Propulsion and Energy Systems. Kimiya KOMURASAKI, Professor, Dept. Aeronautics & Astronautics, The University of Tokyo

Transcription

1 Propulsion and Energy Systems Kimiya KOMURASAKI, Professor, Dept. Aeronautics & Astronautics, The University of Tokyo

2 Schedule Space propulsion with non-chemical technologies 10/5 1) Space Propulsion Fundamentals 10/19 2) Electric propulsion Overview & Hall Thruster 10/26 3) Beamed Energy Propulsion Overview 11/2 4) Beamed Energy Propulsion: Microwave Rocket 11/9 5) Air-Breathing Electric Propulsion by Dr. Schoenherr 11/16 6) Nuclear Thermal Rocket 12/7 7) Powered Flight in Planetary Exploration by Prof. Koizumi 12/14 8) Sailing Propulsion by Prof. Funaki(JAXA) 12/21 9) Microsatellite propulsion by Prof. Koizumi 1/25 Reserve

3 Date and Place Date / Time Monday 15:00-16:30 Lecture Rooms UT-Kashiwa campus Kiban-Bldg., Room 2D8 UT-Hongo campus Eng. Bldg. No.7, Room 72

4 Contact etc. Download materials (Slides & report format) Rating by report submission after each lecture

5 Space Propulsion Fundamentals

6 Important parameters Input 1. Mission velocity increment, ΔV 2. Engine exhaust velocity, V e (average) 3. Structure mass, m inert (mainly tank) Output Payload ratio, m payload /m initial

7 1.1 Mission ΔV (1) Propulsive Energy Propulsive Energy = Orbital Energy + Potential Energy = mv orbit2 /2 (μm/r μm/r 0 ) (horizontal) (vertical) Earth radius: r 0 =6,378km To Low Earth Orbit, required ΔV is approximately, ΔV = v orbit + ΔV grav 7

8 1.1 Mission ΔV (2) Orbital velocity mv r 2 G Mm 2 r GM vorbit r r v: orbital velocity r: circular orbit radius Gravitational parameter: μ=gm=398,600 km 3 s -2 Table 1 Orbital velocity Low Earth Orbit (LEO) at h=170 km Geosynchronous orbit (GEO) r=42,164 km 7.8 km/s km/s 8

9 1.1 Mission ΔV (3) Budget to LEO Table 2 necessary velocities in m/s Vehicle Ariane A-44L Orbit h p x h a Inclination (deg) 170x V LEO ΔV grav ΔV steering ΔV drag ΔV rot ΔV total =ΣΔV Saturn 176 x Titan IV 157 x

10 1.1 Hohmann transfer from LEO to GEO V p r r 2r 2 r apoapsis periapsis V a r r 2r 1 r Perigee Kick (launcher) Vp Vp V km/s Fig. 1 Orbit Transfer between two nonintersecting orbits at the apsides) Apogee Kick (Satellite s motor) Va V2 Va km/s V 1: circular orbit velocity at h=170 km, LEO V 2 : circular orbit velocity at the geosynchronous orbit 10

11 1.2 What is Rocket Propulsion? Momentum Conservation d P system dt 0 d ( d ) d m m V V m V V mv e (1) (2) 11

12 Rocket Equation (Tsiolkovsky rocket equation) Integrating Eq.(2), we have i V Ve ln mf ΔV : velocity increment V e : effective exhaust velocity m f : final mass m i : initial mass m Konstantin Eduardovich Tsiolkovsky (3) 12

13 Effective exhaust velocity, V e Table 3 Exhaust velocity of bipropellant rockets 1 st stage engine propellants Thrust (ton) Exhaust (m/s) Space Shuttle (Main) LOX/LH Energia (RD-0120) LOX/LH H-II (LE-7) LOX/LH Ariane-V (Vulcan) LOX/LH Energia (Booster) LOX/RP Saturn V LOX/RP

14 1.3 Performance limit of single stage launcher(1) Inert mass (for structure and tanks) Inert mass fraction m m m m m m f inert inert i pay prop f pay m inert m inert m prop (4) (5) Substitute Eq.(5) to Rocket Eq. (3), mprop V 1 finert exp 1 mpay Ve 1 finert exp V Ve (6) 14

15 1.3 Performance limit of single stage launcher (2) Necessary condition to avoid division-by-0 in Eq. (6) inert 1 f exp V V 0 e ie. V e V ln 1 f inert V ln 1 f V 2.54 km/s 10 km/s inert e for f inert = 0.08 and V e =4 km/s 15

16 Question With exhaust velocity V e = 4 km/s and required ΔV = 8 km/s, how much propellant is required for the launch of 10 ton payload? Let e 2 =7.4 and f inert =

17 Launch rockets in the world Rocket Japan H-IIA (2 SRB) (4 SRB) H-IIB Launch mass (ton) Payload mass (ton) LEO (altitude, km/ inclination) 10 (300/30.4) 15 (300/30.4) 19 (300/30.4) GTO First launch Boost ers solid solid Propellant 1st stage 2 nd 3rd & 4th LOX/LH 2 LOX/LH 2 LOX/LH 2 LOX/LH 2 USA Falcon 9 Antares (400/28.7) LOX/RP-1 LOX/RP-1 LOX/RP-1 Solid ESA Arian V (ES) (ECA) VEGA (300/5.2) solid LOX/LH 2 Solid N 2 O 4 /CH 6 N 2 LOX/LH 2 Solid Solid & N 2 O 4 / UDMH Russia Zenit 3SL Proton M Soyuz (-/51.6) LOX/ RP-1 LOX/RP-1 N 2 O 4 / UDMH LOX/RP-1 LOX/RP-1 N 2 O 4 / UDMH LOX/RP-1 LOX/RP-1 N 2 O 4 / UDMH N 2 O 4 / UDMH China Long March N 2 O 4 / UDMH N 2 O 4 / N 2 O 4 / UDMH UDMH 17

18 2. Exhaust velocity in Chemical Rockets

19 2.1 Chemical Rocket Engine Cycle (Why is V e limited?) i c t e Isobaric heating + isentropic expansion Fig. 3 Schematic of a chemical rocket p i t h p c pt t e p e v Fig. 4 p-v diagram (left) and h-s diagram (right) s 19

20 2.2 Characteristic exhaust velocity & thrust coefficient Exhaust velocity V e C C F * R T 1 c ep 1 M p c 1 2 (7) (8) C F : Thrust coefficient : nozzle performance C*: Characteristic exhaust velocity, m/s 20

21 Characteristic exhaust velocity C*(m/s) (performance without nozzle) C A P m 1 R T 2 1 c 2 M 1 * t c (11) (12) acoustic velocity A function of T c /M 21

22 Thrust coefficient : nozzle performance C F F A P t c (9) p e p c 1 2 (10) Fig.5 Thrust coefficient as a function of nozzle expansion ratio. 22

23 2.3 Chamber Temperature T c (adiabatic flame temperature) H 2 O molar heat of formation 1 Q1 H H H 2 O H 2 H2 O 2 242kJ/mol (13) Increment of internal energy is T Q C T f d 2 T p,h 2 O 0 (14) Taking a balance, Q 1 =Q 2, T c is determined 23

24 2.4 Combustion reactions 1 H2 O2 H2 O 2 1 O O H H H 1 2 O2 OH 2 2 Table 4 Mixture Ratio and T c, M. (10MPa) MR H 2 O mole fraction T c, K M T c /M 4H 2 +O H 2 +O H 2 +O

25 2.4 Optimum LOX/LH2 mixture ratio optimum condition Highest V e ΔV (m/s) Ve (m/s) ΔV propellant density in tanks V e Propellant density (g/cm 3 ) Stoichiometric MR Propellant mixture ratio (m o /m f ) 0.2 Computed performance for LOX/LH2 engines 25

26 3 Propellant Feed System

27 Rocket engine cycles (Auxiliary combustion chamber necessary to drive turbo pumps) Gas-generator cycle Staged combustion cycle 27

28 LE-7 engine cycle (H-II s first stage ) Staged combustion cycle 28

29 RL10 turbopump(for upper stages of Atlas -5 and Delta -4) LOX pump turbine H2 pump Layout of RL10 engine single shaft TPA turbine impeller inducer AIAA Single Shaft Turbopump Expands Capabilities of Upper Stage Liquid Propulsion 29

30 LE-7 engine failure 1999 H-II rocket No. 8 was failed due to engine trouble. The failed engine was retrived from the seabed about 3000 m deep and its non-destructive and destructive tests were performed. Retrieval of the failed engine. It is said The fuel turbopump had problem in inducer (a propeller-like axial pump used to raise the inlet pressure of the propellant ahead of the main turbopumps to prevent cavitation) where cavitation caused an imbalance resulting in excessive vibration and led to premature engine failure. 30

31 4. Reusable vehicle launched as a conventional and returned as a glider 31

32 5. What is alternative? 10/19 Electric propulsion Overview & Hall Thruster 10/26 Beamed Energy Propulsion Overview 11/2 Beamed Energy Propulsion: Microwave Rocket 11/9 Air-Breathing Electric Propulsion 11/16 Nuclear Thermal Rocket 12/7 Powered Flight in Planetary Exploration 12/14 Sailing Propulsion 12/21 Microsatellite propulsion 32

33 Report at each lecture Evaluate Technical Readiness Level of Alternative Propulsion Systems presented at each lecture. Deadline 17:00 PM, Wednesday. Submit to

34 Evaluation of Technical Readiness Technology Readiness Level (TRL) is NASA s measure to assess the maturity of evolving prior to incorporating that technology into a system. Proposed by Dr. John Mankins 34

35 TRL Basic principles observed and reported This is the lowest "level" of technology maturation. At this level, scientific research begins to be translated into applied research and development. 2. Technology concept and/or application formulated 3. Analytical and experimental critical function and/or characteristic proof of concept Once basic physical principles are observed, then at the next level of maturation, practical applications of those characteristics can be 'invented' or identified. At this level, the application is still speculative: there is not experimental proof or detailed analysis to support the conjecture. At this step in the maturation process, active research and development (R&D) is initiated. This must include both analytical studies to set the technology into an appropriate context and laboratory-based studies to physically validate that the analytical predictions are correct. These studies and experiments should constitute "proof-of-concept" validation of the applications/concepts formulated at TRL 2. 35

36 TRL Component and/or breadboard validation in laboratory environment 5. Component and/or breadboard validation in relevant environment Following successful "proof-of-concept" work, basic technological elements must be integrated to establish that the "pieces" will work together to achieve concept-enabling levels of performance for a component and/or breadboard. This validation must be devised to support the concept that was formulated earlier, and should also be consistent with the requirements of potential system applications. The validation is "low-fidelity" compared to the eventual system: it could be composed of ad hoc discrete components in a laboratory. At this level, the fidelity of the component and/or breadboard being tested has to increase significantly. The basic technological elements must be integrated with reasonably realistic supporting elements so that the total applications (component-level, sub-system level, or system-level) can be tested in a 'simulated' or somewhat realistic environment. 6. System/subsystem model or prototype demonstration in a relevant environment (ground or space) A major step in the level of fidelity of the technology demonstration follows the completion of TRL 5. At TRL 6, a representative model or prototype system or system - which would go well beyond ad hoc, 'patch-cord' or discrete component level breadboarding - would be tested in a relevant environment. At this level, if the only 'relevant environment' is the environment of space, then the model/prototype must be demonstrated in space. 36

37 TRL System prototype demonstration in a space environment 8. Actual system completed and 'flight qualified' through test and demonstration (ground or space) 9. Actual system 'flight proven' through successful mission operations TRL 7 is a significant step beyond TRL 6, requiring an actual system prototype demonstration in a space environment. The prototype should be near or at the scale of the planned operational system and the demonstration must take place in space. In almost all cases, this level is the end of true 'system development' for most technology elements. This might include integration of new technology into an existing system. In almost all cases, the end of last 'bug fixing' aspects of true 'system development'. This might include integration of new technology into an existing system. This TRL does not include planned product improvement of ongoing or reusable systems. 37