# Presentation: 572375

## The Jovian Entry Probe:

**A. Atzei, A. Santovincenzo, H. Ritter,  F. Mazoue, ****P. Falkner and the CDF JEP Team**
**ESA/ESTEC**_[Alessandro.Atzei@esa.int](mailto:Alessandro.Atzei@esa.int)__   tel: +31 071 565 8059_

- A Feasibility Study of a Minimum Resource Jovian Entry Probe

##             Overview:

-             Overview:

- Context of Jovian technology Reference Studies
- Requirements and constraints
- Design approach
- Mission Analysis
- Aerothermodynamics
- Thermal Protection System
- Resulting configuration
- Conclusion

## Jupiter related Technology Reference Studies

_**Goal:**_ Identify and when possible assist in the development of enabling technologies for future science missions to the Jovian system. Furthermore, to support the scientific community in the field of Jovian exploration.
- Presently two studies have been completed and a new study has been initiated:
-  Jovian Minisat Explorer: Focussing on the exploration of     Europa (included Europa polar orbiter and Jovian equatorial relay S/C,       implications of RPS, as well as small impactors and Europa lander) 
-  Jovian System Explorer:
- Study of the Jovian magnetosphere (one or more magnetospheric S/C)
- Jovian Entry Probe: Study of the Jovian atmosphere
- one or more entry probes, up to 100 bar 

## JEP Study requirements and constraints

- Carry the probe to Jupiter and release it
- Perform **entry and descent** into the Jovian atmosphere at near equatorial latitude  (with an option of non-equatorial descent up to -30deg/+30 deg, if possible)
- Measure atmospheric properties **in-situ** down to an altitude corresponding to **100 bar **using a given Strawman payload
- Transmit the data **in real time** to the accompanying Orbiter
- Achieve a final orbit for magnetospheric measurements with the Orbiter
- Achieve multi-probe mission if mass allows

-  Launch vehicle: Soyuz Fregat 2-1b from   Kourou
-  Preferred launch dates: 2016 or 2023
**Payload: 12 kg; 30 W; 5 l; 353 bps****  (Highly integrated)**
-   Avoidance of Jovian ring when   defining probe approach, while not   exceeding distance during comms
-  Design shall be compliant with Beagle 2   Enquiry Board recommendations and    Huygens Lessons Learned
-   Max. heat flux during entry: 500    MW/m2   (assumed as maximum    capability for present TPS technology)

## Mission design drivers

**Jupiter atmosphere**
-  Uncertainties on the physical and chemical parameters
-  High temperature, high pressure at low altitude
-  Strong attenuation for Comms below 20 bar altitude (100 bar ~24 dB)

**High velocity entry**
- Can not be reduced below ~ 47 km/s
- Aerothermodynamic phenomena, in this regime and for the Jupiter atmosphere, not well understood (uncertainties in calculation of the heat fluxes/loads)
- Very high aerothermodynamic heat fluxes at the limit of present TPS technology capabilities
- High TPS mass fraction (50-70%)
- Very high g-load (further qualification of components)
**Synchronisation probe/Orbiter (transmission in real time)**
- Orbiter to Probe link to be phased with deployment and relay phase
**Mass**
- Provides a higher limit for the allowable mass (Launcher performances)
- Fulfilment of mission requirements provides a lower limit

## Probe design : Concurrent design approach

**Heat fluxes **
**Heat load**
**Max 500 MW/m****2**

**Pressure max. 100 bars**

**Data real time transmission**

**Payload (mass,**
**volume, data rate, power, Temp. op.)**

**Parachute size**

**TPS****≈50% of the total probe mass**

**Entry velocity**
**FPA**

**Atmospheric probe trajectory**

**DM minimum sphere volume and mass**

**Probe volume**

**DM spherical-type shape**

**Battery, PCDU, CDMU volume**

**Probe minimum mass**

**Interp. Trajectory + Orbiter final orbit**

**Orbiter Mass 700kg**

**Probe(s) maximum allowable mass**

**DM pressure vessel**

**Jupiter atm. descent**

**Jupiter atm. model**

**Comms architecture**

***≤****** ******?***

## Baseline Summary

| Launch vehicle |
| --- |
| Launch date |
| Number of orbiters |
| Orbiter JOI perijove  (Rj) |
| Orbiter JOI apojove (Rj) |
| Orbiter final perijove (Rj) |
| Orbiter final apojove (Rj) |
| Orbiter final inclination |
| Number of probes |
| Probe mass w margin (kg) |
| Probe release |
| Probe entry latitude (deg) |
| Max P for descent (bar) |

- 1

- 2

- Soyuz-Fregat 2-1b

- 4

- 5

- 200

- 100

- 70

- 15

- 200

- 100

- 70

- Equatorial

- Polar

- 1

- 2

- 150

- 250

- 300

- Hyperbolic, -90d

- Capture

- +3/-7

- ±15

- ±30

- 2016

- 2023

- 40

- 100

## Transfer to Jupiter: VEEGA

***VEEGA Transfer Summary***
- Earth escape 2016/12/18
- Hyperbolic escape velocity: 3.45 km/s
- Declination -30.5 deg
- No DSMs
- Venus swingby: 2017/5/12, 10137 km
- Earth swingby 1: 2018/3/6, 2489 km
- Earth swingby 2: 2020/3/19, 1423 km
- Arrival: 2022/11/17
- Hyperbolic arrival velocity: 5.2 km/s
- Arrival decl. wrt Jup. Equator: 1.3 deg
- Transfer duration: 2161 days (5.9 y)

## Mission Options ­ Probe Release 

- Assume: 4x200 Rj insertion orbit, 15x200 Rj final orbit, probe release *30d* (worst case for ODM) before entry for hyperbolic case, as best  V case JOI for capture (Io swingby)
- Calculate max possible probe mass given 10% launch margin
- Comms between probe and carrier very problematic
- Conclusion: Capture not favourable in terms of mass, entry speed not significantly reduced, direct comms problem => discard this option

|  | Hyperbolic | Capture |
| --- | --- | --- |
| Max Probe mass | 335 kg | 230 kg |
| Radiation | One pass at 4Rj | At least one pass at 4Rj |
| Comms range | ~3 Rj | ~3 Rj |
| Entry velocity | 47-48 km/s equatorial | 46-47 km/s equatorial |
| Operational | Short time between JOI, relay | More manoeuvres (PLM, PRM) |

## Swingby-Augmented JOI

- Relay phase

- 2nd JOI opportunity “after relay”

- Ganymede swingby
- (perijove – 16 h)

- Inbound arc
- (hyperbolic)

- Outbound arc
- (elliptic)

- perijove

- Here: Example with Ganymede swingby
- Alternative: Io swingby

- 1st JOI opportunity “before relay”

## Aerothermodynamics (1/2):Total absorbed heat fluxes assumed over JEP surface (100 bar probe)

**Convective blockage: 90% (conservative assumption)**
**Base cover: 2.5% of convective and 1.25% of radiative stagnation point flux (literature data)**

## Aerothermodynamics (2/2):Stagnation point Radiative heat fluxes at entry latitudes (Ve > 48-49 km/s)

**Heat fluxes are in excess of 800 MW/m****2**** in all cases.**
**This is beyond the capability of present TPS technologies.**

**Note: For Jupiter entry probes to higher latitudes larger TPS mass fractions are expected, e.g. for a 30deg latitude the TPS mass fraction probably exceeds 70%**

## TPS: SEPCORE Concept selected

-  The ablator is mounted on a hot structure, which is    insulated against the inner compartment using      lightweight insulation, possibly fibres
-  Considerable mass savings due to reduced ablator thickness and   use of more efficient insulation.

**Assumptions:**
** ****SEPCORE concept**
** ****Separation after 300s**
** ****Carbon phenolic on C/SiC structure**
** ****Assumed limit: ****1100degC**
** ****C/SiC backside adiabatic   ****  (conservative)**
**Required min. ablator thickness:**
** ****About 42mm ********  ****≈ 60 kg/m****2**** ablator**

**Assumptions:**
** “****Classical” heat shield**
** ****Separation after 300s**
** ****Carbon phenolic on CRFP structure**
** ****Assumed limit: ****150degC**
** ****CFRP backside adiabatic**
**Required min. ablator thickness:**
** ****About 60mm  ********  ****≈ 85 kg/m****2**** ablator**

**Notes:**
Sepcore: split the ablation and insulation function in two separate materials: each function is performed by optimal material
Still need more mass for insulation and structure and insulation, but overall mass is lower for SEPCORE

## Overview of probe designs

- All dimensions in mm

**100 bar probe:**
-  Mass ~ 300 kg
-  P/L resource ~ 12 kg, ~30 W (peak), ~350 bps
-  Entry latitude between -7 and +3 deg
-  One probe + one orbiter
-  Descent time = 1 hour
-  Variable power comms system to cope with very    strong atmospheric attenuation (~24 dB)
-  Comparable in mass to the Galileo probe   (but 100 bar vs. Galileo’s 20 bar)

**40 bar probe:**
-  Mass ~ 270 kg
-  P/L resource ~ 12 kg, ~30 W (peak), ~350 bps
-  Entry latitude between -7 and +3 deg
-  Two probes + one orbiter
-  Descent time = 1 hour
-  Comms scenario complicated but should be feasible

## Accommodation

- Antenna’s

- Pilot chute

- Main chute
- available volume

- Upper Shell

- Lower Shell

- Back cover
- 3 layers: ablator, structure, IFI

- Front shield
- 3 layers: ablator, structure, IFI

**Platform**

## Extended payload analysis

**Hypothesis**: increase of
- 100% of the P/L mass	 	17.4 	kg
- 100% of the P/L volume	10.8 	litres
- 50% of the P/L power	 	34 	W
- 50% of the P/L data rate	530 	bps
**Impact on volume**: DM diameter sphere increases from 600 mm to 650 mm
**Impact on probe mass**:

## JEP Study Conclusions

**Jupiter entry probes face extreme aerothermodynamic challenges: Identification of adequate TPS is very challenging **
**Testing in a representative environment is unfeasible, leading to large uncertainties in theoretical models. Generous TPS margins are therefore required**** **
**Suitability of available TPS materials is to be verified**
**High latitude entry conditions are beyond current material capabilities**
**The minimum probe configuration for the set requirements resulted in a Galileo-class probe, although the maximum depth is increased from 20 to 100 bar**** **
**Due to the very low resources and the atmospheric attenuation, comms are a major problem, driving the probe & S/C design: for this class of probe 100 bars is the limit. Furthermore the distance between probe and relay S/C is limited to the ~4 Rj range**** **
**High complexity and extreme test conditions are major cost drivers**** **
**Single vs. multiple probes = deep vs. “shallow => science trade-off**

## Any Questions ?

## Back-up slides

## Scientific rationale for JEP

- Initial assumptions on atmospheric composition: identical abundances as in Sun
- The present knowledge of the Jovian atmosphere is limited. These data have shown that Jupiter’s composition is not identical to the Sun’s: deviations up to 3 times the solar abundance have been measured
- Composition of atmosphere up to 20 bar is believed to be reasonably well understood, however in situ data very limited
- No quantitative results on O/H in the deep well-mixed atmosphere are available. Water was presumably the original carrier of heavy elements to Jupiter, hence the determination of its abundance in the deep atmosphere is of fundamental importance to the models of formation of Jupiter and the origin of its atmosphere as well as the origin of the solar system
- Furthermore, since meteorological and dynamical effects could cause the mixing ratios of water and possibly other volatiles to vary over the planet, it is essential to measure the full atmospheric composition, simultaneously with the related phenomena, such as winds and cloud properties
- The best way to accomplish this is by deploying deep multiprobes (50-100 bar) into different regions of Jupiter. However this study shows that with the given limitations this is hard to achieve. A new entry probe, even if with the same capabilities of the Galileo probe would contribute significantly, provided a entry zone is targeted consisting of a cloudy zone to contrast the Galileo Probe’s hot spot entry

## Strawman payload

| Instrument | Function |
| --- | --- |
| Atmospheric Structure Package (ASP) [accelerometers, gyros, p&T sensors, TPS recession sensors] |   Provide information about temperature, density, pressure, and    molecular weight of atmospheric gases. These quantities are to    be determined from the measured deceleration of the Probe    during the atmospheric entry phase as well as p&T sensors 
  Gyros: 3 degrees of freedom (for descent reconstruction) |
| Gas Chromatograph/Mass Spectrometer (GCMS) | Atmospheric composition: constituent mixing ratios; isotopic ratios |
| USO (Atomic Clocks) | Determine the vertical wind profile with an accuracy of a few m/s
Provide data for Probe localisation and trajectory reconstruction |
| Polarisation nephelometer | Main: Determine cloud location as a function of the pressure level
Secondary: Characterise composition, size and shape of cloud particles & aerosols |
| Radiometer | Vertical temperature profile/radiant energy flux |
| Speed of Sound Instrument | Speed of sound, through which ortho- to para-H2 ratio is determined (lowest priority as relevance for Jupiter unclear) |

## P/L Budgets

| Instrument | Mass
(kg) | Power
(W) | Volume
(l) | Data rate
Bps (compressed) | Duty cycle
(%) |
| --- | --- | --- | --- | --- | --- |
| GCMS | 5 | 13.5 | 2.6 | 250 | 100 |
| ASP | 0.7 | 0.5 | 0.5 | 30 | 10 |
| USO (Atomic clocks) | 0.5 | 7 | 1.2 | 0 | 100 |
| Nephelometer | 0.2 | 1.7 | 0.4 | 40 | 20 |
| Radiometer | 0.3 | 2.3 | 0.25 | 30 | 10 |
| Speed of Sound Instrument | 1 | 3 | 0.3 | 3 | 100 |
| DPU & co | 0.5 | 0.7 | 0.15 | 0 | 100 |
| Total (inc 20% margin) | 10 | 28.7 | 5 | 353 
(no margin) | -
 |

## From mission obj. to probe req.

***100 bar***

**Meas. altitude range**

***Heat shield dropped off***

**Entry and descent into the Jovian atmosphere**

**In-situ atmospheric measurements**

**Transmission of  the data in real time back to Earth**

**Entry probe + carrier Orbiter**

**Probe payload**

**Probe localisation knowledge**

**Comms link**
**via relay Orbiter**
**or direct to Earth**

**Aerodyn loads**

**Pressure vessel**

**Jupiter atmosphere attenuation model**

**Synchronisation probe/Orbiter (real time)**

**Localisation instrumentation**

**Payload mass,**
**volume, data rate, power constraints**

**Test facilities**

**Mission analysis**

**Battery sizing**

**TPS**

**Mission analysis**

**Jupiter radiation model**

**Jupiter atm. model**

**VLBI during the first part of the entry**

**Latitude entry**

## Measurement requirements

- Probe design shall maximise data return by allowing measurements as early as possible after entry (starting at release of front shield)
- Probe design shall allow atmospheric profile reconstruction with an accuracy of 10 km
- Mission design shall allow relay window to the Orbiter as large as possible during descent
- Probe design shall allow transmission at 353 bps during descent
- Probe shall carry, in addition to science payload, flight instrumentation to validate aerothermodynamic and ablation models
- Probe shall be aerodynamically stable during entry and the DM shall be aerodynamically stable during descent (within 5deg)

## Two-Probe Entry Geometry

- Probe 1: descent at +3.6 deg

- 1st relay phase

- 2nd relay phase

- Trajectories shown in non-rotating frame

- perijove

- Probe 2: descent at -6.8 deg

- 2nd JOI opportunity “after relay”

- Orbiter  trajectory

- 1st JOI opportunity “before relay”

## Probe design options

|  | Number Probes | Mass (kg) | Max Pressure (bar) | Entry Latitude (deg) | Required probe design modifications |
| --- | --- | --- | --- | --- | --- |
| Baseline 
“Minimum probe” | 1 | 301 | 100 | +3 or -7 |  |
| Option 1:
“40-bar probes” | 2 | 270 | 40 | +3 and -7 | Structure, TPS, Comms, Power |
| Option 2: 
“Non-equatorial probe” | 1 |   | 100 | -15.6 | Extremely high heat fluxes (heat flux above 900MW/m2)
Not affordable with present TPS technology
This option has not been further analysed |
| Option 3:
“Extended payload” (mass&volume ×2, data rate & power ×1.5) | 1 | 351 | 100 | +3 or -7
 | Instruments, Config (Increase of the DM diameter sphere to 650mm), TPS, structures |

- Three probe design options:

## Worst case atmosphere

Entry mass corresponding to the total mass
including system margins + 10kg

Worst case entry angle

Aerothermodynamic fluxes computed with fix blockage factor

TPS thickness increased by 20%

TPS mass computed with 20% margin

- Entry mass corresponding to the total mass
- including system margins + 10kg
- Worst case entry angle
- Aerothermodynamic fluxes computed with fix blockage factor
- TPS thickness increased by 20%
- TPS mass computed with 20% margin

- Specific design margin on TPS

## Galileo Recession Data

- Ablation data measured during the entry indicated that
- Stagnation point recession was less than predicted
- Ablation at frustum and shoulder was much higher than predicted
- More dissociation of ablation products
- Also recent mathematical models are not fully capable to explain the observed behavior.
- Possible reasons
- Enhanced turbulence from mass injection
- Particle spallation
- Complex interaction between ablation and shock layer radiation

## Preliminary TPS Sizing

**TPS based on SEPCORE concept**
**Heatshield separation after 170 sec.**
**Assumed limit at hot structure: 1100degC**
**Basecover**
**17mm Nylon-phenolic**
**Mounted on CMC hot structure (3.2mm)**
**Rearside insulated by 20mm fibrous insulation (IFI-type)**
**Frontshield**
**47mm Carbon phenolic**
**Constant thickness assumed**
**Mounted on CMC hot structure (3.2mm)**
**Rearside insulated by 20mm fibrous insulation (IFI-type)**
**Thicknesses include 20% uncertainty. Further 20% system margin are applied on the mass.**

## Mass budget – Baseline MINIMUM PROBE

- This is the minimum configuration probe
- that fulfils all the initial requirements

- It is comparable in mass to the Galileo probe (but 100 bar vs. Galileo’s 20 bar)
- Reaching 100 bar is possible but at the expense of a higher DM structural mass and therefore higher probe mass. Only single probe fits into the mission

## Mass budget for 2 probes at 40 bars

- Limiting the descent to between 20 to 40 bar gives a probe mass reduction which _***just ***_allows a two-probe mission with separate latitudes and longitudes at descent and relays in sequence

## Mission Environments and TPS Constraints

**Mission environments for ablative TPS applications and spallation limitations (from B. Laub and E. Venkatapathy, 2003)**

**Note: For Jupiter entry probes to higher latitudes larger TPS mass fractions are expected, e.g. for a 30deg latitude the TPS mass fraction probably exceeds 70%.**

## Galileo Recession Data

**Ablation data measured during the entry indicated that**
**Stagnation point recession was less than predicted**
**Ablation at frustum and shoulder was much higher than predicted**
**More chemical phenomena (dissociation)**
**Also recent mathematical models are not capable to explain the observed behaviour.**
**Possible reasons:**
**Enhanced turbulence from mass injection**
**Particle spallation**
**Complex interaction between ablation and shock layer radiation**

## Entry point for Different Inclinations

| Inclination [deg] | 25 | 30 | 35 |
| --- | --- | --- | --- |
| Entry angle [deg] | -16.1 | -12.5 | -10.2 |
| Entry velocity [m/s] | 48716 | 49177 | 49813 |
| Entry altitude [km] | 450 |  |  |
| Entry lat. [deg N] | -15.4 | -15.7 | -16.1 |
| Entry azimuth [deg] | 65.0 | 57.8 | 51.1 |
| Peak decel. [g] | 480 | 370 | 300 |

- Assumptions for the option 2 non equatorial analysis*:
- Mass = 500 kg
- Nose Radius = 0.65 m
- Base Diameter = 1.30 m
- *Option 2 discarded during
- Study phase

## Option 3 extended payloadAltitude versus Time

- Option 3 extended payloadAltitude versus Time

**Main chute release point**

**Without Main release**

**Heat Shield trajectory after release**

## Probe baseline design

- “Conventional” Carbon phenolic based Sepcore Front shield

- Back cover sized by the dimensions of the DM and the mortar

- Pilot chute for extraction of main chute

- Sealed Descent Module sized for 100 bar external pressure

- The overall dimensions and mass of the probe result from the dimensions of the DM which is sized by the equipment volume: Probe Mass cannot be reduced below a certain threshold

- Main chute for separation of front shield from DM

## Test facilities (1)

- NASA presentation on re-organising arc jet facilities:

** ****New JEP entry test facility in Europe can be expected to cost around 500 M€ **
- cost more than actual probe, and a very high investment for a facility that may only     be used for one project.

## Test facilities (2)

**Use high-velocity Earth atmosphere entry**
- Use multiple Earth and lunar flyby’s to accumulate velocity, maybe also use upper stage or dedicated propulsion module to further accelerate test probe.
- At high altitude, Earth’s atmosphere is dominated successively by helium, molecular hydrogen, and atomic hydrogen => simulate Jupiter atmosphere entry?
- 34 MW/m2 expected for EVD (16 from convection, 18 from radiation)
- Share data (and costs?) with other projects requiring high-velocity entry data: Aurora, FLPP reusable launcher studies...
**Pre-phase A studies on lower-velocity EVD show max velocity about 12.5 km/s.**

## Test facilities (3)

- Full simulation of JEP entry conditions very probably not possible with any of the existing facilities:
- Not same gas velocity
- Not same atmosphere (can any of the test results for nitrogen atmosphere be extrapolated to He/H atmosphere of Jupiter?)
- Not same radiative / convective energy ratio

## Test facilities (4)

**No facility existing achieving the level of heat-flux reaches during a Jupiter entry.**
**Plasma facilities generally have difficulties in providing radiative fluxes**
**Most powerful (in terms of heat-flux) European  TPS facilities:**
**SCIROCCO (CIRA) - segmented arc heater - 3.8 MW/m****2**
**L3K (DLR) - segmented arc heater -  4 MW/m****2**
**Plasmatron (VKI) -  3 MW/m****2**
**SIMOUN (EADS) – Huels arc heater -  4 MW/m****2**
**COMETE (EADS) – Plasmatron – 7 MW/m****2**
**JP 200 (EADS) - Huels arc heater – 25 MW/m****2**
**High Pressure (EADS) – Huels arc heater – ? MW/m****2**
**PWK4 (IRS) – Magnetoplasmadynamic generator – 100 MW/m****2**

## Test facilities (5)

**EADS-ST plasma test facilities**** **
**(Source: *****Testing of Ablative Material at EADS Space Transportation, J.-M. Bouilly, D. Conte, F. Leleu, P. Plotard, Ablation Working Group, ESTEC, Oct. 13, 2005)*****.**

## Test facilities (6)

**Three facilities may have an interest for a Jupiter entry in term of heat-flux, however: **
**PWK4 works with subsonic flows;**
**The two others are dedicated to military activities (French deterrent force) so little information is available. They are both supersonic and only small models (0.05 m) can be tested.**
**Possible improvements of SCIROCCO:**
**	****	****- With some modifications of the set-up, heat fluxes in the range of ****	****20 MW/m****2**** might be reached for samples of 25 mm.**
**	****	****- Some efforts in this direction are actually investigated in the frame ****	****of a technology study related to EVD (AURORA Programme).**
***Source****: Review of European Facilities for Space Aerothermodynamics, RT 1/06302 DMAE, ONERA, May 2003.*