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Combustion performance studies of aluminum and boron based composite solid propellants in sub-atmospheric pressure regimes
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Propulsion and Power Research 2019;xxx(xxx):1e10 http://ppr.buaa.edu.cn/
Propulsion and Power Research w w w. s c i e n c e d i r e c t . c om
ORGINAL ARTICLE
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Combustion performance studies of aluminum and boron based composite solid propellants in sub-atmospheric pressure regimes Pratim Kumara,*, Mayank Varshneyb, Aniket Manashc a
Department of Aerospace Engineering and Applied Mechanics, IIEST, Shibpur, Howrah, West Bengal, 711103, India b Indian Institute of Technology (Indian School of Mines), Dhanbad, Jharkhand, 862004, India c Aryabhatta Center for Nanoscience & Technology, AKU, Patna, Bihar, 800001, India Received 30 October 2018; accepted 24 September 2019 Available online XXXX
KEYWORDS AP/HTPB; Metallic/non-metallic fuel; Catalyst; Sub-atmospheric pressure; Burning rate; Base-bleed
Abstract The aim of present study is to investigate the burning rate, ignition delay, and flame characteristics of ammonium perchlorate (AP)-hydroxyl terminated poly-butadiene (HTPB) [AP/HTPB] based composite propellants (CSP’s) in sub-atmospheric pressure regimes (13 kPae100 kPa). Several fuels and catalyzed were used to evaluate their effects on the combustion characteristics of AP based propellants in sub-atmospheric pressure regimes. In fuels, aluminum (Al) and boron (B) were selected as metallic and non-metallic fuel respectively. While in catalyst, butyl ferrocene (B.F.) and ferric oxide (F.O.) were selected as liquid and solid catalyst respectively. Apart from these, other ingredients that were used are di-octyl adipate (DOA), toluene di-isocyanate (TDI), and glycerol. The article throws some light on the burning rate and ignition delay properties for these new classes of prepared propellant samples. At subatmospheric pressures, all propellants are susceptible to irregular burning with the ejection of soot’s, fumes, and unburned particles. F.O. based catalyzed propellants can sustain its combustion up to the lowest pressure. ª 2019 Beihang University. Production and hosting by Elsevier B.V. on behalf of KeAi. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/licenses/by-nc-nd/4.0/).
*Corresponding author. E-mail address: [email protected] (Pratim Kumar). Peer review under responsibility of Beihang University.
Production and Hosting by Elsevier on behalf of KeAi
1. Introduction From last several decades there is a remarkable increase of interest in the field of sub-atmospheric pressure combustion studies of composite solid propellants [1e5]. The
https://doi.org/10.1016/j.jppr.2019.09.001 2212-540X/ª 2019 Beihang University. Production and hosting by Elsevier B.V. on behalf of KeAi. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/licenses/by-nc-nd/4.0/).
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main application of sub-atmospheric pressure combustion studies are in the field of base bleed applications of missiles and military shells [6e10], propellant flame structures [11,12], low pressure deflagration limit (LPDL) [13], accurate prediction of early stages of motor ignition transients [14], and start stop of solid rocket motors [14]. Subatmospheric pressure conditions also allows better investigations over the combustion mechanism and flame structure since, (a) gas phase reaction zone consisting mainly of premixed flame (as against competing premixed and diffusion flames at rocket operating pressures), (b) less severe temperature gradient at the burning surface, and (c) better spatial and temporal resolutions. At sub-atmospheric pressure regimes, the diffusion processes became much faster and chemical reactions are slower at lower pressures [15]. It is common for the time and distance scale of flame zone to increase with decreasing pressure. Combustion studies in sub-atmospheric pressure are of interest for flame zone structures in cognizance that controlling chemical kinetics reaction steps and product gas compositions changes with pressure. Solid propellants which have the ability to burn in vacuum at high burning rate can act as a gas generator for reducing the partial vacuum created at the projectile base during flight [16]. The combustion gases from the combustion chamber escape through an unchoked port in the projectile base. Due to this, drag reduction and consequently range enhancement of military shells takes place. This process is known as base-bleed. For effective base bleed (BB) applications of missiles and shells, the propellant must have to sustain its combustion and also to burn more efficiently at sub-atmospheric pressure regimes. To initiate the proper ignition and to increase the burn rate of solid propellants in sub-atmospheric pressure [17e19] is an area of exciting research and have promising future prospects. Although composite solid propellants, as compared to double base (DB) propellants, burn more efficiently at low pressures, the combustion efficiently falls drastically below certain sub-atmospheric pressures culminating in extinction. The pressure at which extinction occurs is termed as low pressure deflagration limit (LPDL), as against the high pressure deflagration limit (HPDL) demonstrated in certain propellants [20]. Excessive fuming, soot generation, and ejection of unburned particles are some of the common characteristics of propellant combustion at sub-atmospheric pressure regimes. These burning characteristics show the poor combustion efficiency at sub-atmospheric pressure regimes. At some pressure, propellant can’t sustain its combustion and this limiting pressure is known as LPDL [21]. As per the Beckstead-Derr-Price (BDP) model, the primary diffusion flame controlled the combustion at the low pressure end, while the final diffusion flame along with the AP decomposition flame controlled the combustion at the high pressure end. In the present paper, an attempt has been made to investigate the burning rate, ignition delay, and flame structure of 18 ammonium perchlorate (AP, NH4ClO4)
Pratim Kumar et al.
based composite solid propellants in the pressure regimes of 100 kPa (760 mmHg) to 13 kPa (100 mmHg). Two ferrous based catalysts were used in the study viz. ferric oxide (F.O.) as a solid catalyst, and butyl ferrocene (B.F.) as a liquid catalyst [22]. Two fuels were also used viz. aluminium (Al) as a metallic fuel, and boron (B) as a nonmetallic fuel. Present study will be useful for obtaining a better composition of AP based propellant for increasing the range of military shells by decreasing their base drag using a gas-generator unit which can sustain its combustion in subatmospheric pressure regimes.
2. Experimental methodology 2.1. Composition, formulation, and preparation of propellant samples As discussed previously, a solid propellant basically consists of oxidizer, fuel, catalyst, and other additives. The selected components for propellant processing are given below: 1. Fuel binder: Hydroxyl-terminated poly-butadiene (HTPB) (supplied by propellant fuel complex, Vikram Sarabhai Space Centre (VSSC), Thiruvananthapuram). Currently HTPB is being extensively used as a polymeric fuel binder because of its capacity to load higher solid filler concentrations, and superior performance. 2. Oxidizer: Ammonium perchlorate (AP) (supplied by Tamil Nadu Chemicals Ltd., Madurai). A unimodal particle size of 74 mm was used in the study. 3. Plasticizer: Dioctyl adipate (DOA) (supplied by Chemporium Co., Mumbai). DOA was used as plasticizer in the present work in order to maintain easy process ability and good mechanical properties of the propellant. 4. Curing agent: Toluene di-isocyanate (TDI) (supplied by Merck Co., Mumbai). It helps in curing of HTPB. 5. Cross linking agent: Glycerol (supplied by Merck Co., Mumbai). It helps in the formation of matrix type structure between fuel and oxidizer particle. 6. Metallic fuel: Aluminum powder (fine, <75 mm) (supplied by: Sigma Aldrich). 7. Non-metallic fuel: Boron powder (amorphous powder, supplied by Thomas Baker, Mumbai). 8. Burning rate modifiers: Ferric oxide [5 mm (solid state), supplied by Central Drug House (CDH)], and butyl Q3 ferrocene (liquid state) supplied by Sigma Aldrich. A total of 18 propellant compositions were formulated using above ingredients in different proportions. Oxidizer percentage varied by 80, 75, and 70, while fuel percentage varied by 20, 25, and 30 in the propellants. Catalyst percentage was kept 2% to the total weight of the propellant sample. Composition of the prepared propellant samples is presented in Table 1. For preparation of propellant samples, requisite quantities HTPB, DOA, and glycerol were mixed thoroughly
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Table 1
Composition of prepared propellants.
Propellant No.
AP/%
HTPB/%
DOA/%
TDI/%
GLYCEROL/%
Al/%
B/%
F.O./wt%
B.F./wt%
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
80 75 70 75 70 70 80 75 70 75 70 70 80 75 70 75 70 70
14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36
4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36
1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13
0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15
0 5 10 0 0 5 0 5 10 0 0 5 0 5 10 0 0 5
0 0 0 5 10 5 0 0 0 5 10 5 0 0 0 5 10 5
0 0 0 0 0 0 2 2 2 2 2 2 0 0 0 0 0 0
0 0 0 0 0 0 0 0 0 0 0 0 2 2 2 2 2 2
inside a mixer till the whole mixture becomes homogeneous. Afterwards, requisite quantities of fuels and catalyst were added in the mixture with continuous mixing to maintain homogeneity. At last, TDI was added and mixing continued for another 30 min. The final slurry was casted in a stainless steel casting box. In the casting box, 12 slots each of size 60 6 6 mm3 were manufactured for obtaining the casted propellant strands. The propellant slurry filled casting box was then kept in a vacuum electric oven at 60 C 2 C for a period of six days to allow it to cure into a polymeric state and to attain required mechanical strength. After six days, the moulds were removed from the oven and allowed to cool at room temperature. The propellant blocks thus obtained were kept in a polythene bag and stored inside desiccators to avoid moisture absorption. Finally, the obtained propellant strands were inhibited from all four sides by quick drying polyethene enamel for obtaining linear burning.
needed for propellant ignition, and for timer reading (shown in Figure 5).
2.2.2. De-pressurization system It consists of a high vacuum pump which was connected at the top of the surge tank through a high vacuum line. At the top of the surge tank, a mercury manometer was attached to read the vacuum level inside the combustion chamber. The advantage of the surge tank is to minimize the pressure rise and fluctuations during the combustion of propellant strands.
2.2.3. Ignition system Hot wire ignition system was employed for igniting the propellant strand. Nichrome wire with resistance of 20 ohm per cm and of diameter 0.27 mm was passed at the top of the propellant strand, and was heated till the propellant sample get ignited by passing a current of 1.5 amp at 30 V.
2.2. Experimental set-up The experimental setup is also known as sub-atmospheric strand burner set up. The detail of experimental set-up is shown in Figure 1. The details of different parts of the experimental set up are discussed one by one.
2.2.1. Combustion chamber A cylindrical bell jar of borosilicate glass of capacity 2000 cc was used as a combustion chamber. It was fixed on a stainless steel plate with the help of retainer rings, nuts, and bolts. Polyvinyl chloride (PVC) washer and a thin layer of silicon grease were used to avoid the leakage from the joints. The provisions were made on the plate to attach the combustion chamberwith a surge tank. The plate also contains four equidistant electrodes for electrical connections Figure 1
Sub-atmospheric strand burner set-up.
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Pratim Kumar et al.
Figure 2
logr vs. logp for AP/HTPB þ metal/non-metal propellants.
Figure 3
logr vs. logp for F.O. added AP/HTPB þ metal/non-metal propellants.
Figure 4
logr vs. logp for B.F. added AP/HTPB þ metal/non-metal propellants.
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vacuum levels. Two sets of readings were taken for each sample to ascertain the repeatability of the process. Reported burn rate is the average of two readings.
3. Burning rate and ignition delay data of propellant samples
Figure 5
F.O. added Al propellant at 100 kPa.
2.2.4. Chronometer Two fusible electrical wires were fixed in the propellant strand at a distance of 40 mm. These wires were connected to a timer via electrode rods. The melting of these wires will start and stop an electrical chronometer respectively. When the first fuse wire gets disconnected due to melting, the electrical chronometer starts; and when the second fuse wire disconnected due to melting, the chronometer stops. The burning rate was obtained by dividing the distance between the fuse wires to the time elapsed between the disconnection of two fuse wires. Burning rate (r) is expressed in mm/s in the entire manuscript.
2.3. Experimental procedure The propellant strand of size 6 6 60 mm3 was placed over the propellant holder. Then the ignition wires were connected to a “common” and “ignition rod”, and the fuse wires “1” and “2” were attached with “common” and “1st fuse rod” and “2nd fuse rod” respectively. These arrangements were then enclosed inside a bell jar, which constitutes the combustion chamber. Silicone grease was applied on the flange of the glass bell jar to make the enclosure leak proof. Afterwards, bell jar was secured tightly with the help of PVC washer and retainer rings, and finally electrical connections were checked properly. For creating the vacuum inside the glass bell jar, a high vacuum pump was used. It was connected to a combustion chamber via a surge tank. The level of vacuum created inside the bell jar was noted in a manometer attached to the surge tank. The set-up has the capacity to create the vacuum up to 13 kPa (100 mmHg). For igniting the propellant strands, igniter current and stop watch were switched on simultaneously. Stopping of stop watch when the propellant ignites, gives the value of ignition delay (I.D.) in seconds (s). The melt down of first fuse wire starts the timer, and melts down of second fuse wire stops the timer, and was used for obtaining the burning time. Similar procedure was applied for all the burning rate measurement under different
A total of 18 propellant samples were prepared using AP as oxidizer, HTPB as fuel binder, aluminum and boron as metallic and non-metallic fuel, and butyl ferrocene (B.F.) and ferric oxide (F.O.) as liquid and solid catalysts respectively as discussed previously. In this section, burning rate (r) and ignition delay (ID) data for each propellant sample were provided in Table 2. The discussion over burning rate, ignition delay, and flame structure are discussed separately in section 4.
4. Discussion over burning rate, ignition delay, and flame structure Composite propellant consists of heterogeneous dispersion of oxidizer particles and other solid additives such as metallic fuel, and catalyst in a polymeric fuel matrix. The combustion mechanism of propellant depends on the physico-chemical properties such as energetic of binder, oxidizer, percentage of oxidizer, oxidizer/fuel (O/F) ratio, oxidizer particle size of propellant, and on the other side, it also depends on the rocket motor conditions such as pressure, temperature, flow over the burning surface under which the propellant is burning. Attempts have been made in the present experimental work to investigate the effect of pressure on the burning rate of the prepared composite propellants made from ammonium perchlorate and HTPB polymeric fuel binder in the sub-atmospheric pressure range. In an attempt to correlate the burning rate to pressure, an equation of the type, r Z a pn, or logr Z loga þ nlogp is used for the all propellant samples [23]. In the above equation, “r” is the burning rate, “a” is the pre-exponential constant, “p” is the chamber pressure, and “n” is the pressure index. Burning rate is highly sensitive to the value of the pressure exponent; n. High values of n can produce large changes in the burning rate with relatively small changes in the chamber pressure, with potentially catastrophic consequences as higher burning rate leads to even greater chamber pressure. If n Z 1, burn rate is directly, or linearly, proportional to the chamber pressure. The slope of log of burn rate vs. pressure curve is a straight line. If the value of the exponent is close to zero, the burning rate is largely insensitive to pressure, and unstable combustion may result. For these reasons, the pressure exponent for a propellant should have a value between 0.4 and 0.8 in the regime of the motor steady state operating condition. Burning rate vs. pressure graphs for all propellants were plotted between logr vs. logp and are presented in Figures 2e4. For calculating the values of “n” and “a”,
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Pratim Kumar et al. Table 2
Burning rate and ignition delay data of propellant samples.
Pressure
100 kPa
40 kPa
27 kPa
Propellant BR/(mm/s) ID/s BR/(mm/s) ID/s BR/(mm/s) ID/s BR/(mm/s) ID/s
BR/(mm/s) ID/s
BR/(mm/s) ID/s BR/(mm/s) ID/s
Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop
0.80 0.75 0.87 0.65 0.76 0.73 1.30 1.12 1.12 1.03 1.08 1.13 1.03 1.64 1.37 1.01 1.18 1.18
0.60 0.60 0.60 0.56 0.53 0.56 0.91 0.94 0.92 0.69 0.66 0.82 0.84 1.24 1.03 0.73 0.85 0.95
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
79 kPa
2.45 1.93 2.07 1.79 2.22 2.14 2.65 2.21 2.7 1.96 2.33 2.65 2.97 3.24 3.14 2.48 3.9 3.31
1.57 1.60 4.20 2.90 3.16 2.60 1.50 1.90 3.70 6.10 1.90 2.50 1.30 2.10 2 2.08 1.6 2.3
1.82 1.42 1.83 1.22 1.60 1.30 2.35 2.04 2.35 1.77 2.06 2.10 2.45 2.92 2.74 1.87 3.30 2.68
66 kPa
4 4.20 6.10 3.60 5.65 3.65 1.90 2.10 4.20 6.80 2.40 2.70 2.60 3.30 2.50 2.45 2 3.40
1.24 1.12 1.45 1.04 1.18 0.97 2.14 1.67 1.89 1.49 1.67 1.77 1.77 2.52 2.33 1.60 2.10 1.99
53 kPa
5.80 6.10 9.20 4.05 6.43 5.22 2.10 4.05 4.60 12.4 5.30 3.20 3.40 4.40 3.00 3.95 2.00 3.80
1.07 1.05 1.14 0.80 0.93 0.91 1.84 1.44 1.49 1.26 1.37 1.40 1.37 2.03 1.91 1.32 1.30 1.48
6.20 6.50 12.30 4.75 8.48 9.40 3.50 4.75 5.10 17.3 5.50 3.80 3.60 7.70 3.30 4.06 3.40 4.40
7 7.90 16.04 5.30 8.79 10.60 5.20 5.30 5.20 21 7.20 5.40 4.80 8.20 6.50 6.03 4.70 5.40
13 kPa
7.90 8.50 22.2 6.20 9.20 13.4 5.40 5.60 6.10 29.1 7.30 6.60 5.10 13.3 10.4 6.90 8.10 6.30
0.44 0.40 0.34 0.32 0.29 0.35 0.42 0.55 0.78 0.39 0.40 CTa CT NBb NB CT 0.45 CT
8.80 9.80 28.30 7.40 13.80 17.80 8.70 6.00 9.20 38.80 9.20 8.70 5.80 NB NB 7.50 15.50 11
a
CT: Combustion terminated.
b
NB: Not burned.
Table 3
Values of “n” and “a”.
S. No.
Propellants
n
a
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop
0.834 0.75 0.915 0.789 0.967 0.805 0.923 0.691 0.649 0.813 0.903 0.970 0.989 0.745 0.869 0.898 1.061 0.873
0.043 0.053 0.031 0.040 0.023 0.040 0.042 0.095 0.12 0.049 0.038 0.036 0.029 0.10 0.060 0.038 0.026 0.048
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
logr Z loga þ nlogp equation was used, and the obtained values of “n” and “a” are presented in Table 3. For elaborate and enhanced understanding, discussion over burning rate, ignition delay, and flame structure of the propellant samples is further divided into four subsections. The subsections are: a) Combustion characteristics of AP/HTPB þ metal/nonmetal propellants; b) Combustion characteristics of ferric oxide added AP/ HTPB þ metal/non-metal propellants;
c) Combustion characteristics of butyl ferrocene added AP/ HTPB þ metal/non-metal propellants; d) Flame structure of propellants. From the next paragraph, all the four subsections are discussed sequentially.
4.1. Combustion characteristics of AP/ HTPB þ metal/non-metal propellants Figure 2 represents the logr vs. logp graph of AP/ HTPB þ metal/non-metal propellant samples. A linear variation of burning rate with pressure in the log scale was observed for all the six propellant samples, which confirms the validation of empirical relation, r Z axpn in the pressure range studied in the present investigation. However, the value of pre-exponential factor “n” and pressure index “a” are different for each propellant. The values of “n” and “a” is presented Table 3. A general trend is that with the decrease in pressure, burning rate decreases with increase in ignition delay timing. Burning rate range was also observed to be narrow for this class of propellants i.e. variation in the burning rates were less. It was also observed that addition of metal powder decreases the burning rate in all the cases. It may be due to that fact that addition of metal powder in the propellant was at the cost of AP. This results in lower percentage of AP in the propellant and hence results in reduction of oxidizing species, for the combustion of fuel binder. Moreover, because of high melting point and boiling point of the metals, these may remain un-combusted or
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partial combustion takes place during the burning of propellant at low pressure. The combustion of composite propellant is a complex process because of heterogeneity at the burning surface. It is controlled by both type of combustion mechanism i.e. premixed combustion and diffusion controlled combustion depending on the pressure at which it is burning. In the low pressure range, the premixed type of combustion dominates the overall combustion. As per Summerfield [24], chemical reaction is rate controlling step at low pressure, while diffusion is rate-controlling step at high pressure as the particle size effects tend to vanish at low pressure while they are all-important at high pressure. Considering the above facts, the decrease of burning rate with decrease of pressure is natural because the premixed flames are governed by single second-order reaction kinetics [21,24]. The gas-phase nth order chemical reactions are governed by the equation, dε=dt Z k$εn $pn1 ZA$eEa=RT $εn $pn1 This in case of second order reaction becomes: dε=dt Z A$eEa=RT $ε2 $p where, ε Z order of the chemical reaction t Z time k Z reaction constant p Z pressure Ea Z activation energy A Z Arrhenius constant R Z gas constant T Z temperature From above equation it can be observed that, dε/dt is proportional to pressure (p) in second order reaction. In other words, the rate of change of order of chemical reaction with time depends directly on the pressure. As we know, in low pressure regimes for AP based propellants premixed combustion dominates the overall combustion. Since in low pressure regimes the flame and surface temperature are low, due to which the rate of decomposition of fuel and oxidizer is also less. In low pressure, the rate controlling process is the chemical kinetics, and the fuel and oxidizer vapors have to completely diffuse together before any chemical reaction occurs. These vapors have to be mixed diffusionally in some height above the burning surface before final combustion in the flame zone. Hence, order of reaction (ε) and activation energies (Ea) of propellant ingredients governs the flame speed. These phenomenons are equated with the help of gas-phase nth order chemical reaction. Hence, in the low pressure regimes the flame speed becomes independent of pressure, and diffusion of gases, and
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decomposition of fuels and oxidizers governs the combustion process. Moreover, secondary flame which results due to the inter-diffusion of fuel vapors and combustion products of premixed flame moves away as the pressure decreases. These all affects the heat transfer from the secondary flame zone to propellant surface and finally the burning rate of the propellant samples.
4.2. Combustion characteristics of ferric oxide added AP/HTPB þ metal/non-metal propellants Figure 3 represents the logr vs. logp graph of F.O. added AP/HTPB þ metal/non-metal propellant samples. A linear variation of burning rate with pressure in the log scale was observed for all the six propellant samples. Addition of F.O. increases the burning rate of propellant samples at all the pressures. This increase of burning rate may be attributed to increased surface and sub-surface reactions occurring during the combustion and due to high flame temperature. The burning rate points at the lower end (13 kPa) move upwards, and the maximum burning rate was observed for propellant no. 12. Similarly, burning rate at higher end (100 kPa) was also observed to increase and the maximum burning rate was for propellant no. 7, 9, and 12. The burning rate range observed to be wider in the case of F.O. added propellant samples i.e. burning rate variations were higher as compared to un-catalyzed propellant samples.
4.3. Combustion characteristics of butyl ferrocene added AP/HTPB þ metal/non-metal propellants Figure 4 represents the logr vs. logp graph of B.F. added AP/HTPB þ metal/non-metal propellant samples. A linear variation of burning rate with pressure in the log scale was observed for all the six propellant samples. From Figures 3 and 4, it can be observed that the F.O. was more active in comparison to B.F. for increasing the burning rate at the lower end of the pressure range as only prop no. 17 was able to sustain its combustion at 13 kPa. The propellant containing B.F. could not sustain the combustion at pressure of 13 kPa in spite of its more uniform distribution in the propellant mix. This may be due to energy required for its activation as catalyst was not sufficient, or may be due to the lack of sufficient heat available to initiate the combustion reactions. It is quite possible that at low pressure regimes, B.F. further retards the combustion reaction resulting in extinguishment of propellants flame. During experimentations, it was found that the metal containing propellant produced more residues as compared to virgin propellants. This may be due to either partial combustion or non-combustion of metals in this pressure range. Especially in the case of aluminized propellants unburned Al particles produced with some amount of soot too.
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Pratim Kumar et al. Table 4
Order of propellants on burning rate values.
Pressure/kPa
Propellants with corresponding B.R. values in decreasing order
100 79 66 53 40 27 13
P17 (3.9) P17 (3.3) P14 (2.52) P14 (2.03) P14 (1.64) P14 (1.24) P9 (0.78)
Table 5
P18 (3.31) P14 (2.92) P15 (2.33) P15 (1.91) P15 (1.37) P15 (1.03) P8 (0.55)
P14 (3.24) P18 (2.68) P7 (2.14) P7 (1.84) P7 (1.30) P18 (0.95) P17 (0.45)
P15 (3.14) P13 (2.45) P17 (2.1) P9 (1.49) P17 Z P18 (1.18) P8 (0.94) P1 (0.44)
P13 (2.97) P7ZP9 (2.35) P18 (1.99) P18 (1.48) P12 (1.13) P9 (0.92) P7 (0.42)
Order of propellants on ignition delay values.
Pressure/kPa
Propellants with corresponding I.D. values in decreasing order
100 79 66 53 40 27 13
P10 P10 P10 P10 P10 P10 P10
(6.1) (6.8) (12.4) (17.3) (21) (29.1) (38.8)
P3 P3 P3 P3 P3 P3 P3
(4.2) (6.1) (9.2) (12.37) (16.04) (22.2) (28.3)
As the pressure decreases, copious amount of white smoke generated from boron added propellants as compared to aluminum added propellants. Table 3, presents the “n” and “a” value of, r Z axpn relation for all the propellants. It was observed that the value of “n” and “a” are propellant specific, and B.F. added propellants have higher value of “n” as compared to other propellants. In Table 4, top five performing propellants were arranged on the basis of their decreasing burning rate values in investigated pressure regimes. While in Table 5, propellants were arranged on the basis of decreasing ignition delay values. In the same table, one column is also provided for the propellant sample with least ignition delay value. This tabulation is done to enhance the understanding of the propellant combustion characteristics. The propellant with highest burning rate, lowest ignition delay, and pressure index in between 0.6 and 0.8 is the best propellant out of the 18 prepared propellant samples. From the above table, it can be observed that the propellant no. 17 (P17) which consisted of 10% boron and 2% by weight of B.F. shows the highest burning rate at 100 kPa and 79 kPa. As the pressure decreased, propellant no. 14 (P14) which consisted of 5% aluminum and 2% by weight of B.F. shows the high burn rate, and in the lowest pressure of 13 kPa, propellant no. 9 (P9) which consisted of 10% aluminum with 2% F.O. by weight possessed the highest burning rate. Another observation reveals that, 100 kPa, five of the B.F. added propellant samples have the highest burning rate as compared to other propellant samples, while
P9 P5 P5 P6 P6 P6 P6
(3.7) (5.65) (6.43) (9.4) (10.60) (13.4) (17.8)
Lowest I.D. P5 (3.16) P2ZP9 (4.2) P6 (5.22) P5 (8.48) P5 (8.79) P14 (13.3) P17 (15.5)
P13 (1.30) P7ZP17 (1.9) P7ZP17 (2) P7ZP17 (3.5) P17 (4.70) P13 (5.10) P13 (5.80)
at the lowest pressure of 13 kPa, three of the F.O. added propellant samples have the highest burning rate. In this way, it can be stated that B.F. as a catalyst was more effective in higher pressure side; while in lower pressure side F.O. was more effective in sustaining the combustion. Similarly, boron based solid propellants shows high burning rate in atmospheric pressure, while aluminum based solid propellants shows high burning rate in the lowest subatmospheric pressure regimes.
Figure 6
F.O. added Al propellant at 13 kPa.
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From Table 5, it can be observed that propellant no. 10 (P10) which consisted of 5% boron and 2% by weight of F.O. possess the highest ignition delay in all the pressure ranges. Afterwards, P3, P5, and P6 show high I.D. values. A little more observation revealed that the propellant no. 3, 5, and 6 (P3, P5, & P6) consisted of only 10% fuels i.e. 10% aluminum, 10% boron, and 5% aluminum and boron respectively. Lowest I.D. values were observed for propellant no. 13 (P13) which consisted of only 2% by weight of B.F. at 100 kPa, 27 kPa, and 13 kPa. Propellant no. 7 (2% weight of F.O.) and propellant no. 17 (10% boron and 2% by weight of B.F.) possesses same I.D. values for 79 kPa, 66 kPa, and 53 kPa. However, I.D. value of propellant no. 17 (P17) increases as the pressure decreases, which may be due to presence of high percentage of boron in it. Hence, it can be stated that B.F. added propellants ignites quiet well in vacuums, while propellants with no catalyst added were difficult to ignites in all the cases. Considering all the data’s i.e. burning rate, ignition delay, and pressure index, propellant no. 14 (P14) seems to be the most advantageous one as compared to other propellant samples. It shows comparatively high B.R. and low I.D. values in all the pressure ranges, and pressure index was also 0.745. Hence, propellant consisting of aluminum as a metallic fuel and B.F. as a catalyst can be utilized in basebleed applications of military shells and other applications.
4.4. Flame structure of propellants Boron based propellants exhibit bright green colored flame upon combustion, while aluminum based propellant combust with bright yellow colored flame. With decrease in pressure, the flame moves upwards and length get shortens and slowly changed into smoldering type combustion with generation of high amount of smoke and unburned particles. The physical mechanism which is responsible for the lifting of the flame in the low pressure is the competition between flow rate and chemical reaction, characterized by their time scales, the ratio of which is called Damkohler number. In higher pressure, chemical reaction time is faster than the flow time of decomposition products, and hence combustion takes place near to the surface; while in low pressure, chemical reaction time is slower as compared to flow time of decomposition products and hence flame moves upwards for completing the reaction [15]. Also in low pressure, since diffusion is taking faster and hence stoichiometric ratio between fuel and oxidizer vapor doesn’t formed due to which improper combustion takes place and flame temperature decreased. The flame images at 100 kPa and at 13 kPa for aluminum based CSP’s is shown in Figures 5 and 6 respectively. From Figure 5, it can be observed that the aluminum þ F.O. based CSP combusting with bright yellow color flame with streak-lines of pale yellow color of burning aluminum powders at 100 kPa. While from Figure 6, it can be observed that surface burning is taking place with AP
9
particles burning on the surface with bright yellow color patches at 13 kPa.
5. Conclusion Propellant combustion studies at sub atmospheric pressure regimes is an area of interest from the last several decades and will be of interest for combustion community scientists in the coming future too. Several physico-chemical processes continuously affect the burning rate, flame structure, flame/surface temperature, combustion products etc. All these parameters finally govern the energetic of the propellant at different pressures. In the present work, preliminary attempts are performed to understand the effects of pressure and propellant compositions on the burning rate, ignition delay, and on the flame structures of AP based CSP’s at sub-atmospheric pressure regimes. Some of the important conclusions that can be drawn from the present studies are summarized below: a) B.F. as a catalyst is more effective in higher pressure side; while in lower pressure side F.O. is more effective in sustaining the combustion. b) Boron based solid propellants shows high burning rate in atmospheric pressure, while aluminum based solid propellants shows high burning rate in the lowest subatmospheric pressure. c) B.F. added propellants ignite quiet well in vacuums, while propellants with no catalyst added were difficult to ignite in all the cases. d) The burning rate was found in the order as, B.F. þ boron > F.O. þ aluminum > Virgin propellant at 100 kPa. e) The burning rate was found in the order as, F.O. þ aluminum > B.F. þ boron > Virgin propellant at 13 kPa. f) F.O. and aluminum added propellant samples have highest ignition delay values in all the pressure regimes. g) B.F. added and B.F. þ boron added propellant samples have the lowest ignition delay values in all the pressure regimes. h) Propellant no. 14 i.e. B.F. þ aluminum added propellant samples was observed b to be the most advantageous one as compared to other propellant samples. It shows comparatively high B.R. and low I.D. values in all the pressure ranges, and pressure index was also 0.745. Hence, propellant consisting of aluminum as a metallic fuel and B.F. as a catalyst can be utilized in base-bleed applications of military shells. i) The burning rate of all AP-HTPB composite propellants increases with increase of chamber pressure, as per Vieille’s law. j) At low pressure, propellants combustion produces lots of white smoke, and in some cases soot particles and unburned particles were found on the walls of the
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combustion chamber. This indicates poor combustion efficiency at low pressure.
Acknowledgement Q4
This work was successfully completed under the guidance of Dr. P.C. Joshi, retired professor, Dpt. of Space Engineering and Rocketry, BIT Mesra, Ranchi. It’s a great pleasure to express my deep sense of unbound gratitude to all staffs of BIT Mesra.
References [1] S. Krishnan, C. Periasamy, Low pressure burning of catalyzed composite propellants, AIAA J. 24 (10) (1986) 1670e1675. [2] P.A. Ramakrishna, P.J. Paul, H.S. Mukunda, Combustion of sandwich propellant at low pressures, Proc. Combust. Inst. 30 (2005) 2097e2104. [3] P.A.O.G. Korting, H.F.R. Schoyer, Propellant combustion at low pressures, Combust. Flame 63 (1986) 317e327. [4] S. Krishnan, R.D. Swami, Effect of catalyst mixing procedure on subatmospheric combustion characteristics of composite propellants, J. Propuls. Power 13 (2) (1997) 207e212. [5] T. Kuwahara, N. Kubota, Low pressure burning of ammonium perchlorate composite propellants, Combust. Sci. Technol. 47 (1e2) (1986) 81e91. [6] L.K. Zhang, R.Y. Tian, Z.W. Zhang, Burning rate of AP/HTPB basebleed composite propellant under free ambient pressure, Aero. Sci. Technol. 62 (2017) 31e35. [7] H. Bournot, E. Daniel, R. Cayzac, Improvements of the base bleed effect using reactive particles, Int. J. Therm. Sci. 45 (2006) 1052e1065. [8] X.F. Guo, Discussion for two important parameters of a base bleed projectile, Int. J. Energetic Mater. Chem. Propuls. 1 (1e6) (1991) 159e170. [9] L.K. Zhang, Y.H. Zhou, Y.G. Yu, Research on range dispersion of base bleed projectile caused by inconsistent working characteristics of base bleed unit, Acta Armamentarii 31 (4) (2010) 442e446. [10] H.A. Abou-Elela, R. Ossama, A.Z. Ibrahim, O.K. Mahmoud, O.E. Abdel-Hamid, The optimization of base bleed grain parameters for maximum ballistic performance, in: 16th International Conference on Aerospace Sciences & Aviation Technology (ASAT), Military Technical College, Cairo, Egypt, May 2015, pp. 26e28.
[11] H.F.R. Schoyer, P.A.O.G. Korting, Low pressure combustion of composite propellant, Propellants, Explos. Pyrotech. 9 (5) (1984) 149e156. [12] J.S. Goldmeer, J.S. T’ien, D.L. Urban, Combustion and extinction of PMMA cylinders during depressurization in low-gravity, Fire Saf. J. 32 (1999) 61e88. [13] R.A. Cookson, J.B. Fenn, Strand size and low-pressure deflagration limit in a composite propellant, AIAA J. 8 (5) (1970) 864e866. [14] K.H. Parker, J. Wenograd, M. Summerfield, The ignition transient in solid propellant rocket motors, in: Solid Propellant Rocket Conference vols. 29e31, Palo Alto, , CA, U.S.A., January 1964, p. 126. [15] J.A. Steinz, P.L. Stang, M. Summerfield, The Burning Mechanism of Ammonium Perchlorate Based Composite Solid Propellants, Aerospace and Mechanical Sciences Report No. 830, Princeton University, Princeton, New Jersey, February 1969. [16] H. Belaidouni, S. Zivkovic, M. Samardzic, Numerical simulations in obtaining drag reduction for projectile with base bleed, Sci. Technol. Rev. 66 (2) (2016) 36e42. [17] T. Sofue, A. Iwana, Ignition of composite propellant at subatmospheric pressures by means of carbon dioxide laser, Propellants, Explos. Pyrotech. 4 (5) (1979) 98e106. [18] V.A. Arkhipov, M.V. Gorbenko, T.I. Gorbenko, L.A. Savel’eva, Effect of ultrafine aluminum on the combustion ofQ composite solid propellants at sub-atmospheric pressures, Combust. Explos. Shock Waves 45 (1) (2009) 40e47. [19] V.A. Arkhipov, L.A. Savel’eva, N. Zolotorev, Effect of AluminumBoron powders mechanical mixtures on the combustion of highenergy materials at sub-atmospheric pressures, MATEC Web of Conferences 23 (2015) 01005. [20] C. Zanotti, P. Giuliani, Pressure deflagration limit of solid rocket propellants: experimental results, Combust. Flame 98 (1e2) (1994) 35e45. [21] M.W. Beckstead, R.L. Derr, C.F. Price, A model of composite solid propellant combustion base on multiple flames, AIAA J. 8 (12) (1970) 2200e2207. [22] G.M. Gore, K.R. Tipare, R.G. Bhatewara, U.S. Prasad, M. Gupta, S.K. Mane, Evaluation of ferrocene derivatives as burn rate modifiers in AP-HTPB based composite propellants, Def. Sci. J. 49 (2) (1999) 151e158. [23] P. Kumar, P.C. Joshi, R. Kumar, Thermal decomposition and combustion studies of AN/KDN based solid propellants, Combust. Flame 166 (2016) 316e332. [24] M. Summerfield, G.S. Sutherland, M.J. Webb, H.J. Taback, K.P. Hall, Burning mechanism of ammonium perchlorate propellants, in: M. Summerfield (Ed.), Solid Propellant Rocket Research, Academic Press, 1960, pp. 141e182.
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Propulsion and Power Research 2019;xxx(xxx):1e10 http://ppr.buaa.edu.cn/
Propulsion and Power Research w w w. s c i e n c e d i r e c t . c om
ORGINAL ARTICLE
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Q5 Q2
Combustion performance studies of aluminum and boron based composite solid propellants in sub-atmospheric pressure regimes Pratim Kumara,*, Mayank Varshneyb, Aniket Manashc a
Department of Aerospace Engineering and Applied Mechanics, IIEST, Shibpur, Howrah, West Bengal, 711103, India b Indian Institute of Technology (Indian School of Mines), Dhanbad, Jharkhand, 862004, India c Aryabhatta Center for Nanoscience & Technology, AKU, Patna, Bihar, 800001, India Received 30 October 2018; accepted 24 September 2019 Available online XXXX
KEYWORDS AP/HTPB; Metallic/non-metallic fuel; Catalyst; Sub-atmospheric pressure; Burning rate; Base-bleed
Abstract The aim of present study is to investigate the burning rate, ignition delay, and flame characteristics of ammonium perchlorate (AP)-hydroxyl terminated poly-butadiene (HTPB) [AP/HTPB] based composite propellants (CSP’s) in sub-atmospheric pressure regimes (13 kPae100 kPa). Several fuels and catalyzed were used to evaluate their effects on the combustion characteristics of AP based propellants in sub-atmospheric pressure regimes. In fuels, aluminum (Al) and boron (B) were selected as metallic and non-metallic fuel respectively. While in catalyst, butyl ferrocene (B.F.) and ferric oxide (F.O.) were selected as liquid and solid catalyst respectively. Apart from these, other ingredients that were used are di-octyl adipate (DOA), toluene di-isocyanate (TDI), and glycerol. The article throws some light on the burning rate and ignition delay properties for these new classes of prepared propellant samples. At subatmospheric pressures, all propellants are susceptible to irregular burning with the ejection of soot’s, fumes, and unburned particles. F.O. based catalyzed propellants can sustain its combustion up to the lowest pressure. ª 2019 Beihang University. Production and hosting by Elsevier B.V. on behalf of KeAi. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/licenses/by-nc-nd/4.0/).
*Corresponding author. E-mail address: [email protected] (Pratim Kumar). Peer review under responsibility of Beihang University.
Production and Hosting by Elsevier on behalf of KeAi
1. Introduction From last several decades there is a remarkable increase of interest in the field of sub-atmospheric pressure combustion studies of composite solid propellants [1e5]. The
https://doi.org/10.1016/j.jppr.2019.09.001 2212-540X/ª 2019 Beihang University. Production and hosting by Elsevier B.V. on behalf of KeAi. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/licenses/by-nc-nd/4.0/).
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main application of sub-atmospheric pressure combustion studies are in the field of base bleed applications of missiles and military shells [6e10], propellant flame structures [11,12], low pressure deflagration limit (LPDL) [13], accurate prediction of early stages of motor ignition transients [14], and start stop of solid rocket motors [14]. Subatmospheric pressure conditions also allows better investigations over the combustion mechanism and flame structure since, (a) gas phase reaction zone consisting mainly of premixed flame (as against competing premixed and diffusion flames at rocket operating pressures), (b) less severe temperature gradient at the burning surface, and (c) better spatial and temporal resolutions. At sub-atmospheric pressure regimes, the diffusion processes became much faster and chemical reactions are slower at lower pressures [15]. It is common for the time and distance scale of flame zone to increase with decreasing pressure. Combustion studies in sub-atmospheric pressure are of interest for flame zone structures in cognizance that controlling chemical kinetics reaction steps and product gas compositions changes with pressure. Solid propellants which have the ability to burn in vacuum at high burning rate can act as a gas generator for reducing the partial vacuum created at the projectile base during flight [16]. The combustion gases from the combustion chamber escape through an unchoked port in the projectile base. Due to this, drag reduction and consequently range enhancement of military shells takes place. This process is known as base-bleed. For effective base bleed (BB) applications of missiles and shells, the propellant must have to sustain its combustion and also to burn more efficiently at sub-atmospheric pressure regimes. To initiate the proper ignition and to increase the burn rate of solid propellants in sub-atmospheric pressure [17e19] is an area of exciting research and have promising future prospects. Although composite solid propellants, as compared to double base (DB) propellants, burn more efficiently at low pressures, the combustion efficiently falls drastically below certain sub-atmospheric pressures culminating in extinction. The pressure at which extinction occurs is termed as low pressure deflagration limit (LPDL), as against the high pressure deflagration limit (HPDL) demonstrated in certain propellants [20]. Excessive fuming, soot generation, and ejection of unburned particles are some of the common characteristics of propellant combustion at sub-atmospheric pressure regimes. These burning characteristics show the poor combustion efficiency at sub-atmospheric pressure regimes. At some pressure, propellant can’t sustain its combustion and this limiting pressure is known as LPDL [21]. As per the Beckstead-Derr-Price (BDP) model, the primary diffusion flame controlled the combustion at the low pressure end, while the final diffusion flame along with the AP decomposition flame controlled the combustion at the high pressure end. In the present paper, an attempt has been made to investigate the burning rate, ignition delay, and flame structure of 18 ammonium perchlorate (AP, NH4ClO4)
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based composite solid propellants in the pressure regimes of 100 kPa (760 mmHg) to 13 kPa (100 mmHg). Two ferrous based catalysts were used in the study viz. ferric oxide (F.O.) as a solid catalyst, and butyl ferrocene (B.F.) as a liquid catalyst [22]. Two fuels were also used viz. aluminium (Al) as a metallic fuel, and boron (B) as a nonmetallic fuel. Present study will be useful for obtaining a better composition of AP based propellant for increasing the range of military shells by decreasing their base drag using a gas-generator unit which can sustain its combustion in subatmospheric pressure regimes.
2. Experimental methodology 2.1. Composition, formulation, and preparation of propellant samples As discussed previously, a solid propellant basically consists of oxidizer, fuel, catalyst, and other additives. The selected components for propellant processing are given below: 1. Fuel binder: Hydroxyl-terminated poly-butadiene (HTPB) (supplied by propellant fuel complex, Vikram Sarabhai Space Centre (VSSC), Thiruvananthapuram). Currently HTPB is being extensively used as a polymeric fuel binder because of its capacity to load higher solid filler concentrations, and superior performance. 2. Oxidizer: Ammonium perchlorate (AP) (supplied by Tamil Nadu Chemicals Ltd., Madurai). A unimodal particle size of 74 mm was used in the study. 3. Plasticizer: Dioctyl adipate (DOA) (supplied by Chemporium Co., Mumbai). DOA was used as plasticizer in the present work in order to maintain easy process ability and good mechanical properties of the propellant. 4. Curing agent: Toluene di-isocyanate (TDI) (supplied by Merck Co., Mumbai). It helps in curing of HTPB. 5. Cross linking agent: Glycerol (supplied by Merck Co., Mumbai). It helps in the formation of matrix type structure between fuel and oxidizer particle. 6. Metallic fuel: Aluminum powder (fine, <75 mm) (supplied by: Sigma Aldrich). 7. Non-metallic fuel: Boron powder (amorphous powder, supplied by Thomas Baker, Mumbai). 8. Burning rate modifiers: Ferric oxide [5 mm (solid state), supplied by Central Drug House (CDH)], and butyl Q3 ferrocene (liquid state) supplied by Sigma Aldrich. A total of 18 propellant compositions were formulated using above ingredients in different proportions. Oxidizer percentage varied by 80, 75, and 70, while fuel percentage varied by 20, 25, and 30 in the propellants. Catalyst percentage was kept 2% to the total weight of the propellant sample. Composition of the prepared propellant samples is presented in Table 1. For preparation of propellant samples, requisite quantities HTPB, DOA, and glycerol were mixed thoroughly
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Table 1
Composition of prepared propellants.
Propellant No.
AP/%
HTPB/%
DOA/%
TDI/%
GLYCEROL/%
Al/%
B/%
F.O./wt%
B.F./wt%
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
80 75 70 75 70 70 80 75 70 75 70 70 80 75 70 75 70 70
14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36 14.36
4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36 4.36
1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13 1.13
0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15
0 5 10 0 0 5 0 5 10 0 0 5 0 5 10 0 0 5
0 0 0 5 10 5 0 0 0 5 10 5 0 0 0 5 10 5
0 0 0 0 0 0 2 2 2 2 2 2 0 0 0 0 0 0
0 0 0 0 0 0 0 0 0 0 0 0 2 2 2 2 2 2
inside a mixer till the whole mixture becomes homogeneous. Afterwards, requisite quantities of fuels and catalyst were added in the mixture with continuous mixing to maintain homogeneity. At last, TDI was added and mixing continued for another 30 min. The final slurry was casted in a stainless steel casting box. In the casting box, 12 slots each of size 60 6 6 mm3 were manufactured for obtaining the casted propellant strands. The propellant slurry filled casting box was then kept in a vacuum electric oven at 60 C 2 C for a period of six days to allow it to cure into a polymeric state and to attain required mechanical strength. After six days, the moulds were removed from the oven and allowed to cool at room temperature. The propellant blocks thus obtained were kept in a polythene bag and stored inside desiccators to avoid moisture absorption. Finally, the obtained propellant strands were inhibited from all four sides by quick drying polyethene enamel for obtaining linear burning.
needed for propellant ignition, and for timer reading (shown in Figure 5).
2.2.2. De-pressurization system It consists of a high vacuum pump which was connected at the top of the surge tank through a high vacuum line. At the top of the surge tank, a mercury manometer was attached to read the vacuum level inside the combustion chamber. The advantage of the surge tank is to minimize the pressure rise and fluctuations during the combustion of propellant strands.
2.2.3. Ignition system Hot wire ignition system was employed for igniting the propellant strand. Nichrome wire with resistance of 20 ohm per cm and of diameter 0.27 mm was passed at the top of the propellant strand, and was heated till the propellant sample get ignited by passing a current of 1.5 amp at 30 V.
2.2. Experimental set-up The experimental setup is also known as sub-atmospheric strand burner set up. The detail of experimental set-up is shown in Figure 1. The details of different parts of the experimental set up are discussed one by one.
2.2.1. Combustion chamber A cylindrical bell jar of borosilicate glass of capacity 2000 cc was used as a combustion chamber. It was fixed on a stainless steel plate with the help of retainer rings, nuts, and bolts. Polyvinyl chloride (PVC) washer and a thin layer of silicon grease were used to avoid the leakage from the joints. The provisions were made on the plate to attach the combustion chamberwith a surge tank. The plate also contains four equidistant electrodes for electrical connections Figure 1
Sub-atmospheric strand burner set-up.
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Figure 2
logr vs. logp for AP/HTPB þ metal/non-metal propellants.
Figure 3
logr vs. logp for F.O. added AP/HTPB þ metal/non-metal propellants.
Figure 4
logr vs. logp for B.F. added AP/HTPB þ metal/non-metal propellants.
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vacuum levels. Two sets of readings were taken for each sample to ascertain the repeatability of the process. Reported burn rate is the average of two readings.
3. Burning rate and ignition delay data of propellant samples
Figure 5
F.O. added Al propellant at 100 kPa.
2.2.4. Chronometer Two fusible electrical wires were fixed in the propellant strand at a distance of 40 mm. These wires were connected to a timer via electrode rods. The melting of these wires will start and stop an electrical chronometer respectively. When the first fuse wire gets disconnected due to melting, the electrical chronometer starts; and when the second fuse wire disconnected due to melting, the chronometer stops. The burning rate was obtained by dividing the distance between the fuse wires to the time elapsed between the disconnection of two fuse wires. Burning rate (r) is expressed in mm/s in the entire manuscript.
2.3. Experimental procedure The propellant strand of size 6 6 60 mm3 was placed over the propellant holder. Then the ignition wires were connected to a “common” and “ignition rod”, and the fuse wires “1” and “2” were attached with “common” and “1st fuse rod” and “2nd fuse rod” respectively. These arrangements were then enclosed inside a bell jar, which constitutes the combustion chamber. Silicone grease was applied on the flange of the glass bell jar to make the enclosure leak proof. Afterwards, bell jar was secured tightly with the help of PVC washer and retainer rings, and finally electrical connections were checked properly. For creating the vacuum inside the glass bell jar, a high vacuum pump was used. It was connected to a combustion chamber via a surge tank. The level of vacuum created inside the bell jar was noted in a manometer attached to the surge tank. The set-up has the capacity to create the vacuum up to 13 kPa (100 mmHg). For igniting the propellant strands, igniter current and stop watch were switched on simultaneously. Stopping of stop watch when the propellant ignites, gives the value of ignition delay (I.D.) in seconds (s). The melt down of first fuse wire starts the timer, and melts down of second fuse wire stops the timer, and was used for obtaining the burning time. Similar procedure was applied for all the burning rate measurement under different
A total of 18 propellant samples were prepared using AP as oxidizer, HTPB as fuel binder, aluminum and boron as metallic and non-metallic fuel, and butyl ferrocene (B.F.) and ferric oxide (F.O.) as liquid and solid catalysts respectively as discussed previously. In this section, burning rate (r) and ignition delay (ID) data for each propellant sample were provided in Table 2. The discussion over burning rate, ignition delay, and flame structure are discussed separately in section 4.
4. Discussion over burning rate, ignition delay, and flame structure Composite propellant consists of heterogeneous dispersion of oxidizer particles and other solid additives such as metallic fuel, and catalyst in a polymeric fuel matrix. The combustion mechanism of propellant depends on the physico-chemical properties such as energetic of binder, oxidizer, percentage of oxidizer, oxidizer/fuel (O/F) ratio, oxidizer particle size of propellant, and on the other side, it also depends on the rocket motor conditions such as pressure, temperature, flow over the burning surface under which the propellant is burning. Attempts have been made in the present experimental work to investigate the effect of pressure on the burning rate of the prepared composite propellants made from ammonium perchlorate and HTPB polymeric fuel binder in the sub-atmospheric pressure range. In an attempt to correlate the burning rate to pressure, an equation of the type, r Z a pn, or logr Z loga þ nlogp is used for the all propellant samples [23]. In the above equation, “r” is the burning rate, “a” is the pre-exponential constant, “p” is the chamber pressure, and “n” is the pressure index. Burning rate is highly sensitive to the value of the pressure exponent; n. High values of n can produce large changes in the burning rate with relatively small changes in the chamber pressure, with potentially catastrophic consequences as higher burning rate leads to even greater chamber pressure. If n Z 1, burn rate is directly, or linearly, proportional to the chamber pressure. The slope of log of burn rate vs. pressure curve is a straight line. If the value of the exponent is close to zero, the burning rate is largely insensitive to pressure, and unstable combustion may result. For these reasons, the pressure exponent for a propellant should have a value between 0.4 and 0.8 in the regime of the motor steady state operating condition. Burning rate vs. pressure graphs for all propellants were plotted between logr vs. logp and are presented in Figures 2e4. For calculating the values of “n” and “a”,
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Pratim Kumar et al. Table 2
Burning rate and ignition delay data of propellant samples.
Pressure
100 kPa
40 kPa
27 kPa
Propellant BR/(mm/s) ID/s BR/(mm/s) ID/s BR/(mm/s) ID/s BR/(mm/s) ID/s
BR/(mm/s) ID/s
BR/(mm/s) ID/s BR/(mm/s) ID/s
Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop
0.80 0.75 0.87 0.65 0.76 0.73 1.30 1.12 1.12 1.03 1.08 1.13 1.03 1.64 1.37 1.01 1.18 1.18
0.60 0.60 0.60 0.56 0.53 0.56 0.91 0.94 0.92 0.69 0.66 0.82 0.84 1.24 1.03 0.73 0.85 0.95
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
79 kPa
2.45 1.93 2.07 1.79 2.22 2.14 2.65 2.21 2.7 1.96 2.33 2.65 2.97 3.24 3.14 2.48 3.9 3.31
1.57 1.60 4.20 2.90 3.16 2.60 1.50 1.90 3.70 6.10 1.90 2.50 1.30 2.10 2 2.08 1.6 2.3
1.82 1.42 1.83 1.22 1.60 1.30 2.35 2.04 2.35 1.77 2.06 2.10 2.45 2.92 2.74 1.87 3.30 2.68
66 kPa
4 4.20 6.10 3.60 5.65 3.65 1.90 2.10 4.20 6.80 2.40 2.70 2.60 3.30 2.50 2.45 2 3.40
1.24 1.12 1.45 1.04 1.18 0.97 2.14 1.67 1.89 1.49 1.67 1.77 1.77 2.52 2.33 1.60 2.10 1.99
53 kPa
5.80 6.10 9.20 4.05 6.43 5.22 2.10 4.05 4.60 12.4 5.30 3.20 3.40 4.40 3.00 3.95 2.00 3.80
1.07 1.05 1.14 0.80 0.93 0.91 1.84 1.44 1.49 1.26 1.37 1.40 1.37 2.03 1.91 1.32 1.30 1.48
6.20 6.50 12.30 4.75 8.48 9.40 3.50 4.75 5.10 17.3 5.50 3.80 3.60 7.70 3.30 4.06 3.40 4.40
7 7.90 16.04 5.30 8.79 10.60 5.20 5.30 5.20 21 7.20 5.40 4.80 8.20 6.50 6.03 4.70 5.40
13 kPa
7.90 8.50 22.2 6.20 9.20 13.4 5.40 5.60 6.10 29.1 7.30 6.60 5.10 13.3 10.4 6.90 8.10 6.30
0.44 0.40 0.34 0.32 0.29 0.35 0.42 0.55 0.78 0.39 0.40 CTa CT NBb NB CT 0.45 CT
8.80 9.80 28.30 7.40 13.80 17.80 8.70 6.00 9.20 38.80 9.20 8.70 5.80 NB NB 7.50 15.50 11
a
CT: Combustion terminated.
b
NB: Not burned.
Table 3
Values of “n” and “a”.
S. No.
Propellants
n
a
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop Prop
0.834 0.75 0.915 0.789 0.967 0.805 0.923 0.691 0.649 0.813 0.903 0.970 0.989 0.745 0.869 0.898 1.061 0.873
0.043 0.053 0.031 0.040 0.023 0.040 0.042 0.095 0.12 0.049 0.038 0.036 0.029 0.10 0.060 0.038 0.026 0.048
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18
logr Z loga þ nlogp equation was used, and the obtained values of “n” and “a” are presented in Table 3. For elaborate and enhanced understanding, discussion over burning rate, ignition delay, and flame structure of the propellant samples is further divided into four subsections. The subsections are: a) Combustion characteristics of AP/HTPB þ metal/nonmetal propellants; b) Combustion characteristics of ferric oxide added AP/ HTPB þ metal/non-metal propellants;
c) Combustion characteristics of butyl ferrocene added AP/ HTPB þ metal/non-metal propellants; d) Flame structure of propellants. From the next paragraph, all the four subsections are discussed sequentially.
4.1. Combustion characteristics of AP/ HTPB þ metal/non-metal propellants Figure 2 represents the logr vs. logp graph of AP/ HTPB þ metal/non-metal propellant samples. A linear variation of burning rate with pressure in the log scale was observed for all the six propellant samples, which confirms the validation of empirical relation, r Z axpn in the pressure range studied in the present investigation. However, the value of pre-exponential factor “n” and pressure index “a” are different for each propellant. The values of “n” and “a” is presented Table 3. A general trend is that with the decrease in pressure, burning rate decreases with increase in ignition delay timing. Burning rate range was also observed to be narrow for this class of propellants i.e. variation in the burning rates were less. It was also observed that addition of metal powder decreases the burning rate in all the cases. It may be due to that fact that addition of metal powder in the propellant was at the cost of AP. This results in lower percentage of AP in the propellant and hence results in reduction of oxidizing species, for the combustion of fuel binder. Moreover, because of high melting point and boiling point of the metals, these may remain un-combusted or
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partial combustion takes place during the burning of propellant at low pressure. The combustion of composite propellant is a complex process because of heterogeneity at the burning surface. It is controlled by both type of combustion mechanism i.e. premixed combustion and diffusion controlled combustion depending on the pressure at which it is burning. In the low pressure range, the premixed type of combustion dominates the overall combustion. As per Summerfield [24], chemical reaction is rate controlling step at low pressure, while diffusion is rate-controlling step at high pressure as the particle size effects tend to vanish at low pressure while they are all-important at high pressure. Considering the above facts, the decrease of burning rate with decrease of pressure is natural because the premixed flames are governed by single second-order reaction kinetics [21,24]. The gas-phase nth order chemical reactions are governed by the equation, dε=dt Z k$εn $pn1 ZA$eEa=RT $εn $pn1 This in case of second order reaction becomes: dε=dt Z A$eEa=RT $ε2 $p where, ε Z order of the chemical reaction t Z time k Z reaction constant p Z pressure Ea Z activation energy A Z Arrhenius constant R Z gas constant T Z temperature From above equation it can be observed that, dε/dt is proportional to pressure (p) in second order reaction. In other words, the rate of change of order of chemical reaction with time depends directly on the pressure. As we know, in low pressure regimes for AP based propellants premixed combustion dominates the overall combustion. Since in low pressure regimes the flame and surface temperature are low, due to which the rate of decomposition of fuel and oxidizer is also less. In low pressure, the rate controlling process is the chemical kinetics, and the fuel and oxidizer vapors have to completely diffuse together before any chemical reaction occurs. These vapors have to be mixed diffusionally in some height above the burning surface before final combustion in the flame zone. Hence, order of reaction (ε) and activation energies (Ea) of propellant ingredients governs the flame speed. These phenomenons are equated with the help of gas-phase nth order chemical reaction. Hence, in the low pressure regimes the flame speed becomes independent of pressure, and diffusion of gases, and
7
decomposition of fuels and oxidizers governs the combustion process. Moreover, secondary flame which results due to the inter-diffusion of fuel vapors and combustion products of premixed flame moves away as the pressure decreases. These all affects the heat transfer from the secondary flame zone to propellant surface and finally the burning rate of the propellant samples.
4.2. Combustion characteristics of ferric oxide added AP/HTPB þ metal/non-metal propellants Figure 3 represents the logr vs. logp graph of F.O. added AP/HTPB þ metal/non-metal propellant samples. A linear variation of burning rate with pressure in the log scale was observed for all the six propellant samples. Addition of F.O. increases the burning rate of propellant samples at all the pressures. This increase of burning rate may be attributed to increased surface and sub-surface reactions occurring during the combustion and due to high flame temperature. The burning rate points at the lower end (13 kPa) move upwards, and the maximum burning rate was observed for propellant no. 12. Similarly, burning rate at higher end (100 kPa) was also observed to increase and the maximum burning rate was for propellant no. 7, 9, and 12. The burning rate range observed to be wider in the case of F.O. added propellant samples i.e. burning rate variations were higher as compared to un-catalyzed propellant samples.
4.3. Combustion characteristics of butyl ferrocene added AP/HTPB þ metal/non-metal propellants Figure 4 represents the logr vs. logp graph of B.F. added AP/HTPB þ metal/non-metal propellant samples. A linear variation of burning rate with pressure in the log scale was observed for all the six propellant samples. From Figures 3 and 4, it can be observed that the F.O. was more active in comparison to B.F. for increasing the burning rate at the lower end of the pressure range as only prop no. 17 was able to sustain its combustion at 13 kPa. The propellant containing B.F. could not sustain the combustion at pressure of 13 kPa in spite of its more uniform distribution in the propellant mix. This may be due to energy required for its activation as catalyst was not sufficient, or may be due to the lack of sufficient heat available to initiate the combustion reactions. It is quite possible that at low pressure regimes, B.F. further retards the combustion reaction resulting in extinguishment of propellants flame. During experimentations, it was found that the metal containing propellant produced more residues as compared to virgin propellants. This may be due to either partial combustion or non-combustion of metals in this pressure range. Especially in the case of aluminized propellants unburned Al particles produced with some amount of soot too.
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Pratim Kumar et al. Table 4
Order of propellants on burning rate values.
Pressure/kPa
Propellants with corresponding B.R. values in decreasing order
100 79 66 53 40 27 13
P17 (3.9) P17 (3.3) P14 (2.52) P14 (2.03) P14 (1.64) P14 (1.24) P9 (0.78)
Table 5
P18 (3.31) P14 (2.92) P15 (2.33) P15 (1.91) P15 (1.37) P15 (1.03) P8 (0.55)
P14 (3.24) P18 (2.68) P7 (2.14) P7 (1.84) P7 (1.30) P18 (0.95) P17 (0.45)
P15 (3.14) P13 (2.45) P17 (2.1) P9 (1.49) P17 Z P18 (1.18) P8 (0.94) P1 (0.44)
P13 (2.97) P7ZP9 (2.35) P18 (1.99) P18 (1.48) P12 (1.13) P9 (0.92) P7 (0.42)
Order of propellants on ignition delay values.
Pressure/kPa
Propellants with corresponding I.D. values in decreasing order
100 79 66 53 40 27 13
P10 P10 P10 P10 P10 P10 P10
(6.1) (6.8) (12.4) (17.3) (21) (29.1) (38.8)
P3 P3 P3 P3 P3 P3 P3
(4.2) (6.1) (9.2) (12.37) (16.04) (22.2) (28.3)
As the pressure decreases, copious amount of white smoke generated from boron added propellants as compared to aluminum added propellants. Table 3, presents the “n” and “a” value of, r Z axpn relation for all the propellants. It was observed that the value of “n” and “a” are propellant specific, and B.F. added propellants have higher value of “n” as compared to other propellants. In Table 4, top five performing propellants were arranged on the basis of their decreasing burning rate values in investigated pressure regimes. While in Table 5, propellants were arranged on the basis of decreasing ignition delay values. In the same table, one column is also provided for the propellant sample with least ignition delay value. This tabulation is done to enhance the understanding of the propellant combustion characteristics. The propellant with highest burning rate, lowest ignition delay, and pressure index in between 0.6 and 0.8 is the best propellant out of the 18 prepared propellant samples. From the above table, it can be observed that the propellant no. 17 (P17) which consisted of 10% boron and 2% by weight of B.F. shows the highest burning rate at 100 kPa and 79 kPa. As the pressure decreased, propellant no. 14 (P14) which consisted of 5% aluminum and 2% by weight of B.F. shows the high burn rate, and in the lowest pressure of 13 kPa, propellant no. 9 (P9) which consisted of 10% aluminum with 2% F.O. by weight possessed the highest burning rate. Another observation reveals that, 100 kPa, five of the B.F. added propellant samples have the highest burning rate as compared to other propellant samples, while
P9 P5 P5 P6 P6 P6 P6
(3.7) (5.65) (6.43) (9.4) (10.60) (13.4) (17.8)
Lowest I.D. P5 (3.16) P2ZP9 (4.2) P6 (5.22) P5 (8.48) P5 (8.79) P14 (13.3) P17 (15.5)
P13 (1.30) P7ZP17 (1.9) P7ZP17 (2) P7ZP17 (3.5) P17 (4.70) P13 (5.10) P13 (5.80)
at the lowest pressure of 13 kPa, three of the F.O. added propellant samples have the highest burning rate. In this way, it can be stated that B.F. as a catalyst was more effective in higher pressure side; while in lower pressure side F.O. was more effective in sustaining the combustion. Similarly, boron based solid propellants shows high burning rate in atmospheric pressure, while aluminum based solid propellants shows high burning rate in the lowest subatmospheric pressure regimes.
Figure 6
F.O. added Al propellant at 13 kPa.
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From Table 5, it can be observed that propellant no. 10 (P10) which consisted of 5% boron and 2% by weight of F.O. possess the highest ignition delay in all the pressure ranges. Afterwards, P3, P5, and P6 show high I.D. values. A little more observation revealed that the propellant no. 3, 5, and 6 (P3, P5, & P6) consisted of only 10% fuels i.e. 10% aluminum, 10% boron, and 5% aluminum and boron respectively. Lowest I.D. values were observed for propellant no. 13 (P13) which consisted of only 2% by weight of B.F. at 100 kPa, 27 kPa, and 13 kPa. Propellant no. 7 (2% weight of F.O.) and propellant no. 17 (10% boron and 2% by weight of B.F.) possesses same I.D. values for 79 kPa, 66 kPa, and 53 kPa. However, I.D. value of propellant no. 17 (P17) increases as the pressure decreases, which may be due to presence of high percentage of boron in it. Hence, it can be stated that B.F. added propellants ignites quiet well in vacuums, while propellants with no catalyst added were difficult to ignites in all the cases. Considering all the data’s i.e. burning rate, ignition delay, and pressure index, propellant no. 14 (P14) seems to be the most advantageous one as compared to other propellant samples. It shows comparatively high B.R. and low I.D. values in all the pressure ranges, and pressure index was also 0.745. Hence, propellant consisting of aluminum as a metallic fuel and B.F. as a catalyst can be utilized in basebleed applications of military shells and other applications.
4.4. Flame structure of propellants Boron based propellants exhibit bright green colored flame upon combustion, while aluminum based propellant combust with bright yellow colored flame. With decrease in pressure, the flame moves upwards and length get shortens and slowly changed into smoldering type combustion with generation of high amount of smoke and unburned particles. The physical mechanism which is responsible for the lifting of the flame in the low pressure is the competition between flow rate and chemical reaction, characterized by their time scales, the ratio of which is called Damkohler number. In higher pressure, chemical reaction time is faster than the flow time of decomposition products, and hence combustion takes place near to the surface; while in low pressure, chemical reaction time is slower as compared to flow time of decomposition products and hence flame moves upwards for completing the reaction [15]. Also in low pressure, since diffusion is taking faster and hence stoichiometric ratio between fuel and oxidizer vapor doesn’t formed due to which improper combustion takes place and flame temperature decreased. The flame images at 100 kPa and at 13 kPa for aluminum based CSP’s is shown in Figures 5 and 6 respectively. From Figure 5, it can be observed that the aluminum þ F.O. based CSP combusting with bright yellow color flame with streak-lines of pale yellow color of burning aluminum powders at 100 kPa. While from Figure 6, it can be observed that surface burning is taking place with AP
9
particles burning on the surface with bright yellow color patches at 13 kPa.
5. Conclusion Propellant combustion studies at sub atmospheric pressure regimes is an area of interest from the last several decades and will be of interest for combustion community scientists in the coming future too. Several physico-chemical processes continuously affect the burning rate, flame structure, flame/surface temperature, combustion products etc. All these parameters finally govern the energetic of the propellant at different pressures. In the present work, preliminary attempts are performed to understand the effects of pressure and propellant compositions on the burning rate, ignition delay, and on the flame structures of AP based CSP’s at sub-atmospheric pressure regimes. Some of the important conclusions that can be drawn from the present studies are summarized below: a) B.F. as a catalyst is more effective in higher pressure side; while in lower pressure side F.O. is more effective in sustaining the combustion. b) Boron based solid propellants shows high burning rate in atmospheric pressure, while aluminum based solid propellants shows high burning rate in the lowest subatmospheric pressure. c) B.F. added propellants ignite quiet well in vacuums, while propellants with no catalyst added were difficult to ignite in all the cases. d) The burning rate was found in the order as, B.F. þ boron > F.O. þ aluminum > Virgin propellant at 100 kPa. e) The burning rate was found in the order as, F.O. þ aluminum > B.F. þ boron > Virgin propellant at 13 kPa. f) F.O. and aluminum added propellant samples have highest ignition delay values in all the pressure regimes. g) B.F. added and B.F. þ boron added propellant samples have the lowest ignition delay values in all the pressure regimes. h) Propellant no. 14 i.e. B.F. þ aluminum added propellant samples was observed b to be the most advantageous one as compared to other propellant samples. It shows comparatively high B.R. and low I.D. values in all the pressure ranges, and pressure index was also 0.745. Hence, propellant consisting of aluminum as a metallic fuel and B.F. as a catalyst can be utilized in base-bleed applications of military shells. i) The burning rate of all AP-HTPB composite propellants increases with increase of chamber pressure, as per Vieille’s law. j) At low pressure, propellants combustion produces lots of white smoke, and in some cases soot particles and unburned particles were found on the walls of the
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combustion chamber. This indicates poor combustion efficiency at low pressure.
Acknowledgement Q4
This work was successfully completed under the guidance of Dr. P.C. Joshi, retired professor, Dpt. of Space Engineering and Rocketry, BIT Mesra, Ranchi. It’s a great pleasure to express my deep sense of unbound gratitude to all staffs of BIT Mesra.
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