avd report final

8/3/2019 Avd Report Final

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Nitish Kumar

Janmejay Jaiswal

Lala Surya Prakash

11/17/2011

Reentry Vehicle Design


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Table of Contents

List of Symbol .................................................................................................................... 3

List of figures ..................................................................................................................... 4

Mission Requirements .......................................................................................................... 5

Vehicle Configurations ......................................................................................................... 6

Initial Mass Estimation9 .................................................................................................... 8Orbit Parameters ................................................................................................................ 9

Payload ........................................................................................................................... 10

Guidance, Navigation and control ........................................................................................ 13

Power System ................................................................................................................... 13

Thermal analysis ............................................................................................................... 14

CFD & FEM ANALYSIS .................................................................................................... 17

Structure ......................................................................................................................... 21

Trajectory ....................................................................................................................... 23

Blackout during Reentry................................................................................................. 28Calculation for blackout time and altitude......................................................................... 29

Propulsion ....................................................................................................................... 30

Recovery System .............................................................................................................. 30

Parachute..................................................................................................................... 30Conclusion ....................................................................................................................... 31

References ....................................................................................................................... 32

APPENDIX ...................................................................................................................... 33


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List of Symbol

R- Base radius of spacecraft

L- Length of Spacecraft

Rn- Nose radius of spacecraft

V- Velocity of spacecraft

r-Distance of spacecraft from center of Earth

A-Apogee

P-Perigee

V-Velocity Impulse

cirV -Velocity of spacecraft in initial circular orbit

en- Reentry flight path angle

q - Heat transfer rate

-Density

hw-Wall enthalpy

ho-Stagnation Enthalpy

-sweep angle of delta wing

LBC-Lifting ballistic coefficient

L/D- Lift to Drag ratio

ne,w-Electron density in wake region

ne,s- Electron density in stagnation region

IspSpecific Impulse

i- initial satte

f-final state


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List of figures

Figure 1 Mission Profile ....................................................................................................... 5

Figure 2 Conceptual design of Subsystem of reentry vehicle ...................................................... 6

Figure 3 Fore body shape Design .......................................................................................... 7

Figure 4 Positioning of Sub system inside reentry vehicle .......................................................... 8

Figure 5 Stagnation heat transfer rate at different altitude ...................................................... 14

Figure 6 Heat transfer rate at windward surface .................................................................... 15

Figure 7 Heat transfer at leading edge of wing ....................................................................... 16

Figure 8 Density variation along the length of vehicle obtained from CFD analysis ..................... 17

Figure 9 Static temperature contour on reentry vehicle obtained from CFD analysis ................... 18

Figure 10 Mach number contour on reentry vehicle obtained from CFD analysis ........................ 18

Figure 11 Static Pressure Variation along length of vehicle obtained from CFD analysis .............. 19

Figure 12 Total Pressure Variation along length of vehicle obtained from CFD analysis ............... 19

Figure 13 Wall Shear Stress with length of vehicle obtained from CFD analysis .......................... 20

Figure 14 Mesh and Loads for FEM ANALYSIS ..................................................................... 21Figure 15 Total Translation (in mm) ..................................................................................... 22

Figure 16 Von Mises Stress (in kPa) .................................................................................... 22

Figure 17 Altitude vs Deceleration ...................................................................................... 24

Figure 18 Altitude vs V/Vre ............................................................................................... 25

Figure 19 Downrange vs V/Vcs .......................................................................................... 25

Figure 20 Time vs V/Vcs ................................................................................................... 26

Figure 21 Entry and touchdown point location ...................................................................... 27

Figure 22 Relation of Volume-Wing area parameter to L/D at hypersonic speeds ....................... 28


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Mission Requirements1. To conduct experiments in micro gravity conditions. We have chosen ISS orbit for our experiments.

2. To perform experiments on rat and to bring them safely to Earth.

3. Deceleration of vehicle should not exit 1g.

Mission Profile

Figure 1 Mission ProfileAfter conducting micro gravity experiment in 400 km circular orbit, we will deboost spacecraft to 200

km circular orbit by using solid propellant. Since we are conducting bone loss experiment, we require

more time. So the duration for the whole mission is 10 days.


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Vehicle Configurations

Figure 2 Conceptual design of Subsystem of reentry vehicle


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We are designing lifting body reentry vehicle for our mission. Following are the advantages of lifting

body

1. Increase the width of entry corridor2. Lesser deceleration than ballistic reentry vehicle.3. Enlarge the landing footprint4. Provide additional trajectory options such as skipping trajectory5. Execute non propulsive plane changes.

During this design process, we are trying to optimize deceleration and heat flux.

Figure 3 Fore body shape DesignData:

L=1.5m

R=1m

rn=0.3m

Half cone angle= 33.6 degree

From elementary geometry,

22

22

R+L

r

R

L=x nt

L

Rx=y tt

22 tnto yr+x=x

noa xx=x


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Solving for xt, yt, xa and xo, we get

xa=0.2413m ; xo= 0.5413m

xt=0.374m ; yt=0.249m

Volume of the vehicle = 1.6277 m3

Figure 4 Positioning of Sub system inside reentry vehicleInitial Mass Estimation9

Sub System Mass (in kg)

Payload 100

Structure 300

Power System 75

Flight Instrumentation & NGC 180

Thermal Protection System 60

Recovery 105

Launcher Interface 15

Total Initial Mass 835

Growth Margin 5%


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Orbit Parameters

Initial Orbit Parameters in Keplerian Coordinate System

Radius 6764.4 km

Eccentricity 0

Inclination 51.6 deg.

De-boost phase

The orbit will become circular of altitude 200km after de-boosting phase .We are doing an impulsive

Hohmann transfer for de-boosting. Velocity impulse required for the transfer is calculated using vis-

viva equation

ar=V

122

Initial Orbit

r = a = 6764.4 km

VI= 7.6763 km/s

Transfer Orbit

rA= 6764.4 km

rP = 6578 km

a = 6671.2 km

VTA= 7.6225 km/s

VTP = 7.8385 km/sV1=VI VTA= 0.0538 km/s

Final Orbit

r=a= 6578 km

VF=7.7843 km/s

V2=VTF VF =0.0542 km/s

Vtotal = V1+V2 = 0.108 km/s

The optimal value of reentry angle can be found by the following formula8

22

2r1

atencir

atcirat

cirrsecr

rrV=V

This gives the minimal flight path angle at which reentry takes place. If we decrease the flight path

angle below this atmosphere may not be able to capture the vehicle.

en = 1.60089 deg


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PayloadWe are carrying three payloads to perform experiments in micro gravity conditions

1. Protein Crystal Growth Experiment (CPCG)The goal of the Commercial Protein Crystal Growth payload is to grow high-quality crystals of selected

proteins so that their molecular structures can be studied. On Earth, gravity often has a negative

impact on growing protein crystals. In microgravity the near weightlessness of space gravitys

disturbances are removed, allowing crystals to grow in a more regular and perfect form.

When the microscopic crystals are returned to Earth, scientists will use X-rays to help analyze the

crystals and to map the locations of a proteins atoms. This information allows them to make

computer models depicting the biological molecules and to determine how the biological substances

function. Knowing the atomic structure may help pharmaceutical companies develop medicines that

fit into a proteins structure much like a key in lock. This research may lead to more effective

medicines with fewer side effects.

Experiment chamber in the CPCG-H consists of a small chamber to hold the protein crystal solutionand reservoir chamber to hold the precipitating agent solution. A small droplet of protein solution is

mixed with a small amount of precipitating agent solution and placed in the protein chamber. The

larger chamber is filled with more concentrated precipitating agent solution, which is captured in a

polymer wicking material to keep the solution from moving around the chamber. During activation,

the protein chamber is rotated so that it is in vapor contact with the reservoir. Water molecules

migrate from the protein droplet through the vapor space into the more concentrated reservoir. As

the volume of the protein droplet decreases, the concentration of protein increases and protein

crystals form. As the experiment proceeds, the crystals become larger.

Advantages

Structural studies using microgravity-grown protein crystals may provide information that can be usedin the development of new drugs. With the advent of genomic information from humans and many

other species, the roles proteins play in diseases and degenerative conditions is becoming more clear

and the need for information about the structure of these proteins more critical.

Benefits from microgravity protein growth experiments have already been seen. Many of the

crystallization experiments conducted on the Space Shuttle have yielded crystals that furthered

structural biology projects. For example, microgravity crystallization experiments have been conducted

with recombinant human insulin. These studies have yielded X-ray diffraction data that helped

scientists to determine higher-resolution structures of insulin formations. This structural information is

valuable for ongoing research toward more effective treatment of diabetes. Other very successful

microgravity crystallization experiments have provided enhanced X-ray diffraction data on a proteininvolved in the human immune system. These studies have contributed to the search for drugs to

decrease inflammation problems associated with open-heart surgery.

To conduct this experiment, temperature is maintained at 22C .Also maximum acceleration should be

less than 1g to perform experiment13

.


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Specification

Mass of the setup: 8 kg

Power Requirement: 0

Volume Required: 356.4 cm3

2. Physiological System Experiment

Osteoporosis is a disease marked by a progressive loss of bone mass. It currently affects a largenumber of people. Immobilization of any kind for a longer time can cause bone loss; hence it is also

the problem faced by bedridden or paralyzed patients. Because exposure to micro-gravity results in a

loss of bone mass similar to the effects of osteoporosis and immobilization, the space environment

serves as a laboratory for studying these conditions. The changes seen in bone tissue after exposure to

microgravity are, in fact, more similar to the changes seen in osteoporosis.

Hence payload to study the capacity of a synthetic protein molecule to halt or slow bone loss in

micro-gravity is taken.. The protein has potential use in countering conditions that involve loss in bone

mass. It may also be useful for combating the bone loss that is likely to be experienced by astronauts

on long-term space flights in the future.

This experiment deals with the bone loss in rats. Four rats are under investigation. Two of them are

trained (like subjected to centrifuge etc) and other two are normal rats.

Specification

Weight of each rat = 250 gm

Weight of food & water = 17 kg

Power required = 28 W

Operational hour=16 hrs per day

Volume required = 35732 cm3

Sub systems required for this experiment:A) Air Quality Control

B) Lighting System

C) Water Refill Box

Water Refill Box is supplementary hardware that can be used to replenish drinking water in the AEM

(Animal Enclosure Module) for missions longer than 5 days.

D) Temperature Recorder

The Ambient Temperature Recorder (ATR) is a self-contained, battery-powered instrument,

approximately the size of a deck of cards. It may be placed in almost any environment (not

submersible in liquid) to provide recording of up to four channels of temperature data. External

probes are flexible to allow the user to place probes at various locations within the sensedenvironment. Standard length for probes is 3 feet, but they may be longer or shorter, if required.

Power for this is provided by two internal Lithium thionyl chloride batteries. An O-ring seal protects

the internal electronics of the ATR from fluids in the environment and permits operation in damp or

humid environments, such as an animal habitat.


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3. Plant Growth Unit

The Plant Growth Unit (PGU) is a self-contained system and designed to hold removable Plant Growth

Chambers. The PGU consists of the support com- ponents and a cavity for growing plants. The PGU is

equipped with a 15 W plant growth lamps (Vita-Lite spectrum), a timer to provide day/night cycling,

temperature sensors, electronically-controlled fans, heater strips for temperature modification, data-

acquisition system, and internal batteries.

For environmental control, two thermostatically-controlled variable-speed fans draw air over the plant

growth chambers. A temperature gradient decreasing from the top to the bottom of the chambers is

maintained to prevent moisture condensation in front of the light. Diurnal temperature cycling is

provided, with a chamber temperature of 25.5 0.61C during the daylight and 23.3 0.61 C during

the night.

Specification

Weight of setup = 35 kg

Power Required = 20 W

Operational hour = 16 hrs

Volume required = 42240 cm3

Temperature controlling unit

To control the temperature within the limits, refrigerator/incubator system is used. Details of

refrigerator/incubator are given below:

Weight = 20 kg

Power Required = 84 W

Operational hour =24 hrs

Volume required = 59119.2 cm3

Calculation of oxygen required for rats14

Density of O2 = 1.429 g/ml

Minute Volume = 1.4 ml/g (body weight)

Total minute volume for four rats = (1.4*4*250) ml/min = 1.4 l/min

Total O2 required for 10 days = 1.4*14400 = 20160 liter

O2 mass = 28.81 kg

In a M6 cylinder, 164 liters of O2 is compressed to nearly 1 liter at a pressure of 137 bars. So in our

case volume of vessel required for storing oxygen is

Volume of vessel = 20160/164 = 123 liters = 0.123 m3

Total payload weight = 113 kg

Total payload volume = 151478 cm3


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Guidance, Navigation and controlThe guidance is performed to maintain defined drag accelerations versus velocity profile, taking into

account the fixed constraints on the heat flux and the aerodynamic loads, and the avoidance of the

active oxidation of the selected thermal protection materials. Antenna is placed in a wake region to

avoid blackout.

The navigation consists of a combination of inertial measurements and GPS updates before 120 km

and a Drag Derived Altitude (DDA) update at 60 km, during the blackout phase.

Guidance is controlled by On Board Computer (OBC).Power required by OBC is 20 W and Operational

hour is 24 hrs

Power SystemNi-H2- The pressure vessel cases of Ni-H2 batteries are generally cylindrical with hemispherical ends.

This makes close packing difficult. Also it is used in large spacecraft. Hence we are not selecting.

Li battery- its energy density is higher compared to other batteries but certain types of Li batteries

experienced teething problem in early application showing a tendency to explode in some situation.

Hence this battery is also rejected.

Ag-Zn It has a good energy density (175Wh/kg, Primary) , high voltage per cell and does not have

above disadvantages. Thus it is selected as the power source.

Total Power required for 10 days = (28*16+20*16+84*24+20*24)*10

=32640 sec

Energy density= 175 Wh/kg

Estimated mass of battery= 185 kg


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Thermal analysisDue to very high Mach number during reentry, temperature at the wall is of the order of 1500 K. So

proper insulation is necessary to safely return vehicle to Earth. Thus we need to accurately measure

the heat transfer rate so that we can select proper material for in insulation.

The simplest method for estimating hypersonic aerodynamic heating refers to the following general

relationship2

baVC=q

If the values of C, a and b are calculated from below formula, we get the stagnation point heat transfer

at fuselage nose ( q ) from the above formula. The graph is plotted

o

w

nh

h

R=C 1

1101.83

4

32

1=b,=a

Figure 5 Stagnation heat transfer rate at different altitude


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In order to perform the vehicle windward surface heating analysis ( FPq ), values of C,a and b are given

by

o

w

h

h

x=C 1cos

sin102.53

5

3.21, =b=a For conservative purpose, we take hw=0 and =55. Also we assume x=0.6m. We get the following

graph.

Figure 6 Heat transfer rate at windward surface

We use heat transfer rate at nose and on windward surface to calculate heat transfer rate on wings

leading edge (LE). It is given by

q+q=q FPN cossincos2

1 22

For getting conservative value, is assumed to be zero.

Also =90-(half cone angle)


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Figure 7 Heat transfer at leading edge of wingWe also tried to calculate the thickness of heat shield but value of thickness is coming out to be 15 cm

everywhere. This value seems to be reasonable for nose cone but not for entire vehicle.


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CFD & FEM ANALYSIS

We have done CFD analysis on two dimensional planar body as shown in fig 3 and FEM analysis on

three dimensional body (fig 14). We have neglected wing to avoid unrealistic results .The flight

condition for which the analysis is done is given below

Mach No. 8.4

Altitude - 50 km

Static Temperature -273 K

Static pressure -92.522 Pa

Density 1.18*10-3

kg/m3

Total Pressure 1241110 Pa

Total temperature 4125 K

Figure 8 Density variation along the length of vehicle obtained from CFD analysis


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Figure 9 Static temperature contour on reentry vehicle obtained from CFD analysis

Figure 10 Mach number contour on reentry vehicle obtained from CFD analysis


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Figure 11 Static Pressure Variation along length of vehicle obtained from CFD analysis

Figure 12 Total Pressure Variation along length of vehicle obtained from CFD analysis


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Figure 13 Wall Shear Stress with length of vehicle obtained from CFD analysis1. From figs. 8,11 and 12, we observe that there is a sharp increase in values of density and static

pressure and a decrease in total pressure. This is due to the formation of bow shock wave

ahead of the vehicle. Then the density and static pressure will decrease due to expansion.

2. The static temperature contour diagram shows that the maximum temperature reached is about4000K at stagnation point. The temperature obtained is unrealistic because we have not

accounted for radiation and convection.

3. The wall shear stress is zero at nose because the velocity is zero at stagnation. Then it increasesfor curved portion and then it decreases. This is due to the fact that the flow is attached in

curved portion and separates after that.


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Structure

We observe from static temperature contour diagram (fig 9) that the average temperature on wall is

about 2000K and the average temperature on nose is much higher. But the temperature will be lower in

actual case due to radiation and other heat transfer effects. Hence we can use Reinforced Carbon-

Carbon composites for nose and silica tiles for rest of the body. The wing leading edge heat transfer

rate is higher than the windward surface as seen from fig.6 and fig 7.So we can use ReinforcedCarbon-Carbon composites on leading edge.

Based on the results obtained from CFD analysis, we observe from fig 11 that static pressure on nose

cone is about 8400 Pa and average static pressure on the rest of the body is about 3500 Pa. Using these

values of pressure, we performed FEM analysis on the reentry body. The mesh and results are as

follows:

Figure 14 Mesh and Loads for FEM ANALYSIS

No. of Elements 38921

No. of nodes 8630

Element type Tetra4

Mesh 3D mesh


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Figure 15 Total Translation (in mm)

Figure 16 Von Mises Stress (in kPa)


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Properties of Materials Used

1. Reinforced Carbon-Carbon CompositesMax. temperature 2000K

Density 1900 kg/m3

Ultimate tensile strength 480 MPa

2. Silica TilesMax. Temperature 1700K

Density 300 kg/m3

Tensile Strength 150 kPa

FEM Results

Translation(mm) Von Mises Stress (kPa)

Max 1.933e-004 3.918e+000

Min 0.000e+000 5.365e-002

From the data obtained, we can observe that the stresses on reentry body are much smaller than theultimate tensile strength of the materials used.

TrajectoryAssumptions

1. Effect of Angle of Attack and flight path angle is neglected2. As V/Vcs become 1, most of the parameter tends to infinite. Hence we consider values of

parameter corresponding to V/Vcs=0.95.

3. We consider exponential density model for atmosphere.Following are the expressions to find different parameter to define reentry trajectory

12.

hcs eLBCgr+=

V

V

2/1

1

000

DL

V

V

=ncs

/

1

2


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2

2

1

1

ln2

1

cs

cs

V

V

V

V+

D

L

g0

r0=t

2

01ln

2 csV

V

D

Lr=S

Here S is Downrange which is the distance travelled in the direction of entry velocity in the verticle

plane.

Figure 17 Altitude vs Deceleration


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Figure 18 Altitude vs V/Vre

Figure 19 Downrange vs V/Vcs


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Figure 20 Time vs V/VcsFrom fig 19, downrange at V/Vcs=0.95 is 1500 km and from fig 20, touchdown time is 0.69 hrs.

Using downrange we can calculate reentry points coordinates.

The reentry point longitude-latitude are found using Napier formula for spherical triangle. Landing

point is fixed in Bay of Bengal at 18 deg N, 88 deg E. This target is selected because it is far away from

landmass.


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Figure 21 Entry and touchdown point locationAerodynamics

Since speed encounter during reentry is very high, weve decided to use De lta wing as our lifting

surface.

Advantage of delta wing

1. Ease of manufacture,2. Higher strength

3. Substantial interior volume for fuel or other equipment.

4. It can be made very robust (even if it is quite thin), and it is easy and relatively inexpensive to build

CD=2*sin2()=0.6125

Volume of vehicle (Vol) =1.6277 m3

Wing Span = 4 m

Wing Area (AW) = 2.4 m2

Vol2/3

/AW = 0.576


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Figure 22 Relation of Volume-Wing area parameter to L/D at hypersonic speedsFrom fig 22 15 we can see that (L/D)max = 1.8

Blackout during Reentry

When a spacecraft enters a planetary atmosphere at a velocity significantly exceeding the speed of

sound, a shock layer forms in front of the body. The sheath of ionized particles, which develops around

the spacecraft, is the result of ionization of the atmospheric gases as they are compressed and heated

by the shock, or heated within the adjacent boundary layer. When the electron density gets

sufficiently high, such that it exceeds the critical plasma density of the link frequency, communications

can be disrupted, with the result being significant attenuation or even blackout. The surroundingplasma on a spacecraft entering a planetary atmosphere will attenuate any radiated signal b

absorption and reflection if its density is sufficiently high. Electrons are the main contributors to

deflection of waves in a plasma gas.

Link Frequency (GHz) Designation Critical Electron Density/cm3

0.401 UHF 1.99 109

2.3 S-band 6.56 1010

8.4 X-band 8.75 1011

32.0 Ka-band 1.27 1013

The ratio of electron number density in the wake region to that at the stagnation point can be

estimated using7

Gse,

we,

P

P=

n

n/1


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Where P /PG is the pressure ratio across the shock region, is the isentropic expansion coefficient

(or ratio of specific heats), which ranges from 1.3 to 1.4 , and is a correction factor normally taken to

be unity.

For this study, a value of = 1.4 for Earth. The approximation of equation can produce different results

because the expansion coefficient, , is not constant as the gases expand, and the wake-region

pressure can be significantly higher than the free-stream pressure, P , during certain stages of theatmospheric entry profile. Thus, the inclusion of the factor in the above equation will allow for a

better approximation of the wake-region electron number concentration. This assumes that the

communications antenna is located in the wake region. The electron density in the stagnation region is

assumed to remain constant as the gases flow around the spacecraft into the wake region.

Thus, the basic rule of thumb for testing for a blackout condition is

1. If the calculated electron number density lies below the critical plasma number density by more

than an order of magnitude, blackout is deemed unlikely

2. If the calculated electron density lies above the critical electron number density by more than an

order of magnitude, blackout is deemed likely.

3. If the calculated electron number density is within an order of magnitude of the critical number

density, blackout is deemed uncertain, but possible.

Calculation for blackout time and altitude

Maximum electron density at stagnation point is taken from MA-6 reentry data se,n = 5*1013

/ cm3.

At two different altitude we calculate electron density in wake region.

Altitude (km) Temperature(K) V/Vcs V (m/s) Mach no.

GP

P we,

n / cm3

60 258 0.6 4819.8 14.97 3.827*10-3

9.385*1011

80 175 0.95 7631.35 28.78 1.035*10-3

3.688*1011

We observe from the above table that the electron density in wake region is one order of magnitude

higher than electron density at stagnation point. Thus in this region, blackout is most likely to happen.


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Propulsion

Total Velocity impulse is 108 m/s. Assuming de-boosting time from 400 km circular orbit to 200 km

circular orbit to be 60 sec, we got acceleration of 1.8 m/s2. Assuming Isp of propellant to be 300 sec,

from Rocket equation we get

mi/mf=1.037

Since mi=850 kg, final mass of spacecraft is 819.37 kg. Thus mass of propellant is 30.63 kg.

Corresponding thrust is 1530 N.

Reaction Control System- We use cold compressed Nitrogen gas as RCS for attitude control of vehicle

for aligning vehicle during reentry.

Recovery SystemThe landing system consists of the Drogue Parachute, Main Parachutes and the Landing Bag. The

Drogue is deployed by means of a pyrotechnic mortar. Primary task of Drogue parachute is to stabilize

the vehicle and reduce descent speed at which the main parachute is deployed.

Parachute

Two-stage parachute systems have been used to recover the spacecraft. The recovery design was

based on a reentry vehicle weighing 850 kg with an initial velocity of 8000 m/sec at a trajectory angle

of 1.60089 degree below the horizontal at an altitude of 200 km.

Recovery was desired to commence at an altitude of about 5 km after the vehicle had passed through

the peak heating and maximum stagnation pressure environment. This altitude is selected to reduce

dispersion from target point. The main function of parachute is to decrease the velocity of vehicle .

Parameters

Material Nylon

Cone angle 20 deg

No. of gores 12

Suspension line strength 3.3 kN

Max dynamic pressure 96000 Pa

We are opening parachute at 5 km altitude to get less dispersion from landing point. Large velocity at

higher altitude does not allow deployment of parachute due to vortex formation.

Velocity at 5 km =136.5 m/s

Velocity near splashdown = 10 m/s

time required to splashdown from 5 km altitude = 16 sec

We computed Deceleration = 7.84 m/s2

From Newtons law we get


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Total tension = m*(g+a)/cos(20) = 15.97 kN

Tension in each cable = 1.33 kN

Dynamic pressure at 5 km altitude = 5500 pascal

Area of the canopy = (Total tension)/( Dynamic pressure) = 2.9 m2

Diameter of canopy (circular) = 1.92 meters

Finally spacecraft splashdown in the sea where flotation bag is opened and spacecraft is recovered.

ConclusionWeve designed reentry vehicle which will successfully complete the mission in 10 days. The

experiments will be done and data will be returned to Earth. Vehicle will splashdown in Bay of Bengal

away from landmass. Weve done analytical and numerical (CFD and FEM) analysis. Analytical analysis

is based on first order approximation of main equation and is not very accurate.

To calculate orbit parameter, weve made matlab code which is given in appendix. 2-D CFD and 3-D

FEM analysis of reentry vehicle are done and results seems to realistic. They are in closeapproximation to the analytical results. 3-D CFD analysis is not done because we were not getting

reasonablr result and problem for this remains unknown to us.

We have included materials for TPS which are available in India. Other materials can give better

performance but we have to take into account the availability and cost of material. A lot of design for

reentry vehicle is possible but we have selected a configuration which is not available in literature

upto our knowledge.


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References1. THERMAL PROTECTION MATERIALS FOR RE-USABLE LAUNCH VEHICLES ,M.R.Ajith, R Gopi, V

Chandrasekaran, M C Mittal, P P Sinha ,VIKRAM SARABHAI SPACE CENTRE

2. Heat Transfer Analysis for a Winged Reentry Flight Test Bed, Antonio Viviani, Giuseppe Pezzella3. Preliminary report on development of Interim parachute recovery system for reentry vehicle,

William Pepper

4. Communication blackout at reentry, F.J.Tisher, University of Alabama Research Institute,19635. ROUGH ESTIMATE Of THE "BLACKOUT" TImE In RE-ENTRY COMMUNICATIONS, F.J.Tisher,

Goddard Space Center

6. Atmospheric Re-Entry, John C. Adams, Jr.7. The Spacecraft Communications Blackout Problem Encountered during Passage or Entry of

Planetary Atmospheres, D. D. Morabito, IPN Progress Report

8. RE-ENTRY TRAJECTORY SIMULATION OF A SMALL BALLISTIC RECOVERABLE SATELLITE, WalkiriaSchulz* Paulo Moraes Jr, ADVANCES IN SPACE DYNAMICS 4: CELESTIAL MECHANICS AND

ASTRONAUTICS,

9. Wijker J Lecture series ae2 S02,Delft University of Technology10.Rocket Propulsion Element,Sutton, Biblarz, Wiley11.Space Vehicle Design ,Griffin , French,AIAA12.Introduction to Space Flight, Francis J Hale13.Method of growing protein crystal by Lawrence, Delucas, Patents14.Journal of Applied Toxicology, Joe L.Manderly,Volume 6, Issue 1, pg 25-30,


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APPENDIX

/*Program to claculate trajectory parameters */

h=0:1:200;

rho=1.225;

r0=6378e3;bet=0.1378;

n=0;

m=850;

CD=0.6125

//lbyd=[1 1.5 2 2.5 3];

lbyd=[1.8 1.8 1.8 1.8 1.8];

g0=9.81;

a=3.14;

bc=m*g0/(CD*a);

for j=1:5

lbc(j)=bc/lbyd(j);

end

lbcnew=gsort(lbc,'g','i');

bcnew=gsort(bc,'g','i');

printf("\n")

Vcs=sqrt(g0*r0);

gamma=1.6;

phi=acot(sqrt(1+.106.*(lbyd).^2));

cr=(r0*(lbyd).^2)./(5.2*sqrt((1+.106.*(lbyd).^2)));

LF=1./cos(phi);

for i=1:5

for j=1:201

Vden=sqrt(1+(rho*r0*g0/(2*lbcnew(i)))*exp(-bet*h(j)));

V(j+201*(i-1))=1/Vden;

Q(j+201*(i-1))=0.5*rho*(V(j+201*(i-1))*Vcs)^2*exp(-bet*h(j));

d(j+201*(i-1))=((V(j+201*(i-1)))^2-1)/lbyd(i);

time(j+201*(i-1))=.5*sqrt(r0/g0)*lbyd(i)*log((1+(V(j+201*(i-

1)))^2)/(1-(V(j+201*(i-1)))^2));

dr(j+201*(i-1))=-.5*r0*lbyd(i)*log(1-(V(j+201*(i-1)))^2);

//q(j+201*(i-1))=(V(j+201*(i-

1))*Vcs)^3*sqrt(lbcnew(i)*sind(gamma));

//printf("%f\n",time(i))

end

// Q(i)=Vcs^2*sqrt(lbyd(1)*bcnew(i)/sind(gamma));

if modulo(n,1)==0 then

//figure(0)

//plot(V(1+n*201:201+201*n),time(1+n*201:201+201*n))

//xlabel('V/Vcs')

//ylabel('Altitude(km)')


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//printf("%f\n",time(i))

//figure(1)

plot(V(1+n*201:201+201*n),dr(1+n*201:201+201*n))

//plot(V(1+n*201:201+201*n),dr(1+n*201:201+201*n))

end

n=n+1;end

/* Program to calculate Heat Transfer Rate at stagnation point*/

h=0:1000:200*10^3;rho=1.225*exp(-h./7200);cpw=.25*4.18*10^3;R=6378*10^3;rcs=R+h(length(h));g=9.81;

vcs=sqrt(g*rcs);hrw=0;h0=vcs.^2/2;rn=.3;C=1.83*10^-4*(1-hrw./h0)/sqrt(rn);b=3;a=.5;lbd=1.8;cd=.6125;A=3.14;W=850*g;v=vcs*(1+(lbd*cd*A*rho*g*R)/(2*W)).^-0.5;q=C.*rho.^a.*v.^b;plot(q,h);hold on

avd report final

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