignitible with nitric acid, and v/ith
HTP. Hydrazine ±a generally used in the hydrate f o m , which is equivalent to a hydrazine concentration of Sl^o,
4.5. Ojd-dants
Discussion of the properties of oxidants has received a thorough treatment in the literature but, for completeness, a brief review is includedi.. Fluorine is considered in this category since its reaction vrith a fuel results in the evolution of energy. Fluorine is the most powerful ' oxidisin,g' appnt known, reacting v/ith practically all organic and inorganic materials. It provides the highest specific impulse and, v/ith its high specific gravity, the hif^hest volume impulse. It has a very high vapour pressure, end. its low boiling point makes necessary the insulation of storage oqixlpment. Reaction vd-th nany metals is slov7 at ambient temperature, resulting in a protective film of metallic fluoride. For this reason, sixch metals as iron,
aluminium, magnesium, copper, m d brass are coüipatible if the temperature is controlled. Fluorine is very toxic.
Liquid oxygen also provides rapid reactions, high flame speeds, and high combustion ter^iperatures. It, too, is a cryogenic material,
vdth its attendant problems of vapour loss and laT-tempcrature coKipatibility, Vacuum jacketing reduces vapour losses to about 0,1^ per day in bulk
storage, but losses from current designs of aircraft tank are of the order of ~/fo per hour. Thermal insulation reduces tanJ-cage costs, but the most effective insulcmts are also inflanT.iable. The Biixture ranges of inflammability of most fuels are widened considerably in combination v/ith oxygen, compared, vdth those in air, end the fire hazard is further increased since the freezing of v/e.ter from the surrounding atmosphere makes for easier generation of sparks. Since the spontrmoous i.gnition temperature of most materials is lower in oxygen tlian in air, liquid oxygen cannot be used as a wall coolant. Availability of oxygen is no problem, and liquefaction can be carried out in mobile plants if necessary. Liquid ozone is more attractive from the j^erformance and density points of view, giving an increase in specific impulse of about 1C^. Explosive decomposition is likely on contamination by dust or other small in^purities, unless the mass concentration of ozone is reduced below 30^ by dilution with ojqygcn. The higlier
volatility of oxj-'gen results in a progressive rise in ozone concentration during storage, and this risk is not eliminated by ixjfrigeration siiice ozone and oxygen then tend to separate out. Ozone is extremely toxic above a concentration of about 0,1 parts per million,
High Test Peroxide pro\ddes a useful level of performance with most fuels and is stable in storage. If protected from temperature rise and the in,grcss of contiiminants, the decomposition loss does not exceed about •]% per annum in temperate zones, and 3^ in the
(80 to lOC^ concentration), and the vapour is very active, so that flushing of the system is necessary after dr-aining, A number of materials,
including silver and the pennanganates, catalyse decomposition and permit the use of HTP as a monopropellant or a source of auxiliary power.
Pure aluminiuïi and stainless steel are suitable structural materials for
use with HEP.
High strength ('fuming') nitric acid, which is prefixed 'red' when it contains up to 30/c of dissolved oxides of nitrogen, gives a performance comparable v/ith HTP. It is slightly hygroscopic, and its corrosivity to ali.iininium increases with water content, It reacts with most organic materials, and may cause fires on contact with wood
or other cellulose products. Fuming nitric acid does not inflame, but will ignite spontaneously with some fuels. The low surface tension
enables the liquid to creep over surfaces, and the fumes evolved are highly toxic. Aluminium, stainless steel, and high silicon irons are
suitable structural materials, pai'ticularly when corrosion inhibitors are incorporated in the acid.
The nitrogen peroxides, in the form of the mixed oxides also give comparable perforraance. A typical mixture of 70% NpO, and 30^ NO
appears to meet the engineering requirements of high chemical stability, good corrosion resistance, low freezing point, and reasonably low
vapour pressure. V/ith dry peroxides, carbon steel and many other structural ms.terials are suitable, but stainless steels are necessary \mder wet conditions,
5. Conclusions
i. A high heat of reaction is a basic requirement for a rocket propellant pair, but this is modified by the thermodynamic behaviour of the combustion gases in the propelling nozzle. Consequently, any realistic indication of propellant performance must be based upon the resultant thrust.
ii. A high combustion temperature is necessary for high thrust, but cooling problems are intensified, and certain propellants only are suitable as coolants.
iii. A general increase in specific impulse is found v/ith a decrease in specific gravity of the propellant mixture,
iv. Liquid hydrogen offers the ultimate in specific impulse for chemical propellants, but handling difficulties are severe. V. Liquid fluorine is the most energetic reacting medium available
for use with rocket fuels.
vi. A general increase in volume iiipulse is fo\ind with an increase in specific gravit.y of the propellant mixture.
vii. Hydrocarbon fuels still represent a practicable compromise for rocket applications,
viii. Non-carbon hydrides appear to be the next most attractive fuels for rocket applications,
ix. Slurry fuels are more suited to air-breathing propulsive •units, Some typical rocket applications of liquid bi-propellants are shown in Table 3.
6, Acknov/ledgements
Pigs. 9 and 11 included with permission from G, P. Sutton, 'Rocket Propulsion Elements', I95é, John Wiley & Sons Inc.
•
Table 3 prepared by R W, Stutchbury, Aircraft Propulsion Department, College of Aeronautics, Cranfield.
7, References 1. Perchonok, E. 2. Offtermatt, ¥ F. 3. Anon. 4. Vïarren, F,A. 5. Sutton, G.P.
6. Rocketdyne Co., U.S.A.
7. Callaghan, Lt. W.M,, U.S.N, 8. Goodger, E.M, 9. Conn, M.E., cS: Dukek, W.G. 10. Breitwieser, R., Gordon, S., & Gammon, B. 11. Pinns, M.L,, eisen, W.T., Bamett, H.C. , cS: Breitwieser, R. 12. Henshaw, J.B.
Perfonnance Evaluation of Ramjet
Propellants. Chemistry of Propellants Meeting. AGARD, June, 1959. Paris.
Tables of Bond and Resonance Energies for Estimating Standard Keats of Formation. Cal. Inst, Technology, AD-24936. 1953c
International Critical Tables. Vol. V. p. 163. McGraw Hill Book Co. New York. Rocket Propellants. p. 82 Reinheid Publishing Corp, New York. 1958, Rocket Propulsion Elements, p. 54.
John W^iley & Sons. Inc., New York. 195^. Theoretical Performance of Rocket
Propellant Combinations, Chart published 1.1.59.
Combustion of Slurry Fuels. College of Aeronautics Thesis, June, 1958.
Unpublished.
Aviation Fuel Problems at High Altitudes and High Aircrai't Speeds, College of Aeronautics Report No. 119.
High Performance Hydrocarbon fuels for Supersonic Propulsion. S A E, National Aeronautic Meeting, MarcV'April 1959. New York.
Summary Report on Analytical Evaluation of Air and Fuel Specific Inipulse
Characteristics of several nonhydrocarbon jet-engine fuels. NACA RM E52L08.
NACA Research on Slurry Fuels. NACA Report 1388, 1958.
Research into Slurry Fuels. College of Aeronautics Thesis, June, 1959. Unpublished,
APPENDIX Derivation of Specific Impulse
Cross-sectional areas A, Inlet „ Velocity i" •—•> e Exhaust Ambient V Velocity Pressure Inlet Plane Exit "fhrust F Plane
The total force exerted on a duct in the direction of flight = F + P ( A - A . ) = change of stream thrust = (m V + p A ) - (fn.V. + p. A . ) .
^ e e -^e e'^ ^ i i ^ -^i i'
In rocket engines, V., A., and ih. are all zero, hence,
F = fli V + ( p - p ) A . e e ^-^e -^a' e
In the optimum case, the nozzle v/ill be designed to give an exit pressure equal to ambient, and the thrust expression reduces to
:-F = fh V . e e
Thrust is a measure of the size and performance of the rocket engine, but a comparison of the performance of propellants is available in the thrust produced by unit consumption rate of propellant mass. This parameter is termed 'specific impulse', and represented by the
symbol I^^.
Hence. I = F/ih = V , in consistent units, * sp ' e e'
Since it is more convenient to deal with thrust in lonits of pounds force, and with propellant mass in units of pounds, this expression may bo written in terms of the British mixed (Engineers') unit system, as follows :
-T _ g (Lb) ^P fii^ (lb/sec)
Therefore, I = ^r}?-i,jL /-i-u; \, ^ consistent B r i t i s h
' sp 32,174 \ ( l b / s e c ) ' absolute u n i t s ,
and also, I^^ = 3 2 ! ^ & ( s l ^ s e c ) - ^ consistent Perry units,
V
= ^
-^
, where 32,174 is the conversion factor
J'-» I 1'+between the various systems of units.
V can be determined by applying the Bernouilli expression to the
inlet and exit planes of the propelling nozzle :
-• 5 - V ^ + C T = - 5 - V ^ + C T , for isentropic conditions
'
^
c
p
o ' ^ ' e ^ p e '
where subscripts c and e represent combustion chamber (nozzle inlet)
and nozzle exit planes respectively. Since V is nominally equal to
zero, this expression may be rearranged to
:-V = 2 C T - 2 C T
e \| P c p e
In an ideal gas, C is constant, hence
:-IT
This may be developed from thermodynamics as follov/s :
-V
e
2 y GT • —
[-(P/Pc) " ]
- -xi y - 1 Mwhere M is the mean molecular weight of the product gases,
This expression is sometimes vsrritten as V = 0 *GT /k,
since the term comprising 9 appears in a number of other performance
parameters. Hence, _
FZ^
Per simplicity, the chemical composition of the combustion gases may be assijmed to remain constant throughout the expansion in the propelling nozzle. Specific inipu].3e calculated under these
'frozen equilibrium' conditions is slightly less (approximately ^o) than that obtained in the 'shifting equilibrium' case which includes the heat released by recombination of the free atoms and radicals due to the falling temperature.
For operation within a gaseous atmosphere, where aerodynamic drag is significant, specific impulse may be based upon propellant liquid volume, in the same way as heating value. This is expressed as the product of I and density, and is termed the volime impulse (or density impulse;?
In the ramjet engine, designed for operation \\dthin the earth's atmosphere, fuel effectiveness may be expressed in terms of the air specific impulse, v/hich is the thrust produced v/ith unit mass
Consumption rate of air. This parameter also indicates the size of engine necessary for a given tlirust level.
Therefore, I = Vo F-» x. /TU> \, ^ consistent B r i t i s h
> sp 32,174 A 3 ( l b / s e c ) ' absolute u n i t s ,
and also, I^^ = 32^,^74 fc (slu&/sec)' ^ consistent Perry units,
V
- .^^ .-:, , where 32,174 is the conversion factor 32,174
between the various systems of units,
V can be determined by applying the Bernouilli expression to the inlet and exit planes of the propelling nozzle
:-^ V :-^ + C T = - | V :-^ + C T , for isentropic conditions
where subscripts c and e represent combustion chamber (nozzle inlet) and nozzle exit planes respectively. Since V is nominally equal to zero, this expression may be rearranged to
:-V = 2 C T - 2 C T
e \1 P c p e
In an ideal gas, C is constant, hence
:-ir
V^ = 2 C Ï (1 - T / T ^ ) , e '^ p 0 ^ e c^
This may be developed from thermodynamics as follov/s :•
2 y GT "^ —
1 - ( P / P , ) y ^ - \ | y -1 M
where M is the mean molecular weight of the product gases. This expression is sometimes vsrritten as V = 9 ^GT /k, since the terra comprising 9 appears in a number of other performanoe
parameters. Hence, ^ (^
Per simplicity, the chemical composition of the combustion gases may be assumed to renain constant throughout the expansion in the propelling nozzle. Specific impuJ.se calculated under these
'frozen eqi:dlibrium' conditions is sliglitly less (ajjproximately ^b) than that obtained in the 'shifting equilibriijn' case which includes the heat released by recombination of the free atoms and radicals due to the falling temperature.
For operation within a gaseous atmosphere, where aerodynamic drag is significant, specific impulse may be based upon propellant liquid volume, in the same way as heating value. This is expressed as the product of I and density, and is termed the volume impulse (or density impulse;?
In the ramjet engine, designed for operation within the earth's atmosphere, fuel effectiveness may be expressed in terms of the air specific impulse, v/hich is the thrust produced v/ith unit mass
Consumption rate of air. This parameter also indicates the size of engine necessary for a given thrust level.
^ \ M A T E K L A L PROPERTY ^ ^ - v ^ ^ LOV.'ER ^ CHU/lb ^ ' " ™ ' 4 CHU/gall VALUE J ' r:-M<'ORi.A'.;c.:: E Ü J E X S.G, •'s 6 0 / 6 0 ° P ^aPtiC.HKAT fS CI!U/ 6 0 " ^ l b ° C . v;.i'OHi:;/iTiON lEAT CHU/lb ' B O i m r c POINT ° c FRiJSZIMC POIOT °C FIASH POINT °C a-ONTA:JrCUS IGN. T r i ï P . irr c AIK • ' 1 a t m . POLÏ
AVTOR RP-1 ryjH^'^ ETHANOL 1 0 , 2 5 0 1 0 , 2 8 0 1 0 , 2 7 0 6 , 6 5 0 6 2 , 0 0 0 8 3 , 0 0 0 3 6 , 9 0 0 5 2 , 7 0 0 1 , 0 0 1 , 0 2 1 . 0 9 0 . 4 2 0 , 6 O.f; O.eCÓ 0 , 7 9 2 0 . 5 0 0 . 6 1 67 2 0 4 1 5 0 , 1&5. 201 f^ 300 274 275 - 4 0 - 4 0 - 6 0 - 1 1 8 3 7 . 8 4 3 . 3 ^' 21 2 5 4 4 3 9 2 BORON AVTOR SIURRY 2 8 , 7 0 0 1 2 , 5 0 0 2 0 , 0 9 0 1 6 4 , 000 0 , 6 9 2 . 4 4 0 . 0 7 1 1,31 b . p . 2 . 3 3 b . p . 109 - 2 5 3 150»-- 2 5 9 150»-- 4 0 5 7 0 B^Hg BjHg m^ NjH^ raMH 1 7 , 4 0 0 1 6 , 2 0 0 5 , 7 2 0 4 , 7 2 0 7 , 8 7 0 7 4 , 8 0 0 9 8 , 8 0 0 3 5 , 0 0 0 4 7 , 7 0 0 6 1 , 4 0 0 1 . 5 5 1 . 9 0 0 . 2 4 0 . 2 7 0 . 5 8 0 , 4 3 0 . 6 1 0,61 1.01 0 . 7 8 b . p . 0 . 4 8 0 . 5 7 1 , 1 2 0 . 7 5 0 , 6 5 125 122 326 1 3 4 - 9 3 63 -53 1 1 4 63 - 1 6 5 - 4 7 - 7 8 2 - 5 8 52 1 651 2 5 0 LPg 1^2 HTP HNO, " ^ ~ \ 1.51 1.14 1.35 1.56 b . p . 0.37 0.4 0.61 0.42 b . p . 41 51 346 115 -188 -183 152 86 -220 -220 -22 -42 ^ - \ TY1'1C>\L
^ . ^ KIATERIAL PROPERTY ^'^-^,^,^^^
1
1
ü e; t-i a p is ic HEAT CONTENT I VïTTH LD„ s p 2 COMBUSTION SPECIFIC GRAVITY VAPOUR PRESSURE TKERtiAL STABILITY VISCOSITY -TEliTERATUBE SPECIFIC KEAT UNIFORJ.ITY AVAILABILITY & COST STORAGE, TRANSFER TRANSPORT OTiEALL RATI HG AIR-HHEATHING ROCSfflT HYDRO-CARBONS 3 4 3 3 1 2 1 2 4 1 1 2 2 " ^ 2 1 1 1 5 5 1 5 1 1 4 5 1 1 SLURRIES 3 5 5 1 2 3 4 2 5 2 4 4 5 NON-CARBON HYDRIDES 2 3 4 4 2 4 2 4 3 4 3 3 4 HYDRAZINE 5 2 2 2 4 4 3 4 2 3 2 4 3VEHICIE 'THOR' 'ATLAS' De-K. EROP, 'BDUE STREAK' SAUNDiES-ROE 'BLACK KKIGHT' L . R , B . A . 'VERONIQUE' AEROJET-GENERAL • A E R O H E E ' MARTIN 'viKnre' 'VANGUARD' STAGE 1 STAGE 2 STAGE 3 'NOVA' STAffl 1 STAGE 2 STAGE 3 STAGE 4 STAGE 5 ENGDfE ROCKETDYNE I f i - 7 9 SUSTAINER ROaCJTDYNE 2 - I E 9 3 BOOST 1 - I B 1 0 5 SUSTAINER ROII£-ROYCE 2 ENGDÏES BHISTOL-SIDDELEY GAiaiA, I K . 201 L . R . B . A . AEROJET REACTION MOTORS XIR 1 0 G . E . C . i AEROJET SEA-L=:VÏÏL TIÏkbST. Lb. 150,0C-0 1 6 5 , 0 0 0 6 0 , 0 0 0 ASSUMED 1 5 0 , 0 0 0 EACH 1 6 . 4 0 0 8 , 8 0 0 4 , 0 0 0 2 0 , 0 0 0 2 7 , 0 0 0 8,ooo(Anr 2,300(AI!r 1 4 X 1 , 5 0 0 , 0 0 c 1 , 5 0 0 , 0 0 0 8 0 , 0 0 0 4 0 , 0 0 0 6 , 0 0 0 FU3L KP-1 EP-1 KP-1 AS&mOSD HYDROCARBON KEROSINE TURPSrJTINE AKIUME ETHANOL KP-1 1 U.D.M.H. i SOL 1 HP-1 RP-1 STORABLE GAIDANT " ^ 2 ^2 ^ 2 ASSUiïED ^ 2 H . T . P . NITRIC ACm X NITRIC ACm X " ' 2 L0„ ITITRIC ACID ID " ' 2 " ' 2 " ^ 2
1 "^2
LEJJIDS H:-3»iAHKS I . H . B . M . • ( R . A . P . ) I . C . B . M . ( U . S , STRAT. AIR COI-a.AND) L . R . 3 . M . UPPt-E A.T:,!OSFHaRE & HE-SNTRY UPPER AB-iOSPHERE ( I . G . Y . ) UPPER ATl.;OaPHS«E SOUNDING SATELLITE CARRIER N . A . S . A , ADVANCED 1 SPACE-LABORATORY PROJECT K Hypergolic p a i r s Turpentine (Pinene) °^(^^(, t . p . 155°C, f . p . -55°C. S.G. 0,858 Aniline ^gHJJHg b . p . 184°C. f . p . - 8°C. S.G. 1.03m
r
m
m
z
iJ
r
z
o
E
m >z
o
O
z
m
m
c
:Dm
m
Reactants at stand;ird state :> Products at changed state Products at changed state + A H = Products at standard state
A H = Heat added to regain standard state after ^ reaction from standard state.
= negative, if reaction exothermic
A H = ZAH.O P r o d u c t s - 2AH„ R e a c t a n t s
r I 1
Thus, Max. Neg, AH r e q u i r e s :
-MAK. NEG. AH^ PRODUCTS cS: MAX. POS. AH^ RB/iCTANTS
N^
V
AH.p = Heat of F o r m a t i o n
Element M o l e c u l e s > . Gaseous Atoms ^ Compounds
.._^., ^J Atomisation Heat Endothezroic Positive Bond Energy Exothermic Negative AH„ = Atomisation Heat - Bond Energy
2 A H
'a S D(X - I)
>/
PRODUCTS MAX. rtEG. AH^
J£
-T'X
REACT/il^S W . . POS. AH^ ( o r min. n e g . ) MIN. POS. ^ A H , C (11.3) better than H2 (52.1)ïvIAX. POS. ^D(X-Y^ H2O ( 1 2 . 3 ) b e t t e r t h a n CO ( 8 . 0 ) or 1 CO2 ( 8 . 0 ) MM. POS. 2AH. H2 ( 5 2 . 1 ) b e t t e r t h a n C ( 1 1 . 3 ) — I "MIN. P O S . £ D ( X - Y T " Min. (C-H)b-»nds C ss C better than C - C
All values in kcan/s^.
5
&ü
ul I -2 -3 : C O . CO-HjO "ALCOHOLSMAX. NEG, A H ? PRODUCTS
I?
X <i
i
ul X - 6 - I O -12 - 1 4 4 5 6 NO OF CARBON ATOMSFIG. 3. HEATS OF FORMATION AND REACTION OF HYDROCARBONS NB AHf VALUES DERIVED FROM REF 2
CODE F U E L DEn9.R.D.N9 AVGAS AVIATION GASOLINE 2485
AVTAG AVIATION WIDE-CUT GASOLINE 2486 AVTUR AVIATION KEROSINE 2 4 8 2 AVCAT AVIATION HIGH FLASH KEROSINE 2 4 8 8
R.P-I ROCKET KEROSINE
IO-4 IO-2
lO-O-i
o o z < I/I O I 9-8 • 9.6 . AVGAS SPECIFIED SG. RANGES^
h-,_je
I AVCAT I AVTUR I I AVTAG I 4 5 0 -GAS Oi-S DIESEL FUELS FUEL OILS 7 0 0 7 O-B SPECIFIC GRAVITY 0.9FIG. 4. TYPICAL HEATING VALUES OF PETROLEUM FUELS
z X o 3 5 0 2 0 0 0 -SCO I500 lOOO 5 0 0 -250 Ul z z 3
i
| 2 0 0 | -ISO 3 I d a z O ICO f so l O SPBQFIC GFIAVITY Ï O1 0 1-5 SPECIFIC GRAVITY
FIG. 6. VAUUES OF PERFORMANCE INDEX IC
[
/}
\ lO'V
y
e7
IQ-2/^('-t^l^l
icr T~ IÖ-^ lO PRESSURE RATIOFIG. 8. INFLUENCE OF PRESSURE RATIO UPON FUNCTION G (SEE ALSO REF 4.]
1 ^.,^^^
1 PRESS \ ^ ~ \ ^ \ 1 / 1 / 1 / ' V E L Y / / / / / 1J »j
^ / 1N
1
-2 1 3 0 0 0 2 5 0 0 o'' I 2 0 0 0 ^ I500 6 0 0 0 4 0 0 0 rFIG. 7 TYPICAL VARIATIONS OF
T . R & V IN A ROCKET NOZZLE
2 4 0 ZOO 6 0 p, / 20 y = /^ ^ 2 ^ 1-3
J
/^y\
'^ 1-25 1 I20 I40 4 0 5 0 BO lOOCOMBUSTION TEMPERATURE/ MEAN MOLECULAR RATIO - ^
FIG 9 INFLUENCE OF ^ RATIO AND V UPON Isp AND Vt (REF S BY PERMISSION]
op
o
o
6 8 0 II.OOO ?
FIG. lO. VARIATION OF GRAVIMETRIC HEATING VALUES AND COMBUSTION TEMPERATURES WITH CARBON/ HYDROGEN RATIO
3 5 0 0 a. 5 Ul I-z g D CD O O 2
a
3 0 0 0 -2 5 0 0 2 0 0 0 -3 4 0 3 2 0 3 0 0 28 O f^ Ï 3 0 0 p.5i.a ( A ) L H 2 / L 0 2 3SOO 2 4 6 B OXIDANT/ FUEL MASS RATIO(^ s 2 0 A T M
2 3 0
( B ) H Y O R A Z I N E / L 0 2
0-25 O'S O 75 I O OXIDANT/FUEL MASS RATIO
FLUORINATION
• HTP A RFNA
0 - 4 0 - 6 0 - 8 l-O 1-2 SPEQFIC GH»/fTY OF PROPELLANT MIXTURE
FIG. 12. RELATIONSHIP BETWEEN PEAK SPECIFIC IMPULSE AND SPECIFIC GRAVITY (DERIVED FROM REF 6)
2 0 0 0 -SHIFTING EQUILIBRIUM Pc = lOOO pi.i.a. ( i . l4-7p.si.a. X L F j + L 0 2 • HTP A RFNA
I
8. 2 8 0 0 2 4 0 0 0-8 l-O 1-2 1-4 1-6 SPECIFIC GRAVITY OF PROPELLANT MIXTUREP^ • l O O O pisxa. Rt= 14-7 pi»i.a. SHIFTING EOUIUBIUM (+) FROZEN EQUILIBRIUM (•)
FUELS PLUS UQUIO OXYGEN
FIG. I3A. RELATIONSHIP BETWEEN PEAK VOLUME IMPULSE AND SPECIFIC GRAVITY
(SMALL SCALE PLOT ) ( DERIVED FROM REF 6)
0 « 0 - 7 0 « 0 - 9 l-O l-l SPECIRC GRAVITY OF PROPELLANT MIXTURE
FIG I3B RELATIONSHIP BETWEEN PEAK VOLUME IMPULSE AND SPECIFIC GRAVITY
RG. 14. COMPARISON OF PEAK SPEORC IMPULSE VALUES FOR Bl - PROPELLANTS (DERIVED FROMRER6)
E
I
^^
WITH ALUMINIUM SLURRY FUEL (REF.T) O 5 AVTLIR SPECIFICATION, LIMIT ^ ^ k^2 L^-ift?, - 2 0 0 HSO -lOO -50 TEMPERATURE t
FIG 18 TEMPERATURE VARIATION OF PROPELLANT VISCOSITY
20 4 0 6 0 6 0 lOO *4MASS CONCENTRATION SOUD IN OCTENE-1
FIG 2 0 SPECIFIC IMPULSE VALUES FOR SUJRRY FUELS (DERIVED FROM REF6)
o^ UJ o: Ul a. 2 UJ 4 0 0 3 0 0 2 0 0 IOC O - lOO [- \ A.S.TM. \ SPONTANEOUS \ IGNITION \ TEMPERATURES \ \ ^ k — , ^^^ ^^^^ DISTILLATION RANGES " o ^ < o p ^ 1 r i ^
r
^ ^ — ^ ^ . - ' ' ' ^ . . - " . ' - ' • —1
1 AVTAG ( er > < 1r
^L
^ - ' ^ ^DISTILLATION LIMIT fcRACKINCl
~T - 1 1 H-5
i
r °
!5 13 -^ •Jë
- I hl Ulc
o ^ » ^ ^ • ^ — _l 5 - 1ë
b. ^ ^ - ^ -J JL f U P P E R J R I C H ^ ^ I ^^^^ ^•^"^ INFLAMMABLE RANGE ^^ " " ^ ^ ^ _ ^ . * ^ - - " " T S W E R WEAK I „ « ^ ^ ^ A S H POINTS -AMBIENT y ^ ^ / ' ^ R E E Z I N G POINTS 1 AVCAT I AVTURL
1
11
0 - 7 SPECIFIC GRAVITY 0.9FIG. 19. TYPICAL VALUES SHOWING TRENDS IN PETROLEUM FUEL PROPERTIES ( R E F 8 . )
