PPT, Electrospray, μCAT and other Low-Thrust Systems

[NirViaje]

Pulsed Plasma Thruster (PPT) and other Low-Thrust Systems

• https://en.wikipedia.org/wiki/Pulsed_plasma_thruster

the PPT on Earth Observing-1 mission (EO-1)

Note 1: Efficiency is defined as the thruster efficiency, PPU efficiency not included Note 2: EO-1 mass driven by solar array torque environment Note 3: EO-1 was fuel limited to 1,500 N-sec, however, key components demonstrated life to 20 million pulses (20x10^6 * 750mN-sec = 15,000N-sec)

The main features of the EO-1 PPT system are as follows:

• Two electrode pairs, each with a 9.65 square cm fuel face, opposing each other on one capacitor
• A single fuel feed housing that is shared between the two fuel bars
• A total impulse capability of 1560 N-sec (limited by available fuel)
• Impulse bit dynamic range is <100 µN-sec -900 µN-sec
• 40 µF main capacitor, with a demonstrated life of over 25 million pulses at 43 joules
• Conical mounting bracket that supports the PPT cluster at a 30 degree angle with the outside of the spacecraft
• Operates on a 28 V power bus (22-34 V unregulated)
• One power connector, one command/telemetry connector

• A Micro Pulsed Plasma Thruster (PPT) for the “Dawgstar” Spacecraft, Primex Aerospace Company, 2000
  • PPT Development Efforts at Primex Aerospace Company, W. Andrew Hoskins, et. al., Primex Aerospace Company, 1999
  • On-Orbit Testing of the EO-1 Pulsed Plasma Thruster, nasa, 2003
  • https://en.wikipedia.org/wiki/Earth_Observing-1
  • https://eo1.gsfc.nasa.gov/new/validationReport/index.html

• Liquid PPT for attitude and orbit control of space vehicles, 2017

• Design of a High-energy, Two-stage Pulsed Plasma Thruster, Alfven, Princeton, 2002

Figure 2 shows a typical signal of displacement in measurement of impulse bit. Sensitiveness of the thrust stand is variable by changing the weight mounted on the top of the pendulum as shown in Fig.3. A calibration of the thrust stand is carried out by collisions of balls to the pendulum with various balls from various distances corresponding 15-1400 μNs.

• Electric Rocket Engine System R&D, Osaka Institute of Technology, on PROITERES, 2012
  • Development of Electrothermal Pulsed Plasma Thruster System Flight-Model for the PROITERES Satellite, 2011
  • R&D, Launch and Initial Operation of the Osaka Institute of Technology 1st PROITERES Nano-satellite with Electrothermal Pulsed Plasma Thrusters and Development of the 2nd Satellite, 2013

Figures 11, 12, 13, and 14 show the experimental results. As a result of experiments, the impulse bit showed the maximum value when the discharge room length was 50 mm. The impulse bit was 2.47 mNs, the mass shot was 738μg, the specific impulse was 342 s, the thrust efficiency was 13.1%. It was confirmed that the impulse bit becomes larger as the discharge room length becomes longer, and the specific impulse becomes smaller. Figure 16 is shown that unexpected discharged were induced in some discharge rooms in which no firing was selected. The following two causes were inferred: 1) The plasma which was generated in the 1st MDR-PPT leaked out from the gap between the propellant and the nozzle or the anode. 2) Excessive ablation of propellant occurred by repetitive operation under high-temperature conditions lead to a short between the discharge rooms.

VIII. Conclusions

The following results were obtained:

1. A high-power PPT head showed initial performance of impulse bit at 2.47 mNs, the mass shot was 738 μg, the specific impulse was 342 s, the thrust efficiency was 13.1% at 31.59 J per shot with a discharge room length of 50 mm.
2. A new PPT system with the 1st MDR-PPT for long-time operation was developed. However, with the 1st MDR-PPT unexpected discharges occurred in the discharge rooms.
3. The induced problem that occurred in the 1st MDR-PPT was solved by the 2nd MDR-PPT design.
4. The 3rd MDR-PPT which was designed considering the drawbacks of 2nd MDR-PPT.
5. The discharge chamber switching experiment using PPU and two PPTs was conducted and the ignition discharge was confirmed in both the atmosphere and the vacuum environment in each selected discharge chamber.
6. A total impulse of 92.0 Ns was obtained in a repetitive 110,000 shot operation with a discharge room length of 50 mm of a single PPT head.
7. Design change was successfully made to make the 3rd MDR-PPT a PPT system for mounting on a nano-satellite.
   • Research and Development of Electrothermal Pulsed Plasma Thruster Systems for Powered Flight onboard the Osaka Institute of Technology 2nd PROITERES Nano-Satellite, 2017

Further reference

• An Overview of Cube-Satellite Propulsion, Wichita StateU, 2017
• An Exploration of CubeSat Propulsion Andrew Davis Hine Western Michigan University, 2016
• Small Satellite Propulsion, Arizona State University, AstroRecon, 2015
• Related conference https://github.com/ExponentialDeepSpace/exponentialdeepspace.github.io/issues/9#issuecomment-445451870

[NirViaje]

Electrospray and others

• TILE-V1 Data Sheet

• State-of-the-Art for Small Satellite Propulsion Systems, Khary I. Parker, NASA/Goddard, 2017

[NirViaje]

Micro-Cathode Arc Thruster, μCAT

• Agile development process of flight hardware for a quad-channel Micro-Cathode Arc Thruster (μCAT) subsystem for the 1.5U BRICSat-P cubesat mission, GWU, 2014

• Micro-Cathode Arc Thruster for Nanosatellite Propuslion, GWU, 2013

• Performance Characteristics of Microcathode Arc Thruster, GWU, 2013

• Micro-Cathode Arc Thruster for PhoneSat Propulsion, GWU, 2013

• Micro-Cathode Arc Thrusters for CubeSat Propulsion, GWU, 2018
• High thrust-to-power ratio micro-cathode arc thruster, GWU, 2016
• Micropropulsion and Nanotechnology Laboratory, MpNL@GWU

• https://en.wikipedia.org/wiki/BRICSat-P
  • BRICSat-P, 2015
  • Communication and Data Handling System for BRICsat Satellite, Czech, 2012
  • BRICSat-P (Ballistically Reinforced Communication Satellite)
• 微阴极电弧推力器研究进展, BUAA, 2017 HTML

[NirViaje]

VAT, Vacuum Arc Thruster

VAT - with magnetic -> μCAT

• Miniature Vacuum Arc Thruster with Controlled Cathode Feeding, Isreal, 2017
• Vacuum Arc Thruster development and testing for micro- and nanosatellites, Japan, 2015
• Vacuum Arc thruster development for Horyu-4 satellite, KIT, JAXA, 2013

[NirViaje]

MPCS-IV

The forth International Conferences on Micropropulsion and CubeSats May 9-11, 2019 Lab of Space Propulsion Technology

Beijing Institute of Control Engineering, Beijing, China

• https://www.micropropulsion.org/mpcs-4

[NirViaje]

• https://www.comat-agora.com/space/

The Science Behind the Neumann Drive

• http://neumannspace.com/science/

“Fuelling the Future”

We’ve built a brand new kind of ion engine (that’s a kind of rocket), that has just broken the world record for specific impulse previously held by NASA’s HIPEP thruster.

This level of fuel efficiency is so good that one of these engines could send a probe to Mars and back on a single fuel rod. But it could also be great for keeping satellites in their proper position in orbit, or cheaply sending all the heavy equipment ahead of a manned mission somewhere.

How does it work?

The Neumann Drive uses solid fuel and electricity to produce thrust. It is a “wire-triggered pulsed cathodic arc system” and works kind of like an arc welder.

Arc welders have a cathode (the welding rod, charged negative) and an anode (the work piece, which has the anode-lead clamped to it). When the tip of the welding rod gets close enough to the work piece, an arc of electricity sparks between them. This happens because the electric field between the cathode and anode is strong enough to rip electrons off the air molecules between them and causes a giant “spark” to jump. The arc allows electrical current to flow through the cathode, which heats the material on the tip of the welding rod.

As the electrons jump off the end of the rod and enter the arc, they carry along with them some atoms from the rod in the form of plasma. In an Arc Welder, these iron and carbon atoms then get deposited on the work piece at high energy, creating a small melt pool and the desired weld. In the Neumann Drive, these atoms will be hurled off into space, producing thrust in the drive itself.

In our system the cathode is a cylindrical rod of conducting material that we choose to use as fuel (eg magnesium, vanadium etc). The cathode is charged negative with respect to the anode, and a charging voltage of between 80 and 250V is typically used. The anode is a hollow cylinder that is aligned coaxially with the cathode, but offset slightly forwards, as shown in the diagram below.

The cathode rod has a hole bored down the centre of it, which holds the insulated trigger pin. It needs to be insulated so that the arc can be triggered only at the right time. Our system uses an electrical flashover system to trigger the arc, which means that we pulse a high voltage signal from the trigger pin to the cathode, creating the conditions necessary for plasma formation to occur. There are other methods for triggering an arc (including lasers), but they are not as robust as this.

Once the arc has been triggered, plasma will be created in very small and very bright spots located on the cathode surface close to the trigger location; these plasma generation sites are called “cathode spots” for obvious reasons. The cathode spots erode material from the cathode, ionising and accelerating it into the vacuum chamber so that the plasma moves downstream at high velocity through the anode mouth.

This so-called “drifting plasma” is the exhaust of our rocket, and pushes the rest of the system forwards as it hurtles away. Higher exhaust velocities mean more efficient fuel use which is measured in specific impulse, often called “bounce per ounce.”

What fuels does it use?

We’ve tried out all kinds of materials as fuel. We’ll be posting a showcase about several of them over the coming few weeks, and I’ll add links to these posts as they appear. The list below contains a non-exhaustive list of the things that we’ve tried:

• Molybdenum – our fastest fuel – best for sending people to Mars
• Magnesium – our most efficient fuel – best for sending equipment on long missions
• Aluminium – best for recycled space junk
• Carbon – our most interesting fuel – reusing ahem waste-products from astronauts to get them to where they’re going
• Titanium
• Vanadium
• Tin – a pretty poor fuel that we learned a lot from
• Bismuth – our most useless fuel… just don’t bother

Footnotes and details

aligned coaxially: this means that the central axis of the fuel rod is aligned with the central axis of the cylinder

other trigger methods: Other triggering methods include mechanical methods (similar to the welding rod analogy above) or laser triggering (blast some material off the cathode surface, and use this to create a runaway discharge), but neither of these are as robust as an electrical system.

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• http://neumannspace.com/blog/fuel-of-the-week-tin/

[NirViaje]

Busek RF Ion Thruster

• http://www.busek.com/technologies__ion.htm
• http://www.busek.com/index_htm_files/70011950%20RevA%20Data%20Sheet%20for%20BIT-1%20Ion%20Thruster.pdf
• https://www.nasa.gov/sites/default/files/atoms/files/vlad_busek_iodine_rf_ion_thruster_tagged.pdf
• https://icubesat.files.wordpress.com/2012/06/icubesat-org-2012-c-2-3-_presentation_hruby_201205291229.pdf
• http://mstl.atl.calpoly.edu/~workshop/archive/2012/Summer/Day%202/1545-Williams-PropulsionSolutions.pdf

Controller

• http://comfiletech.com/cb320/
• http://comfilewiki.co.kr/en/doku.php?id=cubloc:index

Enpulsion IFM Nano

• https://www.enpulsion.com/order/ifm-nano-thruster/
  • https://www.enpulsion.com/uploads/products/IFM-Nano-Thruster/ENP_-_IFM_Nano_Thruster_-_Product_Overview.pdf
  • ENP2019-086.B-IFM-Nano-Thruster-COTS-Product-Overview-1.pdf
  • ENP-IFM-Nano-Thruster-Product-Overview.pdf
  • The-IFM-350-Nano-Thruster-Introducing-very-high-delta-v-Capabilities-for-Nanosats-and-Cubesats_v2.pdf
• https://www.cubesatshop.com/product/ifm-nano-thruster/, for 30000 euro
  • https://www.cubesatshop.com/product/cluster-ifm-nano-thruster-smallsats/

[NirViaje]

ThrustMe公司在6月16日发布消息称，该公司的电推进系统主要采取了两项创新措施，使其尺寸仅为传统电推进系统的40%，但推力可提高一倍。第一，采用固态碘作为工质，易于形成高压气体，减少了传统气态工质所必需的高压贮箱和复杂的流量控制装置。第二，加速栅格和射频电压栅格并排设置，使等离子体形成“二极管”效应，直流电压可在持续加速阳性离子束喷出的同时，周期性地将电子喷出，实现中和器的效果，大幅简化了推进系统结构设计。（中国航天系统科学与工程研究院 陈建光）

• http://thrustme.fr/
  • https://mp.weixin.qq.com/s/iUwR7DAmPjoY4YJ2ZWEGvg

[NirViaje]

Water plasma propulsion, Vigoride from Momentus, CTO Joel Sercel

• http://fiso.spiritastro.net/telecon/Sercel_12-12-18/Sercel_12-12-18.pdf
  • Space Transportation Services- A Breakthrough Business Model Enabled By Radical Silicon Valley Innovation, Joel C Sercel, 2018_Sercel_12-12-18.pdf
• https://momentus.space/docs/Momentus-Datasheet-Vigoride.pdf

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RCS, Arcjet

• electrothermal_thrusters.pdf

[NirViaje]

• Recent progress and perspectives of space electric propulsion systems based on smart nanomaterials, NTU, 2018_s41467-017-02269-7.pdf

[NirViaje]

• https://www.mouser.com/catalog/supplier/library/pdf/TamuraCatalog.pdf
• https://www.tamuracorp.com/products/electronic_components/category/piezoelectric_ceramic/piezoelectric_ceramic.html

[NirViaje]

APSCO-SSS-1卫星由7个分系统组成。为避免太空垃圾，离轨分系统采用了由北航宇航学院研制的电喷雾推力器(ILT-50)，该推力器是国内首台由高校研制并执行在轨飞行任务的电喷雾推进装置，也是国际上首次容性储电实现单台电喷雾推力器自中和工作并应用于卫星在轨任务。电喷雾推进研制团队由宇航推进系汤海滨教授、杨立军教授、刘宇教授、富庆飞教授、任军学副教授、杨文将副教授、王一白副教授等及二十余名博士、硕士和本科生组成

在ILT-50电喷雾推进装置研制过程中，项目团队针对泰勒锥的形成、锥射流破裂雾化及带电粒子在电场中的运动等电喷雾基本过程开展了基础理论研究，提出了非定常电场作用下的电雾化动力学理论，发现了交变电压激励下锥射流的多种脉动新模式，指导了电推力器的自中和设计；基于APSCO-SSS-1卫星离轨任务和电推力器高比冲的需求，研发配制了特殊的电喷雾推进剂，应用了新型非金属发射材料，开展了电喷雾推力器设计和试验，突破了推力器自中和、推力器/电源集成和微牛级推力精确测量等关键技术，电推力器比冲达到1400s，在国际上首次通过容性储电方式实现了单台电喷雾推力器的自中和工作并应用于卫星在轨任务。相关研究成果学术论文发表在流体力学顶级期刊《Journal of Fluid Mechanics》、《Physics of Fluids》、《Journal of Aerosol Science》等高水平期刊上：

1. Yang Li-jun, Wang Chen, Fu Qing-fei. Weakly nonlinear instability of planar viscous sheets. JOURNAL OF FLUID MECHANICS, 2013, 735(1). [2. Hai-Bin Tang, Chao-Jin Qin, Yu Liu.Characterization of colloid thruster beams and plumes,Journal of Aerosol Science 42 (2011) 114–126
2. Xie Luo, Jia Bo-qi, Cui Xiao, Yang Li-jun, and Fu Qing-fei. Linear instability analysis of a non-Newtonian liquid jet under AC electric fields. Aerospace Science and Technology, 2020: 106121.
3. Xie Luo, Yang Li-jun, Ye Han-yu. Instability of gas-surrounded Rayleigh viscous jets: Weakly nonlinear analysis and numerical simulation. Physics of Fluids 2017, 29, 074101
4. Xie Luo, Jia Bo-qi, Cui Xiao, Yang Li-jun, and Fu Qing-fei. Linear analysis and energy budget of viscous liquid jets in both axial and radial electric fields. Applied Mathematical Modelling 2020, 83, 400-418
5. Zhongkai ZHANG, Guanrong HANG, Jiayun QI,Zun ZHANG, Zhe ZHANG, Jiubin LIU,Wenjiang YANG and Haibin TANG.Design and fabrication of a full elastic sub-micron-Newton scale thrust measurement system for plasma micro thrusters,Plasma Sci. Technol. 23 (2021) 104004
6. Fawzi DERKAOUI, Zhaoxin LIU, Wenjiang YANG, Yu QIN, Kunlong WU, Peng ZHAO, Juzhuang YAN , Junxue REN and Haibin TANG.Design and research of magnetically levitated testbed with composite superconductor bearing for micro thrust measurement,Plasma Sci. Technol. 23(2021) 104010
7. 王群，富庆飞. 交变电场作用下单液滴蒸发的分子动力学模拟，力学学报，2021， 53（5）：1-10
8. 富庆飞，葛斐，成锦博，杨立军，任军学，汤海滨. 非定常电场下锥射流的振荡行为研究, 实验流体力学，2021，在线出版

• http://www.sa.buaa.edu.cn/info/1050/7791.htm

西北工业大学 陈茂林

• https://teacher.nwpu.edu.cn/chenmaolin.html

[NirViaje]

• Additively_manufactured_electrohydrodynamicionic.pdf

MIT IEPS

• https://spacepropulsion.mit.edu/electrospray-thruster-engineering

Reference

1. D. Krejci, F. Mier-Hicks, R. Thomas, T. Haag, P. Lozano "Emission Characteristics of Passively Fed Electrospray Microthrusters with Propellant Reservoirs," Journal of Spacecraft and Rockets, Vol. 54, No. 2, 2017, pp. 447-458.
2. D. Krejci, F. Mier-Hicks, C. Fucetola, P. Lozano, A. Hsu Schouten and F. Martel, "Design and characterization of a scalable ion electrospray propulsion system," IEPC-2015-149 34th INTERNATIONAL ELECTRIC PROPULSION CONFERENCE, Kobe, Japan.
3. D. Krejci and P. Lozano, "Current Capabilities of Scalable Ionic Liquid Electrospray Thrusters for Nano-Satellites," 39th AAS Guidance & Control Conference, Breckenridge, CO.
4. C. Guerra-Garcia, D. Krejci and P. Lozano, "Spatial uniformity of the current emitted by an array of passively fed electrospray porous emitters," Journal of Physics D: Applied Physics, Vol. 49, No. 11, 115503 (12pp).

[NirViaje]

Research on ultra-light high Isp thruster system

1. Objective

• Thrust
  • 100~500uN
• Isp
  • 1000~6000s or more
• Total Impulse
  • 1000~3000Ns
• Life span
  • 1000~3000 hours of continual thrusting
• Power consuming
  • < 100W
• Weight
  • < 100g (thruster)
  • < 200g (system)

2. Select the suitable principle of operation

• Meet the objective
• Potential to be lightweight
• Easy to integrate with film system

3. Understanding the physical limitation of the thruster system

the limitation of these four:

• Thrust
• Isp
• Total Impulse
• Weight

4. Reference

• Accion Systems | TILE-V1
• ENP-IFM
  • https://www.enpulsion.com/wp-content/uploads/The-IFM-350-Nano-Thruster-Introducing-very-high-delta-v-Capabilities-for-Nanosats-and-Cubesats_v2.pdf
  • https://www.enpulsion.com/wp-content/uploads/ENP2019-086.B-IFM-Nano-Thruster-COTS-Product-Overview-1.pdf
  • https://www.cubesatshop.com/wp-content/uploads/2017/04/ENP-IFM-Nano-Thruster-Product-Overview.pdf
• Máximo D V M, Velásquez-García L F. Additively manufactured electrohydrodynamic ionic liquid pure-ion sources for nanosatellite propulsion[J]. Additive Manufacturing, 2020, 36: 101719.

[NirViaje]

Characterization of Vacuum Arc Thruster Performance in Weak Magnetic Nozzle

• Characterization of Vacuum Arc Thruster Performance in Weak Magnetic Nozzle.pdf

[NirViaje]

• Cal Poly Satellite Positioning Systems_ Thrust or Bust!.pdf