Design, Fabrication and Thrust Measurement Approaches for Microthrusters

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Design, Fabrication and Thrust Measurement Approaches for Microthrusters

Design, Fabrication and Thrust Measurement Approaches for Microthrusters

Nisha. R

UG scholar

Dept of Aeronautical Engineering, Rajalakshmi Engineering College Thandalam , Chennai,Tamilnadu-India

Dharani. V1

UG scholar

Dept of Aeronautical Engineering Rajalakshmi Engineering College Thandalam, Chennai,Tamilnadu-India

Abstract: In order to address the need for small amount of thrust for pico and nano satellites , it is suggested to use micro thrusters which is a vital outcome of MEMS technology .The objective of this work is to design a micro thruster with minimum mass and to withstand heavy working conditions. These micro thrusters are used for altitude alterations and other reaction control system .The required thrust for satellite weighting from 1-10kg will be 10µN to 100mN.Design calculations were plotted for thrust 60µN and mach 2. Here monopropellant hydrogen peroxide is suggested because it has less specific impulse which will suit our case. Catalyst will be used to decompose the monopropellant which will be placed in the bed of thrust chamber. Lithography process suits for material patterning to desired shape. The thrust measurement stand with a force sensor, propellant feed system with a motor driven syringe pump and imaging devices were installed for the firing test of the micro thruster.

Key words: Micro thruster,fabrication of thrust,

I. INTRODUCTION

MEMS technology finds one of its best applications in space,because space launching of satellites has become increasingly cost sensitive. It is not a coincidence that more and more satellite researchers are involved in the development of tiny satellites which cost only a fraction of their larger counterparts. Launching a nanosatellite costs 1/10001/100 of launching a larger satellite. Along with the miniaturization of satellites, new concepts of deploying them to carry out practical missions have been proposed by operating the nanosatellites in a squad formation. This concept allows nanosatellites to carry out missions reliably, effectively and inexpensively . The new operational concept of nanosatellites, however, is based on the premise that each satellite in the squad is capable of relocating its position and adjusting its velocity. Unless the nanosatellites are equipped with a micro propulsion system, orbital transfer of satellite motion is impossible. The required thrust depends on the weight of a satellite as well as the type of maneuvering required by the mission .MEMS research responded to this need and micro propulsion became one of the major fields in power MEMS devices over recent decades . Most of the past research on micro propulsion devices has been based on solid propellant. However, conventional HTPB/AP composite solid propellant could not sustain stable combustion in a

microthruster, due to the excessive heat loss associated with the size of the device. Varying degrees of successful fabrication of MEMS thrusters was reported, using a high explosive as an alternative solid propellant. The energy content of high explosives was enough to overcome the excessive heat loss and still provide thermal energy to be converted into the kinetic energy of the burnt gas to propel the thruster. Successful fabrication and firing tests of solid propellant microthrusters were reported by a number of research groups. However, the inherent problems of solid propellant persisted; they were unable to maintain the thrust axis to coincide with the desired action line and were incapable of continuous firing. Additionally, the combustion of a high explosive propagates as a shock wave and precise prediction of the thrust was impossible. To avoid the problems related to solid propellant, micro propulsion devices using alternative propellants have been proposed and tested; bipropellant microthrusters, high pressure cold gas microthrusters , and monopropellant microthrusters. A bipropellant microthruster has high specific impulse but the system is complex because both oxidizer and fuel must be pressure-fed and metered. The construction of a cold gas jet microthruster is simple but the specific impulse is too low. Considering all the options, we concluded monopropellant is the appropriate choice for a MEMS propulsion system, although monopropellant has a specific impulse less than bipropellant. Past investigations on monopropellant microthrusters reported limited success.

Eqn. 1 The chemical reaction of hydrogen peroxide

PRELIMINARY DESIGN CALCULATIONS:

For design calculations it is necessary to fix propellent and oxidizers for calculating chamber temperature.it is irrelevant to fix the catalyst since it only enhances the chemical reaction not involving in it.

TABLE 1 DIFFERENT THRUSTER TYPES

The combustion chamber temperature can be evaluated using the classical reaction enthalpy balance equations or, as an alternative, one of the several available software tools for chemical equilibrium calculations.

The specific heat ratio and the molecular weight of the exhaust gases are evaluated in a similar way. Then, the characteristic

velocity of the propellant is calculated by means of the classical equation:

where Tc is the combustion chamber temperature, is the specific heat ratio of the exhaust gases and R is their gas constant (which, in turn, is a function of the molecular

weight).According to the frozen-flow approximation, the throat temperature (Tt), pressure (pt), density ( t) and velocity (ut) are given by the following relations:

where pc is the combustion chamber pressure. Exhaust Mach number (Me), temperature (Te)

and density ( e) are calculated using the relations:

where pe is the exhaust pressure. Then, isentropic nozzle equations give the possibility of calculating

THRUSTER

CLASS

SPECIFI

TYPICAL

MICRO-

C

IMPULS

TYPE

E

PROPELLENTS

SCALING

(s)

ISSUES

Cold gas

Chemical

40-80

Freons, N2, Ar

Drag losses

Monopropellent

Chemical

180-220

N2H4, H2O2

Drag losses

Bipropellent

Chemical

300-440

N2H4+ N2O4, H2

+ O2

Propellent mixing,high flammable temperature.

Solid

Chemical

100-290

Nitocellulose + Nitroglycerine

Ignition

where pa is the ambient pressure. Finally the exhaust velocity, throat area and propellant mass flow rate are given by the following equations:

where F is the required thrust.

With above equations ,the calculated values must be used as a reference to build chamber volume,feedline length and its diameter.

DESIGN OF MICROCHANNEL AND CHAMBER:

The inlet feed line should ensure uninterrupted propellant supply to the chamber at required flow rate. At the microscale domain, various factors like fluid resistance, surface tension,viscosity, surface to volume ratio, etc are important factors and may result in slow delivery or even total blockage of the microchannel if not properly designed . In microfluidic channels, the Reynolds numbr (Re) is very small thereby ensuring laminar flow of the fluid. The governing equation for the Reynolds number is given by .

The fluidic resistance (R) offered by the microchannel is very high compared to the macroscale device due to very small crosssectional area through which fluid has to move. For a rectangular microchannel with low aspect ratio (i.e. w

h), the resistance can be stated as,

For a rectangular microchannel with high aspect ratio (w

<< h), the fluid resistance (R) can be approximated as,

12

R=

the exhaust/throat area ratio and the thrust coefficient:

3

where L, w and h are the length, width and height of the

microchannel, respectively. Surface roughness of the

channel is another important factor in controlling the fluid flow through themicrochannel. This is related with the Reynolds number by the Darcy friction factor ( f ) = 64/Re. Since Re is very small in the microfluidic flow, the Darcy friction factor f is very high. This causes high fluidic resistance to the fluid flow and is inversely proportional to the hydraulic diameter of the channel.Thus, a suitable geometrical design of the microfluidic channel is necessary for the normal flow of the propellant from its reservoir to the VLM chamber with minimum pressure drop in the channel. Considering various force factors acting against the fluid motion in the microchannel, the friction factor has been analytically calculated for various hydraulic diameters of the channel and the result is shown in figure 2. It is observed from the graph that the friction factor f is appreciably low

Fig.1 Plot of Darcy friction factor vs Diameter

and does not vary much for the Dh value more than 500 m.However, the f factor increases rapidly to the higher values as the Dh value decreases below 500 m. Considering this analytical result, the width of the microchannel was designed to 500 m in this study to achieve minimum friction factor effects resulting minimum pressure drop in the channel during fluid flow. The pressure drop in themicrochannel has been calculated using the Darcy Weisbach equation.

where pdrop is the pressure drop (Pa) in the channel and is the friction coefficient.A typical graph for the variation of pressure drop with channel length is shown in figure considering the hydraulic diameter 500 m and inlet flow rate 2 mg s1. In order to optimize between the friction factor and pressure loss, the hydraulic diameter (Dh) and channel length (L) were designed to 500 m and 1000 m respectively in this study by interpreting the above results.

Fig.2 Pressure drop vs Channel length.

  1. DESIGN OF NOZZLE:

    Among various types of nozzles layout, the De-Lavel

    i.e. converging-diverging (C-D) type nozzle is the most suitable configuration for present applications. After decomposition in catalyst chamber, the gaseous products are delivered to the converging area of the exit nozzle throat, subsequently accelerate to supersonic speed in the diverging part of the nozzle. For a steady flow operation, the thrust produced at exit location of the nozzle can be approximated as F ~ m* Vexit (4) where m* = mass flow rate, Vexit = exit velocity of the gas.

    The nozzle, having symmetrical configuration along its axial direction will be most suitable for the present applications. The gradual contour serves the purposes of straightening the outlet flow and producing a uniaxial thrust normal to the exit plane. However, a limited configuration of nozzle design is practically feasible by micromachining technology. In most of the MEMS nozzles, the cross section is rectangular and hence it is not asymmetric.The flow in a MEMS supersonic nozzle can be substantially affected by viscous effects.

    This is evidenced by the Reynolds number of the flow, which is a measure of the relative magnitudes of inertial and viscous forces. While designing the nozzle following point should be taken into consideration.

    1. Low Reynolds Number:

      The divergence angle should be chosen to keep the Reynolds Number <200 to avoid the discontinuity at the wall.

    2. Choking of the mass flow:

      In the diverging nozzle section the viscous boundary layer extends away from the wall and into the bulk flow. The end result of the boundary layer growth is choking of the mass flow rate and result in reduction of thrust and efficiency.

    3. Larger divergence angles:

    Results in the literature have suggested that low-Reynolds microscale nozzles will perform better with larger divergence

    angles ~28º as opposed to the macroscale design with smaller angle of 15º20º.

    FABRICATION:

    The material used for a micro monopropellant thruster should resist erosion caused by hydrogen peroxide. The microthruster wall should have a low thermal conductivity to prevent thermal energy from escaping. Previous studies on micro scale thrusters conserved heat energy using ceramics and glasses. We selected photosensitive glass that has been proved to be excellent in reacting flow MEMS devices. The thermal conductivity of glass is less than one hundredth of that of silicon. Additionally, wet etching glass wafer at high aspect ratio is efficient and inexpensive. Figure shows the MEMS fabrication procedure used in the present study. The design of the micro monopropellant thruster was printed on a film mask and then transferred onto a chromium mask on a quartz wafer by lithography. Photosensitive glass diced by 20 mm × 20 mm was exposed to 310 nm wavelength UV light through the chromium mask. Heat treatment at 585 °C for 45 min made the UV exposed portion of the photosensitive glass crystallized around silver atoms. The UV exposed part is easily etched in the hydrofluoric at an etching rate of 12 m per minute. After the etching process, the wafer was polished to make the surface smooth for bonding. The complete structure of the microthruster consisted of five layers of prefabricated glass wafer. Four layers of thruster were joined by thermal bonding and the last layer by a UV bonding technique. Before UV bonding of the fifth layer enclosed the thrust chamber, prepared catalyst grains were inserted into the chamber along with a thin nickel foam screen to prevent the catalyst grains from escaping the catalyst bed region. Among the five glass layers , the layer in the center has the pattern of injector and nozzle profile. The center layer is 250 m thick and thinner than the upper and lower layers for precise fabrication of injector and nozzle. The center layer is joined by thicker upper and lower layers that are 1 mm thick. The top and bottom layers are also 1 mm thick.Thermal bonding was conducted at 500 °C for 12 h under 2100 Pa in a furnace. UV bonding was done with liquid adhesive (Optical Adhesives & Coatings, OP24 produced by DYMAX) on a spin coater rotating at 4000 rpm for 50 s. Then 1.5 min of exposure under 310 nm UV light completed the bonding process.

    Fig.3 The MEMS fabrication procedure

    1. LITHOGRAPHY: Lithography in the MEMS context is typically the

      transfer of a pattern to a photosensitive material by selective exposure to a radiation source such as light.

      A photosensitive material is a material that experiences a change in its physical properties when exposed to a radiation source.

      If we selectively expose a photosensitive material to radiation (e.g. by masking some of the radiation) the pattern of the radiation on the material is transferred to the material exposed, as the properties of the exposed and unexposed regions differs

      • Important in this technique is

        Mask

        Radiation

        Developer solution for etching.

        Fig.4 Process of lithography

        VII.LIMITATIONS:

        • Different photoresists exhibit different sensitivities to different wavelengths. The dose required per unit volume of photoresist for good pattern transfer is somewhat constant.

      • However, the physics of the exposure process may affect the dos actually received. For example a highly reflective layer under the photoresist may result in the material experiencing a higher dose than if the underlying layer is absorptive, as the photoresist is exposed both by the incident radiation as well as the reflected radiation. The dose will also vary with resist thickness.

      • At the edges of pattern light is scattered and diffracted, so if an image is overexposed, the dose received by photoresist at the edge that shouldn't be exposed may become significant. If we are using positive photoresist, this will result in the photoresist image being eroded along the edges, resulting in a decrease in feature size and a loss of sharpness or corners.

      • If we are using a negative resist, the photoresist image is dilated, causing the features to be larger than desired, again accompanied by a loss of sharpness of corners. If an image is severely underexposed, the pattern may not be transferred at all, and in less sever cases the results will be similar to those for overexposure with the results reversed for the different polarities of resist.

      Fig.5 The difference in the exposed pattern

      1. TESTING:

        The experimental setup for the firing of the micro monopropellant thruster and thrust measurement is illustrated in figure . The thrust measurement stand with a force sensor, propellant feed system with a motor driven syringe pump and imaging devices were installed for the firing test of the microthruster.The flow of the propellent can be measured by using flometer. To avoid spontaneous decomposition of hydrogen peroxide, the propellant can be chilled to 5 °C before use. The length of tubing for monopropellant supply can be minimized to reduce the delay of the thrust action. We used a 1/32 in. Teflon tube that will compatible with hydrogen peroxide. A check valve can be used on the tubing to prevent possible backflow o f propellant due to unexpected pressure build- up in the thrust chamber or clogging in the feed line. Thrust can be measured by a Kistler model 9205 load cell along with a Kistler 5015 A charge meter. The transient voltage signal that are proportional to the thrust was transferred to a computer via an NI data system. The

        firing of the thruster was recorded by a camcorder or a camera.

        Fig.6 The experimental setup for the firing of the micro monopropellant thruster and thrust measurement.

      2. REFERENCE:

[1] Jeongmoo Huh and Sejin Kwon, Design, fabrication and thrust measurement of a micro liquid monopropellant thruster, Korea Advanced Institute of Science and Technology.

[2] Siegfried W. Janson, Henry Helvajian, William W. Hansen and Lt. John Lodmell, Microthrusters for Nanosatellites, Center for Microtechnology, Technology Operations, The Aerospace Corporation.

[3] Pijus Kundu, T. K. Bhattacharyya, S. Das A Monopropellant Hydrazine MEMS Thruster for Attitude Control of Nanosatellites Indian Institute of Technology, Kharagpur.

[4] David Hinkley and Brian Hardy , Picosatellites and Nanosatellites at The Aerospace Corporation , In-Space Non- Destructive Inspection Technology Workshop ,Feb 29 Mar 1, 2012 ,Johnson Space Center .

[5] Sungyong An and Sejin Kwon, Scaling and Evaluation of Pt/Al2O3 Catalytic Reactor for Hydrogen Peroxide Monopropellant Thruster, Journal of Propulsion and Power

,Vol. 25, No. 5, SeptemberOctober 2009.

[6] F. Protza ,Machining of carbon and glass fibre reinforced compositers,The Korean Society for Aeronautical and Space Sciences 41(4), 299-304(2013).

[7] Pijus Kundu, Tarun Kanti Bhattacharyya and Soumen Das, Design, fabrication and performance evaluation of a vaporizing liquid microthruster, Journal of Micromechanics and Microengineering, IOP publications.

[8] Angelo Cervone*, Lucio Torre, Luca d,Agostino,Development of Hydrogen Peroxide Monopropellant Rockets, ALTA S.p.A. – Via Gherardesca, 5 – 56121 Ospedaletto, Pisa, Italy

[9] George.P.Sutton,Rocket propulsion Elements,Oscar Biblarz,Wiley India.

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