re-entry space capsule

8/14/2019 Re-entry space capsule

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INDIAN INSTITUTE OF SPACE SCIENCE AND TECHNOLTHIRUVANANTAPURAM

Re-Entry Missile

K.Raghava,

Pranav Nath

Parag jyothi

Rudranaraya


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MISSION STATEMENT

Carry a nuclear warhead from an orbit around Earth targeting anylocation

A great boom to national defense

ICBMs has a max. range 5500km

Multiple Independently Targetable Reentry vehicle has range of

about 13000km


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TRAJECTORY

Ideal 7.612km/s delta V (non-realistic)

500km polar orbit (minimum time)

Time period= 1 hr 35 min

3hr time, plane change maneuver

Crudely 30min to 4hr 45min


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TRAJECTORY Contd.

V achievable is 1.884km/s

V of 0.144km/s for reaching opposite corner

V left is 1.47km/s able to provide 130 plane change maneuver.


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TRAJECTORY Contd.


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TRAJECTORY Contd.


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TRAJECTORY Contd.


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TRAJECTORY Contd.


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TRAJECTORY Contd.


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TRAJECTORY Contd.


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TRAJECTORY Contd.

= 177194.39 W/m^2

, = 1873.254 Pa

= 29.331923

Q = 24436183 J

= 27g m/s^2

V hi l ifi ti


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Vehicle specifications

Reentry module : 1300 kg

1. Structure - 400 kg,

2. heat shield - 400 kg,

3. reaction control system - 10kg,

4. warhead disintegration system - 50 kg,

5. navigation equipment (GPS) - 60 kg,

6. electrical equipment - 120 kg,

7. communication system - 25 kg,

8. propellant - 35 kg

De-boosting module : 1700 kg

Propulsion system: 4-20 N attitude control thrusters, 10000 N main engine

Warhead: Thermonuclear : 500 kg

Estimated unit cost : Rs. 60crores


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AEROTHERMODYNAMICS

trade-off b/n warhead size & optimized shapes -usedRlike Apollo, Beagle-2 and Gemini capsule.

minimum possibility of length to place warhead andparachute subsystem so that the whole payload can beminimized.

four surfaces: a sphere segment, a torus segment, aconical frustum and another sphere segment.

shape is axially symmetric.


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Nomenclature

nose radius -RN,

side radius -RS,

rear cone half-angle -c,

mid radius- Rmand

rear conical part length- L

length -L

auxiliary angleN-angulnose sphere.


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Optimized range of values

Parameter Minimum value Maximum Value

RN 2.0 m 7.0 m

RS 0.02 m 0.40 m

C 50 600

Rm Fixed value Fixed value

The Minimum and maximum shape parameter values .

shoulder radius & D - chosen intuitively - range of values specified for the opti

shapes of RVs.


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Selected capsule geometry


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Aerodynamic constants

D = 3

RN = 2

Rs = 0.2

Rs = 0.2 N(deg) = 24.905

C(deg) = 35

Alpha(deg) = 0

L = 3.755

Cpmax = 1.83

Area = 26.2 Volume = 11.0

Cd = 1.28

For the selected capsule,- Cd and max. Cp values using Re-entr

Newtonian software is found.

Good agreement with conventionally used RVs.


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Computational Analysis

2-D , axis-symmet

assumed to descangle of attack tdrag force.

Model - AutodesProfessional

Meshing- Pointwissoftware

computational grid modeled -3 regions each with different grid f

Fine grid cells were made in the vicinity of capsule boundaries.

The grid is coarsened gradually with the distance away from cap


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Meshing

Grid was extended in the axial direction - the subsonic wak

significantly farther downstream compared to supersonic ahypersonic wakes.

Initially, the computation analysis -unstructured meshing in taway from the capsule.

The convergence criteria couldnt be achieved for lower C

(Courant-Freidrichs-Lewy) no.s. Structured grids - used - superior convergence and accura

Approximately 0.3 million quadrilateral cells.


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Boundary conditions

Mach number = 5

free stream condition

P= 1000 Pa

operating pressure= 0

the chance of numerthe low pressures resu

solution.

The free stream temp

Grid was made large enough around the capsule to study the overf Pressure-far-field condition is used for the extreme grids.


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Computational Methods

NavierStokes simulations(i.e., energy model was used).

Spalart-Allamaras Viscous model with default values.

Density based solver - absolute velocity formulations.

The gravitational effect was neglected as there are high devalues comparatively in the reentry phase.

Ideal-gas law - determine the air density,

Sutherlands lawto calculate the air viscosity with the defas defined in FLUENT.


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Computational Methods

checked for Roe-FDS flux type solver

observed to be converging for the density based solver witas 0.005.

Later, AUSM flux-type solver with the same CFL number for results.

Least square cell-based discretization gradient technique .

first-order upwind approximations discretization scheme for and modified turbulent viscosity

convergence residuals has been set for 10-6 .


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Convergence criteria

mass flow rates across the pressure-far-field aft the capsule has been ofor the convergence criteria.

Solution Steering with FMG initialization up to 5 levels for hypersonic flowwith hybrid initialization of iterations.

50000 iterations

Net mass flow rates throughout the domain is observed to be less than the inlet mass flow rate.

Boundary Mass flow rate (

Inlet 6449.0937

Outlet -6452.9462

Net -3.85248

Contours of Temperature


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Contours of Temperature


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Contours of Velocity Magnitude

F hi h M h b


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For higher Mach numbers :

Repeated with the same solving techniques for different free-conditions.

For a free-stream Mach number of 9, the convergence criteriobtained for density based solver with lower courant number

Results were observed to be diverging for higher mach numb


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Propulsion system


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Selection of propellant

Hydrazine monopropellant- space storability,lower maintenance equipment.

Difficulty of using cryogenic propellant for long

durations.

Gives velocity impulses as required.

Situations for maximum performance used for

designing.


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Deboosting engine


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Attitude Control system

4- 20 N thrusters generating about 20 N-m moment on the veh

Inside the reentry module. 40 kg of propellant

Blowdown type, BR=4

Zero-g propellant control

- Bladder tank


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Thrusters

Astrium 20 N thruster

Works on hydrazine mono propellant

Canted nozzle


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Plumbing


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Heating system : Propulsion system

Hydrazine must not freeze, freezing point 2 C. Line rupture can occur due to thawing if liquid is trappe

behind the frozen hydrazine.

Radiative heaters near lines, valves and tanks.

Catalyst beds are also heated to increase performance

and bed life.


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Electric Components : -

Gyroscope sensor :

To establish an inertial reference coordinate frame.

To measure angular rotation (position and rate) of the space vehicleabout the reference axis.

Power source :-

Lithium ion batteries 150 kWh inside the re-entry module and

70 kWh in the de-booster module will be used.

Accelerometer : As an inertial measurement of velocity and position.

As a sensor of inclination, tilt or orientation in 2-3 directions, asreferenced from the acceleration due to gravity.

As a vibration or impact sensor.

i


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Electric Components

3-axis models of both accelerometer and gyroscope sensors have

been used to get a 6 degree of freedom motion.

ANTENNA :

converts electrical power into radio waves

used with radio transmitter and radio receiver

designed for point-to-point operation .

Thrusters :

to provide stabilization about all three axes

Astrium thrusters ; attitude or velocity control of a RV.

limitations are fuel usage, engine wear, and cycles of the controvalves.

El t i C t


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Electric Components

Communication with RV through modulated plasma : Signal antennae are mounted on the casing wall between forward nose

and trail ends of an aerospace launched vehicle for radio communicationthrough a radiation conducting sheet of plasma

re-entry path is maintained by vehicle guidance at a steep angle inresponse to data transmitted to the antennae to enhance data transmissio

GLOBAL POSITIONING SYSTEM (GPS) :

to provide three dimensional position of the vehicle at any time and place

GPS receiver measures its distance from the satellite based on the traveltime of the radio signals

intervening medium effects the signals due to unavoidable causes leadinto errors.

STRUCTURE


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STRUCTURE

Huge aerodynamic loads (high deceleration) Withstand max. deceleration and dynamic pressure

Provide safe compartment for warhead

Adequate ejection mechanism for the warhead

STRUCTURE C td


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STRUCTURE Contd.

Materials From historical data

Titanium for the support structure

Aluminium honeycomb for covering

A proper base since most load is experienced and heat shield shouldnot collapse


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STRUCTURE Contd


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STRUCTURE Contd.

Max. Stress less than tensile stress Displacement 0.09mm

Total weight13000kg, pre-decided - 700kg

Circular beams converted to smaller cross-section, thickness of thebottom plate reduced


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Name Minimum Maximum

Volume 39713300 mm^3


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Volume 39713300 mm 3

Mass 179.10684 Kilogram

Von Mises Stress 0 MPa 51.0098 MPa

1st Principal Stress -18.0792 MPa 46.4904 MPa

3rd Principal Stress -62.0645 MPa 11.6888 MPa

Displacement 0 mm 1.09837 mm

Safety Factor 5.40288 ul 15 ul

X Displacement -0.933833 mm 0.93195 mm

Y Displacement -0.967126 mm 0.889766 mm

Z Displacement -0.903069 mm 0.000655074 mm

Equivalent Strain 0 ul 0.000463033 ul

1st Principal Strain -0.000000021521 ul 0.000358405 ul

3rd Principal Strain -0.000523155 ul 0.000000506277 u


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STRUCTURE Contd.


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STRUCTURE Contd.

Further work to be done

Structure able to bear 27g load with 1mm displacement

Outer cover of aluminium honeycomb structure to sustain maximumdynamic pressure

SEPARATION OF WARHEAD


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S O O

A cap at the top removed

Launching a parachute from inside attached to the warhead

Due to drag, warhead comes out and capsule continues in its path

Parachute attachment is removed and warhead hits the target

CLAMPING OF WARHEAD WITH


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CAPSULE

Warhead


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Thermonuclear warhead - high explosive destructive force.

Warhead - loaded on to the Re-entry module.

structure for holding & supporting the warhead : protecting from heat lvibrations.

Nuclear bomb carries out three levels of destruction.

a) By direct blast that will account for about 40 % of total energy.

b) Thermal radiations comprising of 30 % energy

c) By Ionizing and residual radiations

Warhead


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similar to B28 developed by the United States.

Warhead : 80 cm in diameter & 200 cm long.

Uses fission bomb in the primary zone to activate fusion reactions in the szone.

Warhead


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Fired about 20 s from the release from reentry capsule.

Destruction range of about 1 Megatons of TNT.

After falling 2 km from the release, timer would switch on the closed parwarhead circuit.

Other sequential switches will be opened to charge the batteries to chanumber of capacitors and turn on the radar fuses.

Radar signals is to be reflected from the surface of earth and trigger thecurrent to the detonation signal.

Thermal Design


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g

In-orbit control and Atmospheric entry

Thermal environment requirements of various subsystems

In-orbit:

Average spacecraft temperature299 K


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Qsun

the heat transfer by direct sunlight,

Qer be the heat transfer due to earth reflected sunlight,

Qibe the internally generated heat,

Qss be the net power radiated to space and

Qsebe the net power radiated to earth.

Qsun =

Qer =(zero) assuming zero view factor from sunlit earth

Qi = 30 W (internal heaters)

1400* * 1400(0.7)(13.4) 13132n

A W


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The view factor for spacecraft in 500 km orbit, Fs,e= 3.9309/ = 0.3128

4 4,( ) ( )s s s s s s e e sun er iA T A F T Q Q Q

4 4

,( ) ( ) sun er i

s s e e

s s

Q Q QT F T

A

Multi-layer insulation


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Each layer is of Mylar material almost 7 micrometer thick, coasides by aluminium using vapor deposition methods.

Outer layers- 120 micrometers thick

Small holes for venting gases trapped. Spacers of Dacron Nearly 20 layers

Heaters


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Patch heaters at required places ~ 1 W each

Electrical resistance element sandwiched between two

sheets of flexible electrically insulating material Kapton.

Fuses and relays provided

Thermistors sense temperature

Thermal protection system


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Maximum heating rate: 508 kW/m2

Total Heat load: 24436.183 kJ

Ablative Heat shield

PAN Carbon-Carbon composite as TPS material.

Estimated Weight = 400 kg

Estimation of Heat Shield thickness


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=

Heat of fusion = 200 kJ/kg

Density of ablative material is 1500 kg/m3

From above equation,

24436.183 x A = 200 x 1500 x A x t

t = 8.14 cm.

Half the thickness for rear-body

Further optimization required.

Vehicle geometry


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Length: 6 m

Max. diameter: 3 m

Surface area: 26 m2

Ballistic coefficient: 38.46

Center of gravity

1. 1m from fore body shell,

2. 1.5m from attitude control thruster position towards

main propulsion system.


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re-entry space capsule

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