5 For a typical launch vehicle headed to an orbit, aerodynamic drag losses are typically quite small, on the order of 100 to 500 m/sec.Gravitational losses are larger, generally ranging from 700 to 1200 m/sec depending on the shape of the trajectory to orbit.By far the largest term is the equation for the space velocity increment.
6 REACHING ORBITThe lowest altitude where a stable orbit can be maintained, is at an altitude of 185 km.This requires an Orbital velocity approximately 7777 m/sec.To reach this velocity from a Space Center, a rocket requires an ideal velocity increment of 9050 m/sec.The velocity due to the rotation of the Earth is approximately 427 m/sec, assuming gravitational plus drag losses of 1700 m/sec.A Hydrogen-Oxygen system with an effective average exhaust velocity (from sealevel to vacuum) of 4000 m/sec would require Mi/ Mf = 9.7.
7 Geostationary orbitA circular geosynchronous orbit in the plane of the Earth's equator has a radius of approximately 42,164 km (26,199 mi) from the center of the Earth.A satellite in such an orbit is at an altitude of approximately 35,786 km (22,236 mi) above mean sea level.It maintains the same position relative to the Earth's surface.If one could see a satellite in geostationary orbit, it would appear to hover at the same point in the sky.Orbital velocity is 11,066 km/hr= 3.07 km/sec (6,876 miles/hr).
9 MULTISTAGE ROCKETSWith current technology and fuels, a single stage rocket to orbit is still not possible.It is necessary to reach orbit using a multistage system where a certain fraction of the vehicle mass is dropped off after use thus allowing the non-payload mass carried to orbit to be as small as possible.The final velocity of an n stage launch system is the sum of the velocity gains from each stage.
10 ANALYSIS OF MULTISTAGE ROCKETS M0i : The total initial mass of the ith stage prior to firing including the payload mass,ie, the mass of i, i+1, i+2, i+3,...., n stages.Mpi : The mass of propellant in the ith stage.
11 Msi : Structural mass of the ith stage alone including the mass of its engine, controllers and instrumentation as well as any residual propellant which is not expended by the end of the burn.ML : The payload massMass ratio
16 MOMENTUM BALANCE FOR A ROCKET Rocket mass X Acceleration = Thrust – Drag -gravity effect
17 EFFECTIVE EXHAUST VELOCITY The total mechanical impulse (total change of omentum) generated by an applied force, FT, is:The total propellant mass expended is
18 The instantaneous change of momentum per unit expenditure of propellant mass defines the effective exhaust velocity.
19 Rocket PrinciplesHigh pressure/temperature/velocity exhaust gases provided through combustion and expansion through nozzle of suitable fuel and oxidiser mixture.A rocket carries both the fuel and oxidiser onboard the vehicle whereas an air-breather engine takes in its oxygen supply from the atmosphere.
20 Criteria of Performance Specific to rockets only.thrustspecific impulsetotal impulseeffective exhaust velocitythrust coefficientcharacteristic velocity
21 Thrust (F) For a rocket engine: Where: = propellant mass flow rate pe = exit pressure, paamb = ambient pressureUejects = exit plane velocity, Ae = exit area
22 Specific Impulse (I or Isp) The ratio of thrust / ejects mass flow rate is used to define a rocket’s specific impulse-best measure of overall performance of rocket motor.In SI terms, the units of I are m/s or Ns/kg.In the US:with units of seconds - multiply by g (i.e m/s2) in order to obtain SI units of m/s or Ns/kg.Losses mean typical values are 92% to 98% of ideal values.
23 Total Impulse (Itot) Defined as: where tb = time of burning If FT is constant during burn:
24 Thus the same total impulse may be obtained by either : high FT, short tb (usually preferable), orlow FT, long tbAlso, for constant propellant consumption (ejects) rate:
25 Effective Exhaust Velocity (c) Convenient to define an effective exhaust velocity (c), where:
26 Thrust Coefficient (CF) Defined as:where pc = combustion chamber pressure,At = nozzle throat areaDepends primarily on (pc/pa) so a good indicator of nozzle performance – dominated by pressure ratio.
27 Characteristic Velocity (c*) Defined as:(6)Calculated from standard test data.It is independent of nozzle performance and is therefore used as a measure of combustion efficiency – dominated by Tc (combustion chamber temperature).
28 Thermodynamic Performance - Thrust Parameters affecting thrust are primarily:mass flow rateexhaust velocityexhaust pressurenozzle exit area
30 Thermodynamic Performance - Specific Impulse Variable Parameters - ObservationsStrong pressure ratio effect - but rapidly diminishing returns after about 30:1.High Tc value desirable for high I - but gives problems with heat transfer into case walls and dissociation of combustion products – practical limit between about 2750 and 3500 K, depending on propellant.Low value of molecular weight desirable – favouring use of hydrogen-based fuels.Low values of desirable.
31 Thrust Coefficient (CF) Maximum thrust when exhausting into a vacuum (e.g. in space), when:(11a)
32 Thrust Coefficient (CF) - Observations More desirable to run a rocket under-expanded (to left of optimum line) rather than over-expanded.Uses shorter nozzle with reduced weight and size.Increasing pressure ratio improves performance but improvements diminish above about 30/1.Large nozzle exit area required at high pressure ratios – implications for space applications.
33 Actual Rocket Performance Performance may be affected by any of the following deviations to simplifying assumptions:Properties of products of combustion vary with static temperature and thus position in nozzle.Specific heats of combustion products vary with temperature.Non-isentropic flow in nozzle.Heat loss to case and nozzle walls.Pressure drop in combustion chamber due to heat release.Power required for pumping liquid propellants.Suspended particles present in exhaust gas.
34 Internal BallisticsLiquid propellant engines store fuel and oxidiser separately - then introduced into combustion chamber.Solid propellant motors use propellant mixture containing all material required for combustion.Majority of modern GW use solid propellant rocket motors, mainly due to simplicity and storage advantages.Internal ballistics is study of combustion process of solid propellant.
35 Solid Propellant Combustion Combustion chamber is high pressure tank containing propellant charge at whose surface burning occurs.No arrangement made for its control – charge ignited and left to itself so must self-regulate to avoid explosion.Certain measure of control provided by charge and combustion chamber design and with inhibitor coatings.