CHAPTER 1 INTRODUCTION 1
CHAPTER 1
INTRODUCTION
1.1 INTRODUCTION OF THRUSTERS USED IN SPACECRAFT
A thruster is a propulsive device used by spacecraft for station keeping, attitude control, in the reaction control system, or long-duration, low-thrust acceleration. All thrusters work by exhausting a propellant in one direction so that the spacecraft goes in the other. They can provide either positioning or, by having two thrusters that point in opposite directions on either side of the spacecraft, rotate it in place.
Basically there are four kinds, depending on the mass of the spacecraft.
Bi-propellant thrusters: (e.g. liquid-oxygen/kerosene or (Di-nitrogen tetroxide plus hydrazine) (Hypergolic propellant))are used for main thrusters.
Figure 1: Bi-propellant thruster
These provide a lot of power but are difficult to turn on and off quickly.
Monopropellant thrusters, such as use hydrazine or hydrogen-peroxide, ignited by a catalyst, can be either main thrusters or attitude control thrusters
Figure 2: Monopropellant thruster
These provide a moderate amount of thrust and can be turned on-and-off quickly.
Cold-gas thrusters: which consist simply of a tank of gas, such as nitrogen, a nozzle, and a valve
Figure 3: Cold gas thruster
Exotic electrical thrusters: like ion drives Ion Propulsion and Pulsed Plasma Thrusters .
These don’t provide much thrust, but are extremely efficient because their energy isn’t stored in the fuel but comes from the main electrical source, usually solar panels. These type of thrusters are very safe and compact.
Ion drives are used to adjust geosynchronous satellites in their orbits and for long-duration space probes.
Pulsed Plasma thrusters are finding a niche in providing attitude control for micro-satellites, since they can be made very small and light-weight.
For all these the mono propellant thrusters are widely using in the space craft upper stage propulsion system and also as the attitude control thrusters for their reliable performance for long duration as well as multiple start-stop-restart capabilities.
1.2 Current state of the art for Monopropellants:
Hydrazine (N2H4) is by far the most commonly used propellant for primary spacecraft propulsion and attitude control thrusters. Hydrazine thrusters, which typically consist of an electric solenoid valve, a pressurant tank, and a catalyst bed of alumina pellets impregnated with iridium, are feature a simple design architecture, are highly reliable, and offer reasonable performance. In a typical design the catalyst initiates an exothermic decomposition of the hydrazine to produce ammonia, nitrogen, and hydrogen gases with approximately 1600 kJ/kg of heat. Although hydrazine decomposition using the Shell 405 catalyst can be performed with no additional heat input to the catalyst, typical designs pre-heat the catalyst bed to insure more reliable ignition and consistent burn profile. In preheated configurations vacuum specific impulse (Isp) greater than 220 seconds can be achieved.
Unfortunately, hydrazine is a powerful reducing agent that poses serious environmental concerns. Hydrazine is extremely destructive to living tissues, and is a probable human carcinogen. Exposure produces a variety of adverse systemic effects including damage to liver, kidneys, nervous system, and red blood cells. In addition to these biological and toxicological impacts, hydrazine presents significant environmental dangers for the spacecraft and launch vehicle. As a chemical prone to rapidly decompose or explode when struck, vibrated, or otherwise agitated, hydrazine is listed among the most sensitive shock-sensitive chemicals by the US Department of transportation. Most significantly, hydrazine has a moderately high vapor pressure at room temperature, approximately 1000 kPa (145 psia). For many applications a high storage vapor pressure is a good propellant feature because it offers the option of delivering the propellant to the combustion chamber using only its natural vapor pressure and offers a potential for further reduction of the system simplicity. Unfortunately, because of this higher vapor pressure, hydrazine fumes significantly at room temperature and presents a high risk as a respiratory hazard. All hydrazine servicing operations must be performed with the use of Self Contained Atmospheric Protective Ensemble (SCAPE) suits.
Although procedures are in place to allow hydrazine to be managed safely on tightly controlled military reservations and at government-operated launch facilities; the toxicity and explosion potential of hydrazine requires extreme handling precautions. Increasingly, with a growing regulatory burden, infrastructure requirements associated with hydrazine transport, storage, servicing, and clean up of accidental releases are becoming cost prohibitive. As space flight operations continue to shift from government run organizations to private companies and universities operating away from government-owned test reservations, servicing payloads requiring hydrazine as a propellant becomes operationally infeasible. Extreme handling precautions generally do not favor hydrazine as a propellant for secondary payloads.
A recent study by the European Space Agency’s European Space Research and Technology Center (ESA/ESTEC) has identified two essential design elements to achieving low cost space access; 1) Reduced production, operational, and transport costs due to lower propellant toxicity and explosion hazards, and 2) Reduced costs due to an overall reduction in subsystems complexity and overall systems interface complexity.
1.3 Hydrazine Replacement Monopropellant Options
A useful monopropellant replacement for hydrazine must be chemically and thermally stable (for storage), but must easily decompose and have good combustion properties. Cryogenic or high freezing point propellants requiring temperature control are not appropriate space propulsion applications. Although mass-specific impulse is important, volume-specific impulse (density impulse) is an even more important consideration, and a high propellant storage density is preferred. Most importantly the propellant must be sufficiently stable to allow technicians and engineers to safely work with the propellant.
Hydrogen Peroxide (H2O2) is sometimes used as an oxidizing agent for bipropellant systems, and is currently being proposed as a “less toxic” alternative to hydrazine. Unfortunately, H2O2 offers a significantly lower overall performance than hydrazine with vacuum Isp slightly below 170 seconds. More importantly propulsion-grade solutions of H2O2 have an even higher room temperature vapor than hydrazine, approximately 1200 kPa (175 psia). Thus, while not as toxic as hydrazine, peroxide still presents a significant respiratory hazard. Propellant grade peroxide solutions are also unstable and present a moderate explosion risk. The reduced performance, coupled with the still significant objective and health hazards, do not favor hydrazine as a “green” alternative to hydrazine.
For the past 15 years the US Department of Defense (DoD) and the Swedish Space Corporation (SSC) subsidiary Ecological Advanced Propulsion Systems (ECAPS) have been pursuing green-propellant alternatives based on aqueous solutions of ionic liquids. Ionic liquids are water-soluble substances that normally exist in solid form at room temperature, but melt below the boiling point of water. When dissolved in water these materials exhibit very strong ion-to-ion interactions. Two of two of the currently most promising ionic liquid replacement options for hydrazine are Ammonium dinitramide (ADN) and Hydroxylamine nitrate (HAN). ADN melts at approximately 90-93 C, and HAN melts at approximately 44-45 C. In solid form both ADN and HAN are highly energetic salts with both reducing and oxidizing components. Consequently, in solid form both materials are unstable and potentially explosive.
Thus as mentioned in the previous paragraph for propellant applications both ADN and HAN are used as a used in concentrated aqueous solutions to order to limit the explosion potential. In typical applications a fuel component such as methanol is added to the solution to increase the propellant performance. Because these propellants are mixed in aqueous solutions, they possess a very low vapor pressure at room temperature, and do not present a respiratory hazard. Thus servicing operations can be performed with the use of SCAPE suits. This low vapor pressure is one of the primary reasons that these propellants are considered to be significantly less hazardous than either hydrazine or peroxide.
1.4 Characteristics of HAN based monopropellants
Physical properties
HAN/HN-based monopropellants have a density of 1.4 to 1.5 g/cm3, which is much higher than the density of 1.0 g/cm3 that those based on hydrazine have. This improves the density specific impulse and enables a reduction in the weight of the fuel tanks and other structures to be achieved. The solidifying point is as low as -35°C, which improves usability in low temperature environments. Their viscosity is slightly higher than that of those based on hydrazine but, depending on their composition, this is not so high as to make it impossible to supply these monopropellants.
Toxicity
Hydrazine is highly toxic and evaporates easily. When handling it, therefore, self-contained atmospheric protective ensemble (SCAPE) suits and the air supply devices for them need to be used, and special operation personnel and medical personnel need to be employed. What is more, the performance of any other work in parallel with hydrazine loading is prohibited. This reduces the work efficiency of launch site operations.
To evaluate the toxicity of HAN/HN-based monopropellants, the authors measured the amounts of toxic vapor they generated and evaluated the obtained measurements using the following hazard index. Their findings revealed that the monopropellant’s toxicity is as low as about 1/10 000 of that of hydrazine and 1/600 of methanol.
HAZARD INDEX (mmHg/PPM) = Vapor pressure (mmHg at 25°C)
LC50 (PPM 4H)
HAN/HN-based monopropellants can be handled with gloves, goggles, simplified masks, and other protective gear similar to that used for handling ordinary chemicals.
Self-sustaining combustion characteristics
HAN-based monopropellants undergo a rapid reaction in thermal and pressure load environments, and this reaction can suddenly retrograde as a result of the temperature rise caused by heat transfer. This could cause supply pipes and even fuel tanks to detonate. HN has been successfully used as a stabilizer for HAN, which acts as a reducer, in nuclear-related reprocessing processes, so it is expected to prove effective in inhibiting such detonations.
Figure 4.shows the burning rate of the HAN/HN-based monopropellant compared with the results for other compositions. HN free compositions exhibited very high burning rates within a range of one to several hundreds of millimeters per second under a pressure of 4 MPa, while the HN mixed monopropellant composed of four compositions had no self-sustaining combustion even under a pressure of 7 MPa, which is the upper limit for the equipment used. This indicates a high degree of safety for this monopropellant. However, increasing the content of water, which acts as a solvent, was not effective in inhibiting self-sustaining combustion. These results indicate that the addition of HN is an effective means of inhibiting self-sustaining combustion.
914400top
Figure 4: Burning rate of HAN/HN-based monopropellant
Here in this project we considered that the HAN is the best replacement of the Hydrazine and we carried out the analytical and simulation studies on standard size thruster used by some space agency in India for 11N thrust requirements.
1.5 Nozzle geometry
Space research organizations have been working on the green propellant attitude control thrusters for spacecraft applications and they have designed a standard size nozzle for the Hydrazine thrusters. Here in this project we considered the same nozzle for HAN green propellant thrusters,
The nozzle geometry is given below,
Fig 6.1: Nozzle geometry
For analytical calculations of pressure, temperature and velocity distribution along the nozzle the divergent part is divided by five parts as shown below,
Z in mm Area at Z location (AX10-6 M2) Area Ratio
0 A* = 3.97 1
3.73 A1 = 35.73 9
6.99 A2 = 63.52 16
11.19 A3 = 99.25 25
16.32 A4 = 142.97 36
22 Ae = 191.24 48
6.2 Assumed Model
Here in this the Steady, quasi-one-dimensional model is assumed. There are gradual variations in the geometry, so that the flow which is near to the nozzle walls is not strictly along the z-direction. However, the flow angularity is very small. The variation in properties can be calculated assuming that the properties are constant in each cross-section. The cross-section area, A, is a function of z-alone. Thus, all properties are functions of z alone.
A= A(z); u = u(z); T=T(z), p= p(z) etc.
Area ratio and Mach no relation:
AeA*=1Me2?+11+?-12Me2?+12?-1Isentropic relations:
T0Te=1+?-12Me2P0Pe=1+?-12Me2??-1By using the above equations, the pressure and temperature distribution along the nozzle is calculated. As we move from throat to the exit the cross sectional area of the nozzle is going to increase so here in these calculations the divergent section length is divided equally by five parts and at each part the area is calculated. After all five areas are known then mach no at each area is calculated by assuming each area is an exit area.
Here in these equations the isentropic index is the most influencing parameter. Its value for Hydrazine is around 1.22 and for HAN is 1.35. The area ratios along the length of the nozzle are (Ae /A*) = 1, 9, 16, 25, 36 and 48 at z =0 , 3.73 ,6.99 ,11.19 , 16.32, 22 mm.
1.6 NASA CEA Run
CEA stands for Chemical Equilibrium Application – designed and developed by NASA Glen Research centre. CEA is used to calculate the chemical product concentrations from any reactants and determines thermodynamic and transport properties for the given product mixture. This Built-in application includes the calculation of theoretical rocket performance, combustion properties, Chapman detonation parameters and shock tube parameters.
CEA INTERFASE
EXECUTION CONTROL
OUT PUT INTERFASE
CHAPTER 2
LITERATURE
Fukuchi B Apollo, Nagase Sakac , Maruzumi Haruki , Ayabe Munaco : “HAN/HN-Based Monopropellant Thrusters” , vol.43 No. 1-2010.
This paper presents the development and hot firing test results of a mixture of HAN/HN/TEAN/H2O. The HN addition realized a reliable catalytic ignition and combustion characteristics with 1 N – 20 N class thrusters and safety characteristics without detonability.
K. Neff and P. King : ” High Performance Green Propellants for Satellite Applications”, (45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit 2 – 5 August 2009, Denver, Colorado) Copyright © 2009 by Moog Inc. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission
This paper provides an overview of LMP-103S, a “High Performance Green Propellant” (HPGP) and outlines the benefits of LMP-103S as compared to hydrazine for monopropellant satellite applications. Corresponding performance and design characteristics of HPGP thrusters in the range of 1 N to 5 N are also reviewed. Satellite propulsion systems are then considered that utilize the HPGP thrusters and propellant.
Angelo Pasini, Lucio Torre, Giovanni Pace and Dario Valentini and Luca d Agostino: “Pulsed Chemical Rocket with Green High Performance Propellants” , (49th AIAA/ASME/SAE/ASEE Joint Propulsion)Conference Copyright © 2013 by ALTA. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission
This paper reports the concept and main ideas behind the project and explicates the baseline and improvements brought by the Pulsed Chemical Rocket concept by means of a preliminary assessment of the propulsion performance and system mass budget.
Klein, N., “Ignition and Combustion of the HAN-Based Liquid Propellants, “Proceedings of 27th JANNAF Combustion Subcommittee Meeting, CPIA Pub.557, Vol. 1, Chemical Propulsion Information Agency, Laurel, MD, 1990, pp.443-450.
This paper described the combustion of HAN-based propellants containing triethanol ammonium nitrate (TEAN) as the fuel component as a three-stage reaction: (1) Initiation,(2) ignition, and (3) combustion. The whole reaction is initiated by the decomposition of HAN in condensed-phase producing heat and gases. A spray of molten TEAN droplets was also produced in this stage. The “ignition” stage started as TEAN droplet started to become involved in the reaction and more heat was released so that the reaction could be self sustained. The final stage was defined as “combustion” where most of the energy stored in the propellant is released.
Vosen, S. R., “The Burning Rate of Hydroxylammonium Nitrate-Based Liquid Propellants,” 22nd Symposium (International) on Combustion, Combustion Institute, Pittsburgh, PA,1988, pp. 1817-1825.
This paper suggested the combustion of LP 1846 occurred in two stages: (1) Exothermic decomposition of HAN in condensed-phase and (2) gas-phase reaction between the product of HAN decomposition and TEAN.
Vosen, S. R., “Hydroxylammonium Nitrate-Based Liquid Propellant Combustion – Interpretation of Strand Burner Data and the Laminar Burning Velocity,”Combustion and Flame, Vol. 82, Nos. 3-4, 1990, pp. 376-388
Vosen observed a luminous flame in the strand burner tests of LP1846 when the pressure was higher than 26.7MPa. The luminous flame was positioned above a transparent zone referred to as a dark zone in nitramine propellants on top of the liquid-gas interface which was identified as the reaction front. At lower pressures, no visible flame was observed. The transparent zone became opaque as pressures were lowered. The different reaction zones observed in the strand burner are one indication of the staged-combustion characteristics.
Chang, Y.-P., and Kuo, K. K., “Assessment of Combustion Characteristics and Mechanism of Hydroxylammonium Nitrate-Based Liquid Monopropellant,” Journal of Propulsion and Power, Vol. 18, No. 5, September 2002, pp. 1076-1085
Chang and Kuo observed the same phenomena as Vosen in strand burner tests of XM 46, which was formerly referred to as LP1846. A luminous flame appeared when the pressure was higher than 28MPa while the reaction was not self-sustainable at pressures lower than 1.82MPa. Different reaction zones were seen in the strand burner. The measured time-temperature profiles were able to be associated with different combustion stages. At a relatively low pressure of 13.8MPa where no flame appeared, the temperature measured in the dark opaque zone was around 300oC. The start of the clearing of the dark opaque gases at different pressures corresponded to a temperature of 600oC. At 33.8 MPa, the appearance of the luminous flame caused a sudden rise of temperature to 1700oC.
Chang, Y.-P., Boyer, E. and Kuo, K. K., “Combustion Behavior and Flame Structure of XM46 Liquid Propellant,” Journal of Propulsion and Power, Vol. 17, No. 4, July 2001, pp. 800-808
Chang et al. suggested that the opaque gases might include brown-colored NO2 and some heavy carbon-containing intermediate species. The temperature rose to a higher level when the opaque gases disappeared into transparent species, and the reaction subsequent to the transparent species finally produced the luminous flame, where most of the chemical energy in the propellant was released.
C. Scharlemann, M. Schiebl, R. Amsüss, M. Tajmar, “Development of Miniaturized Green Propellant Based Mono- and Bipropellant Thrusters”- 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit 8 – 11 July 2007, Cincinnati, OH
This paper discusses the successful development and test of a two thruster systems operating with green propellants: a monopropellant and a bipropellant thruster. The monopropellant system developed by ARC was further improved and tested. The catalytic decomposition with its advanced monolithic catalyst has obtained efficiencies up to 99% resulting in decomposition temperatures of 650°C. Performance measurements on a thrust balance has shown that the thruster can generate thrust between 150 and 1700 mN with a specific impulse of 153 s. Additional improvements presently implemented have the potential to increase the specific impulse to 164 s.
Stephen A. Whitmore, Daniel P. Merkley, Shannon D. Eilers, Michael I. Judson, “Development and Testing of a Green Monopropellant Ignition System”- 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit San Jose Convention Center , San Jose, CA, 15-17 July 2013
This paper will detail the development and testing of a “green” monopropellant booster ignition system. The proposed booster ignition technology eliminates the need for a pre-heated catalyst bed, a high wattage power source, toxic pyrophoric ignition fluids, or a bi-propellant spark igniter. The design offers the simplicity of a monopropellant feed system features non-hazardous gaseous oxygen (GOX) as the working fluid.
Ronald A. Spores, Robert Masse, Scott Kimbrel, ” Green Propellant Infusion Mission (GPIM) Technology Demonstration Mission (TDM) will demonstrate an operational AF-M315E green propellant propulsion system.”- 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Exhibit AIAA 2013-3849 15-17 July 2013, San Jose, California
This paper statuses the propulsion system module development, including thruster design, system design and system component materials compatibility testing. Major system components of the propulsion system module include: propellant tank, latch valve, service valve and thruster valve.
B.V.V. Naga Sudhakar, B Purna Chandra Sekhar, P Narendra Mohan, Md Touseef Ahmad, “Modeling and simulation of Convergent-Divergent Nozzle Using Computational Fluid Dynamics”- International Research Journal of Engineering and Technology (IRJET) e-ISSN: 2395 -0056 Volume: 03 Issue: 08 | Aug-2016.
This paper aims to study the behavior of flow in convergent divergent nozzle by analyzing various parameters like pressure, temperature and velocity using computational fluid dynamics software (C.F.D).These results were further plotted comparing them with analytical values.
Rocket Propulsion Elements- by George P Sutton and Oscar Biblarz
Applied Gasdynamics- by Radhakrishna
Modern compressible flow -by S M Yaya
Gasturbines -By V Ganeshan
Summary of the literature:
Hydroxyl ammonium nitrate (HAN), the nitrate salt of hydroxylamine with chemical formula NH3OHNO3, is a solid at room temperature and melts at 48 C. It can be mixed with water up to a concentration of 95% to make a salt mixture in ionic form with cationNH3OH+ and anion NO3-.
HAN, by itself, could undergo exothermic decomposition when proper activation energy is provided while the excess of oxygen in HAN makes it suitable as an oxidizer. HAN-based propellants which consist of HAN as the oxidizer, and fuel and water, as the solvent and stabilizer have long been considered as potential propellants in liquid gun propulsion and rocket propulsion. The properties and energy content of the propellant can be altered by adjusting the amount of water and utilizing different fuels depending on the required performance. More recently, HAN-based propellants were also proposed as monopropellants for thruster application and often seen as a potential candidate as an alternative to hydrazine because of its low toxicity and high energy density over the state-of- the-art monopropellant hydrazine.
CHAPTER 3
METHODOLOGY
The combustion properties and adiabatic flame temperatures were simulated in NASA CEA run for Hydrazine and HAN
With reference to the CEA results of adiabatic flame temperature and isentropic index of the products of combustion at particular pressure, the theoretical calculations made for standard size of thruster nozzle for Hydrazine and HAN
With the same nozzle geometry and combustion chamber pressure the CEA run is made for the rocket propulsion parameters for Hydrazine and HAN
Analysis of the nozzle is done with respect to the combustion chamber pressure and adiabatic temperature using Ansys work bench for both Hydrazine and HAN
CHAPTER 4
CASE STUDY
4.1 Case study 1: For Hydrazine
Properties of Hydrazine
4.1.1CEA run For Hydrazine
INPUT:
problem case = CE01 o/f=1,
rocket equilibrium frozen nfz=1
p,bar=16,
sup,ae/at=9,16,25,36,48,
react
fuel=N2H4(L) moles=1 t,k=450
oxid=N2H4(L) moles=1 t,k=450
OUTPUT:
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NASA-GLENN CHEMICAL EQUILIBRIUM PROGRAM CEA2, FEBRUARY 5, 2004 BYBONNIE MCBRIDE AND SANFORD GORDON
REFS: NASA RP-1311, PART I, 1994 AND NASA RP-1311, PART II, 1996
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prob case=CE011776 ro equilibrium
! iac problem o/f 1
p,bar16
supar 9 16 25 36 48 reac
fuelN2H4(L) wt%=100 t,k=450 oxid N2H4(L) wt%=100 t,k=450 outputshort
output trace=1e-5 end
THEORETICAL ROCKET PERFORMANCE ASSUMING EQUILIBRIUM COMPOSITION DURING EXPANSION FROM INFINITE AREA COMBUSTOR
Pin =232.1 PSIA CASE = CE011776
REACTANTWT FRACTION
(SEE NOTE) ENERGY
KJ/KG-MOL TEMP
K
FUEL N2H4(L) 1.0000000 66331.049 450.000
OXIDANT N2H4(L) 1.0000000 66331.049 450.000
O/F=1.00000%FUEL= 50.000000R,EQ.RATIO= 0.000000PHI,EQ.RATIO=-1.000000
CHAMBER THROAT EXIT EXIT EXIT EXIT EXIT
Pinf/P1.0000 1.8725 82.510 171.35 297.95 465.07 658.48
P, BAR16.000 8.5445 0.19392 0.09337 0.05370 0.03440 0.02430
T, K1041.58 882.43 428.48 396.31 375.60 360.66 349.88
RHO, KG/CU M1.9779 0 1.2486 0 6.4490-2 3.4648-2 2.1533-2 1.4643-2 1.0819-2
H, KJ/KG2069.92 1602.58 -92.871 -300.85 -444.25 -552.02 -631.91
U, KJ/KG1260.99 918.27 -393.56 -570.34 -693.63 -786.97 -856.50
G, KJ/KG-14533.7 -12464.1 -6923.22 -6618.38 -6431.58 -6301.21 -6209.33
S, KJ/(KG)(K)15.9408 15.9408 15.9408 15.9408 15.9408 15.9408 15.9408
M, (1/n)10.706 10.722 11.848 12.227 12.522 12.763 12.953
(dLV/dLP)t-1.00223 -1.00370 -1.07776 -1.09226 -1.10075 -1.10611 -1.10942
(dLV/dLT)p1.0142 1.0274 2.0616 2.3443 2.5356 2.6733 2.7701
Cp, KJ/(KG)(K)2.9913 3.0143 12.7352 15.8184 17.9891 19.5991 20.7608
GAMMAs1.3603 1.3659 1.1855 1.1682 1.1581 1.1513 1.1466
SON VEL,M/SEC1049.0 966.8 597.0 561.1 537.4 520.1 507.5
MACH NUMBER0.000 1.000 3.484 3.881 4.173 4.403 4.581
PERFORMANCE PARAMETERS Ae/At 1.0000 9.0000 16.000 25.000 36.000 48.000
CSTAR, M/SEC 1325.4 1325.4 1325.4 1325.4 1325.4 1325.4
CF 0.7294 1.5692 1.6429 1.6918 1.7277 1.7538
Ivac, M/SEC 1674.6 2224.4 2301.3 2353.6 2392.6 2421.2
Isp, M/SEC 966.8 2079.8 2177.5 2242.4 2290.0 2324.6
MOLE FRACTIONS *H2 6.6480-1 6.6355-1 5.7567-1 5.4608-1 5.2306-1 5.0427-1 4.8949-1
NH3 2.2429-3 3.7424-3 1.0920-1 1.4470-1 1.7233-1 1.9487-1 2.1262-1
*N2 3.3296-1 3.3271-1 3.1513-1 3.0922-1 3.0461-1 3.0085-1 2.9790-1
4.1.2 Theoretical calculations of Hydrazine:
From literature,
P0=16 barT0 = 1183 K
? = 1.22
we know that , area ratio is given by,
AA*=1M2?+11+?-12M2?+12?-1Substitute the value of ?= 1.22 in the above equation we get ,
AA*15*M15 =0.90+0.099M2At throat, z= 0 the area ratio is one therefore mach no at the throat is one,
At z=3.73mm from the throat the area ratio is 9 then,
9^15*M1^15 =0.90+0.099M1^2 By solving the above equation for Me, by inspection
We get M1 = 3.35
At z=6.99, area ratio is 16, then
16^15*M2^15 =0.90+0.099M2^2On solving, we get M2= 3.74
At z=11.19, area ratio= 25,
25^15*M3^15 =0.90+0.099M3^2The M3= 4.02
At z= 16.32, area ratio=36,
36^15*M4^15 =0.90+0.099M4^2Therefore M4= 4.25
At the exit of the nozzle (z=22 mm), area ratio=48,
48^15*Me^15 =0.90+0.099Me2 On solving the above equation we get Me= 4.42.
For the pressure and temperature distribution
We know that
T0Te=1+?-12Me2P0Pe=1+?-12Me2??-1By using the above equations we can calculate the value of pressure and temperature
Z(mm) A/A* Me at Z Pe=Pz (bar) Te=Tz (K)
0 1 1 8.97 1063
3.73 9 3.35 0.1860 529
6.99 16 3.74 0.09175 465
11.19 25 4.02 0.05573 425.9
16.32 36 4.25 0.03727 396
22 48 4.42 0.02781 376
As at the exit of the nozzle the mach no is 4.42 and the temperature is Te= 376K therefore speed of sound is found from a=?RTe . we know that for Hydrazine ?=1.22, R= 256 J/Kg-K then a=1.22*256*376 = 342 m/s. The exit velocity is given by the relation Ve= Me*a , Ve= 4.42*342=1511 m/s.
Then thrust is given by F= m*Ve +Pe-P?Ae,
Where, m=mass flow rate (4.6grams/second).
P?= Atmospheric pressure (for vacuum P?=0 )
Pe = Pressure at the exit of the nozzle (0.02781 bar)
Ae = Area of the exit (191.23 E-6 m2)
Ve = velocity at the exit (1511 m/s)
F= 7.5 N
Then specific impulse is given by,
Isp= Fmxg Seconds
= 7.54.6X10-3X9.81Isp= 166 seconds
Theoretical Specific Impulse of Hydrazine is 166 seconds
4.1.3Ansys simulation
4.2 Case study 2: For HAN
Properties of HAN
4.2.1 CEA run FOR HAN
INPUT:
problem case = han01 o/f=5.2,
rocket equilibrium frozen nfz=1
p,bar=16,
sup,ae/at=9,16,25,36,48,
react
oxid=NH4NO3(L) mass fraction= 5 t,k=443
oxid=H2O(L) mass fraction= 8 t,k=443
fuel=CH3OH(L) mass fraction= 21 t,k=390
oxid=HAN mass fraction= 95 t,k=443
OUTPUT:
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**************************************************************************** NASA-GLENN CHEMICAL EQUILIBRIUM PROGRAM CEA2, MAY 21, 2004
BY BONNIE MCBRIDE AND SANFORD GORDON
REFS: NASA RP-1311, PART I, 1994 AND NASA RP-1311, PART II, 1996
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THEORETICAL ROCKET PERFORMANCE ASSUMING EQUILIBRIUM
COMPOSITION DURING EXPANSION FROM INFINITE AREA COMBUSTOR
Pin = 232.1 PSIA
CASE = han01
REACTANT WT FRACTION ENERGY TEMP
(SEE NOTE) KJ/KG-MOL K
OXIDANT NH4NO3(L) 0.0457831 -331414.678 443.000 OXIDANT H2O(L) 0.0746988 -274780.257 443.000 FUEL CH3OH(L) 1.0000000 -230404.520 390.000 OXIDANT HAN 0.8795181 -399000.000 443.000
O/F=4.20000 %FUEL= 19.230769 R, EQ.RATIO= 1.068208 PHI,EQ.RATIO= 1.180290
CHAMBER THROAT EXIT EXIT EXIT EXIT EXIT
Pinf/P 1.0000 1.7714 73.989 162.46 297.61 487.48 719.37
P, BAR 16.000 9.0324 0.21625 0.09848 0.05376 0.03282 0.02224
T, K 2244.54 2046.46 1040.76 888.12 782.33 703.17 645.11
RHO,KG/CU M 1.9066 0 1.1812 0 5.5620-2 2.9684-2 1.8395-2 1.2495-2 9.23-3
H, KJ/KG -5407.39 -5865.74 -7959.39 -8242.33 -8431.02 -8567.84
-8665.79U, KJ/KG -6246.57 -6630.42 -8348.18 -8574.10 -8723.27 -8830.52 -8906.78
G, KJ/KG -31233.5 -29412.8 -19934.6 -18461.3 -17432.6 -16658.6 -16088.6
S, KJ/KG-K 11.5062 11.5062 11.5062 11.5062 11.5062 11.5062 11.5062
M, (1/n) 22.239 22.251 22.257 22.257 22.257 22.257 22.258
(dLV/dLP)t -1.00044 -1.00013 -1.00000 -1.00000 -1.00000 -1.00001
-1.00005
(dLV/dLT)p 1.0131 1.0041 1.0000 1.0000 1.0000 1.0001 1.0008
Cp, KJ/(KG)(K) 2.4083 2.2697 1.8916 1.8137 1.7525 1.7044 1.6729
GAMMAs 1.1889 1.1988 1.2461 1.2594 1.2709 1.2808 1.2880
SON VEL,M/SEC 998.9 957.5 696.0 646.4 609.4 580.0 557.1
MACH NUMBER 0.000 1.000 3.246 3.684 4.035 4.335 4.582
PERFORMANCE PARAMETERS Ae/At 1.0000 9.0000 16.000 25.000 36.000 48.000
CSTAR, M/SEC 1414.8 1414.8 1414.8 1414.8 1414.8 1414.8 CF 0.6767 1.5969 1.6831 1.7382 1.7771 1.8044
Ivac, M/SEC 1756.1 2431.3 2520.5 2578.0 2618.6 2647.2Isp, M/SEC 957.4 2259.2 2381.2 2459.1 2514.1 2552.8
MOLE FRACTIONSCH4 0.00000 0.00000 0.00000 0.00000 0.00000 0.00000 0.00001
*CO 0.02865 0.02673 0.00848 0.00468 0.00254 0.00137 0.00077
*CO2 0.10482 0.10682 0.12510 0.12890 0.13104 0.13221 0.13280
*H 0.00033 0.00014 0.00000 0.00000 0.00000 0.00000 0.00000*H2 0.03311 0.03460 0.05274 0.05653 0.05867 0.05983 0.06039
H2O 0.65713 0.65652 0.63878 0.63498 0.63284 0.63168 0.63111*NO 0.00011 0.00002 0.00000 0.00000 0.00000 0.00000 0.00000*N2 0.17471 0.17485 0.17491 0.17491 0.17491 0.17491 0.17491
*O 0.00001 0.00000 0.00000 0.00000 0.00000 0.00000 0.00000*OH 0.00108 0.00032 0.00000 0.00000 0.00000 0.00000 0.00000*O2 0.00006 0.00001 0.00000 0.00000 0.00000 0.00000 0.00000
THEORETICAL ROCKET PERFORMANCE ASSUMING FROZEN COMPOSITION
Pin = 232.1 PSIA
CASE = han01 REACTANT WT FRACTION ENERGY TEMP
(SEE NOTE) KJ/KG-MOL K
OXIDANT NH4NO3(L) 0.0457831 -331414.678 443.000
OXIDANT H2O(L) 0.0746988 -274780.257 443.000FUEL CH3OH(L) 1.0000000 -230404.520 390.000OXIDANT HAN 0.8795181 -399000.000 443.000
O/F= 4.20000 %FUEL= 19.230769 R,EQ.RATIO= 1.068208 PHI,EQ.RATIO= 1.180290
CHAMBER THROAT EXIT EXIT EXIT EXIT EXIT
Pinf/P 1.0000 1.7744 75.177 166.10 305.77 502.67 743.75P, BAR 16.000 9.0173 0.21283 0.09632 0.05233 0.03183 0.02151
T, K 2244.54 2038.70 1019.02 863.68 756.72 677.36 619.57RHO,KG/CU M 1.90660 1.18300 5.5863-2 2.9830-2 1.8495-2 1.2569-2 9.2869-3H, KJ/KG -5407.39 -5866.27 -7948.02 -8226.57 -8411.21 -8544.37 -8639.30
U, KJ/KG -6246.57 -6628.49 -8329.01 -8549.48 -8694.12 -8797.62 -8870.94G, KJ/KG -31233.5 -29324.0 -19673.1 -18164.2 -17118.1 -16338.2 -15768.2S, KJ/(KG)(K)11.5062 11.5062 11.5062 11.5062 11.5062 11.5062 11.5062
M, (1/n) 22.239 22.239 22.239 22.239 22.239 22.239 22.239
Cp,KJ/(KG)(K) 2.2517 2.2058 1.8321 1.7537 1.6986 1.6576 1.6278
GAMMAs 1.1991 1.2041 1.2564 1.2710 1.2822 1.2913 1.2982
SONVEL,M/SEC 1003.1 958.0 691.9 640.6 602.3 571.8 548.4
MACH NUMBER 0.000 1.000 3.258 3.707 4.069 4.380 4.636
PERFORMANCE PARAMETERS Ae/At 1.0000 9.0000 16.000 25.000 36.000 48.000
CSTAR, M/SEC 1411.8 1411.8 1411.8 1411.8 1411.8 1411.8 CF 0.6786 1.5967 1.6820 1.7362 1.7742 1.8009
Ivac, M/SEC 1753.6 2423.2 2510.5 2566.5 2605.9 2633.5Isp, M/SEC 958.0 2254.2 2374.5 2451.0 2504.8 2542.4
MOLE FRACTIONS *CO 0.02865 *CO2 0.10482 *H 0.00033
*H2 0.03311 H2O 0.65713 *NO 0.00011 *N2 0.17471 *O 0.00001 *OH 0.00108 *O2 0.00006
4.2.2 Theoretical calculations of HAN:
From literature,
P0=16 barT0 = 2273 K
? = 1.28
We know that, Area ratio is given by,
AA*=1M2?+11+?-12M2?+12?-1Substitute the value of ?= 1.28 in the above equation we get,
AeA*^14*Me^14 =0.8771+0.1228Me2At throat, z= 0 the area ratio is one therefore mach no at the throat is one,
At z=4.4mm from the throat the area ratio is 9 then,
9^14*Me^14 =0.8771+0.1228Me2By solving the above equation for Me, by inspection
We get Me = 3.45
At z=8.8, area ratio is 16, then
16^14*Me^14 =0.8771+0.1228Me2On solving, we get Me= 4
At z=13.2, area ratio= 25,
25^1/4*Me^14 =0.8771+0.1228Me2The Me= 4.45
At z= 17.6, area ratio=36,
36^14*Me^1/4 =0.8771+0.1228Me2Therefore Me= 4.75
At the exit of the nozzle (z=22 mm), area ratio=48,
48^14*Me^14 =0.8771+0.1228Me2On solving the above equation we get Me= 4.9.
For the pressure and temperature distribution
We know that
T0Te=1+?-12Me2P0Pe=1+?-12Me2??-1By using the above equations we can calculate the value of pressure and temperature
Z A/A* Me at Z Pe=Pz (bar) Te=Tz (K)
0 1 1 8.78 1993
3.73 9 3.45 0.1807 852
6.99 16 4.0 0.07416 701
11.19 25 4.45 0.03699 602
16.32 36 4.75 0.02369 546
22 48 4.9 0.01905 521
As at the exit of the nozzle the mach no is 4.9 and the temperature is Te= 521K therefore speed of sound is found from a=?RTe . we know that for HAN ?=1.28, R= 252 J/Kg-K then a=1.28*252*521 = 409 m/s. The exit velocity is given by the relation Ve= Me*a , Ve= 4.9*409=2004 m/s.
Then thrust is given by F= m*Ve +Pe-P?Ae,
Where, m=mass flow rate (4.6grams/second).
P?= Atmospheric pressure (for vacuum P?=0 )
Pe = Pressure at the exit of the nozzle (0.01905 bar)
Ae = Area of the exit (191.23 E-6 m2)
Ve = velocity at the exit (2004 m/s)
F= 9.58 N
Then specific impulse is given by,
Isp= Fmxg Seconds
= 9.584.6X10-3X9.81Isp= 212 seconds
Theoretical Specific Impulse of HAN is 212 seconds
Ansys simulation for HAN
Results
CHAPTER 5
Results and Discussion
5.1 Validation plots:
For Hydrazine:
lefttop
The pressure, temperature and velocity distribution along the nozzle calculated analytically and simulated in NASA CEA Run are closely matching as shown in the above plots and the specific impulse fond from analytical and CEA are 166 seconds and 230 seconds respectively , the deviation between these are within 30% so the results of CEA run are valid for Hydrazine.
For HAN:
The pressure, temperature and velocity distribution along the nozzle determined analytically and simulated in NASA CEA Run are closely matching as shown in the above graphs and the specific impulse fond from analytical and CEA are 212 seconds and 256 seconds respectively, so the results of CEA run are valid for HAN.
5.2 Comparison plots:
Above plots compares the results of pressure distribution temperature distribution and velocity distribution calculated for HAN and Hydrazine from Analytical and CEA. The exit pressure for HAN and Hydrazine thrusters are found almost same but the exit temperature and velocity for HAN is higher than that of the Hydrazine because the adiabatic flame temperature of the HAN combustion is around 2273K which is almost double the value of Hydrazine (1183K). The velocity of the exit gases of HAN combustion is around Mach 4.9 and that for Hydrazine is 4.58, so the specific impulse is large for HAN because the thrust produced for the constant mass flow rate, is higher than that of Hydrazine thrusters. So the HAN monopropellant is the best replacement for Hydrazine.
CHAPTER 6
CONCLUSION
References:
