1. Principles of Propulsion and its Application in Space Launchers Prof. Dr.-Ing. Uwe Apel Hochschule Bremen 13.07.2012REVA Seminar1

2. Overview How Rockets are Propelled Thrust Generation in a Rocket Engine Rocket Engine Performance Parameters Classification of Space Vehicles Application of Rocket Engines Classification of Rocket Propulsion Systems Physical Limits of Chemical Space Propulsion The Rocket Equation Staging of a Rocket 13.07.2012REVA Seminar2

3. How Rockets are Propelled The Change of the state of motion of a rocket follows the principle of repulsion Newton‘s law applies: ACTIO = REACTIO Any force acting on a mass creates an force of the same size in the opposite direction! By ejection of a mass at a high velocity (usually a hot gas flow ) from the rocket engine a force is produced changing the momentum of the rocket. Important: According to Newton‘ law of momentum conservation the sum of the momentum changes of working fluid and vehicle equals 0 ! 13.07.20123REVA Seminar

4. Functional Principle of a Rocket Thrust is generated exits nozzle with velocity 13.07.20124REVA Seminar

5. Thrust Generation in a Rocket Engine 13.07.20125REVA Seminar

6. Rocket Engine Performance Parameters 13.07.20126REVA Seminar

7. The Rocket Equation Describes Movement of a rocket in force-free space Calculates velocity change achievable with a rocket geaturing a certain mass ratio and average specific Impulse Differential form: Integral form: 13.07.2012REVA Seminar7

8. Classification of Space Vehicles 13.07.20128REVA Seminar

9. Classification of Rocket Propulsion Systems Origin of propulsion energy –Chemical –Nuclear –Solar Propellants and their aggregate state –Solid propellants –Liquid propellants –Hybrid engines –Cold gases Thrust level –High thrust (> engine weight) –Low thrust (< engine weight) 13.07.20129REVA Seminar

10. Application of Rocket Engines 13.07.201210REVA Seminar

11. Typical Performances of Rocket Engines 13.07.201211REVA Seminar

12. Rocket Engine Performance Map thrust to mass [N/kg] acceleration [m/s] specific impulse [m/s] 13.07.201212REVA Seminar

13. ∆V Requirement The ∆V requirement of a space mission is dependent on: –Size and orbit of launch planet –Size and orbit of destination planet –Propulsion concept (thrust level, propulsion time) –Chosen trajectory and resulting flight time –Accuracy of orbit and attitude control system –Vehicle aerodynamics 13.07.201213REVA Seminar

14. ∆V Calculation 13.07.2012REVA Seminar14

15. Typical ∆V Requirements 13.07.2012REVA Seminar15

16. Elements of a Space Transportation System 13.07.201216REVA Seminar

17. Elements of a Rocket The take-off mass of a rocket consists of three major mass elements: Structure and Engine(s) –Body and tankage –Engines and related equipment –Non-usable propellant residuals –Usable propellant reserve –Recovery equipment (parachutes, wings, landing gear, etc.) –Instrumentation and avionics Propellants –Expected propellant consumption during flight –Propellants expended prior to lift-off Payload 13.07.201217REVA Seminar

18. Design Parameters According to the rocket equation a maximisation of the ratio between the initial mass m 0 and the cut-off mass m c is required for a high velocity capability Thus 80% ÷ 90% of the initial mass of a rocket is propellant mass This requires an ultra-light structural design and small, efficient engines with a very high power density! Key design parameters of a rocket are: –The propellant mass fraction –The propellant ratio –The payload ratio 13.07.201218REVA Seminar

19. Technological limits for a rocket The performance of a single-stage rocket is limited by the technologically achievable values for the mass ratio R and the exhaust velocity C and the ∆V requirements of the mission: Limits: – useful minimum payload mass fraction of >=1 % – achievable propellant mass fraction of µ= 0.90 – today’s engines performance ofC 0 = 4300 m/s C vac =4600 m/s –minimum velocity increment to reach orbit∆V= 9100 m/s Thus, it is very difficult to design a one-stage launch vehicle! 13.07.201219REVA Seminar

20. High Development Risk! Technological Limits: Single-stage to Orbit (SSTO) m payload m structure m propellant 13.07.201220REVA Seminar

21. Staging of a rocket The problem can be overcome by "staging" the rocket which means distributing the total propellant mass over more than one tank for each propellant component and not further accelerating empty tankage by cutting it off In theory a rocket with an infinite number of stages would provide a maximum payload ratio Practically the number of stages is limited by the propellant mass fraction of each stage which increases with decreasing stage size because tanks and engines cannot be downsized linear For transportation in orbits around Earth, 2-3 stages provide an optimum performance depending on the selected propellant combination and other design aspects 13.07.2012REVA Seminar21

22. Influence of staging on payload mass (example) Assuming a launch vehicle based on following design data: Mission velocity requirement (Earth to orbit):∆V=9200 m/s Average specific Impulse of engines: C=4400 m/s Launch mass: m 0 =100 Mg Propellant mass fraction: µ=0.9 One-stage design Two-stage design 13.07.2012REVA Seminar22

23. Influence of staging on vehicle mass and payload One-stage designTwo-stage design 13.07.2012REVA Seminar23

24. Optimum staging of a launch vehicle Optimum distribution of total ∆V between the stages of a rocket depends of specific impulses of stage engines and stage propellant mass fractions For a two-stage vehicle, the payload mass fraction of the rocket with respect to a given mission ∆V can be obtained from the following equation 13.07.2012REVA Seminar24

25. Optimum staging of a launch vehicle For a rocket with the same average specific impulse and propellant mass fraction in each stage, the  - Function has its maximum at U 1 =U 2 =∆V/2 This means, that the first stage of a two-stage rocket should have a mass which is 3.6 times the mass of the second stage if the same technology is used in both stages For a launch vehicle going from Earth‘s surface to an orbit the described theoretical optimum is additionally influenced by the ascend trajectory due to: –gravity and drag losses (changes theoretical ∆V distribution) –engine performance (C depends on ambient pressure) 13.07.2012REVA Seminar25

26. Optimum staging of a launch vehicle (Example) 13.07.2012REVA Seminar26